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Model 1900/1900C Airliner (UA-1 and After) (UB-1 and After) (UC-1 and After)

Maintenance Manual

Volume 1 Introduction thru Chapter 28

Copyright © 2017 Beechcraft Corporation. All rights reserved. Hawker and Beechcraft are trademarks of Beechcraft Corporation. P/N 114-590021-7 Issued: November 12, 1982

P/N 114-590021-7C12 Revised: January 1, 2017

Published by Beechcraft Corporation P.O. Box 85 Wichita, Kansas 67201-0085 USA

The export of these commodities, technology or software are subject to the US Export Administration Regulations. Diversion contrary to US law is prohibited. For guidance on export control requirements, contact the Commerce Department’s Bureau of Export Administration at 202-482-4811 or visit the US Department of Commerce website.

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

LIST OF EFFECTIVE REVISIONS Part Number

Date

Chapters Affected

114-590021-7C

November 1, 2009

All (Reissue)

114-590021-7C1

February 1, 2010

Introduction, 04, 05, 12, 27, 33 and 56

114-590021-7C2

May 1, 2010

Introduction, 04, 05, 06, 11, 12, 20, 32, 33, 39, 52, 56, 57, 71, 73, 76, 77 and 91

114-590021-7C3

August 1, 2010

Introduction, 05, 12, 32, 54, 61, 73 and 79

114-590021-7C4

November 1, 2010

Introduction, 20, 21, 28, 61, 76, 79 and 91

114-590021-7C5

May 1, 2011

Introduction, 05, 12, 21, 24, 25, 27, 28, 32, 52, 57 and 91

114-590021-7C6

November 1, 2011

05, 27, and 91

114-590021-7C7

May 1, 2012

05, 20, 26, 27, 28, 32, 33 and 34

114-590021-7C8

August 1, 2012

05, 27 and 32

114-590021-7C9

November 1, 2012

05, 20, 28, 30 and 55

114-590021-7C10

November 1, 2013

Introduction, 05, 06, 12, 20, 21, 24, 27, 28, 32, 52, and 56

114-590021-7C11

November 1, 2015

27

114-590021-7C12

January 1, 2017

39

C12 A

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

List of Effective Pages Section

PAGE

DATE

Title Page (Vol 1) Jan 1/17 Logo Page Title Page (Vol 2) Jan 1/17 Logo Page

Introduction

“A” Page “B” Page

C12 C12

1 thru 18

Nov 1/13

NOTE - The chapter List of Effective Pages is located in the front of each chapter.

C12 B

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Record of Revisions Note: When a revision is inserted, the revision number, the date the revision is inserted into the manual, and the initials of the person(s) inserting the revision should be recorded on this page. Rev No.

Date Inserted

Init

C

12/16/2009

ATP/PC

C1

2/10/2010

ATP/VP

C2

5/28/10

ATP/PC

C3

8/20/2010

ATP/MB

C4

12/22/2010

ATP/GM

C5

6/15/2011

ATP/VDR

C6

11/16/2011

ATP/RLL

C7

7/20/2012

ATP/VDR

C8

10/5/2012

ATP/PC

C9

1/16/2013

ATP/RLL

C10

12/24/2013

ATP/LC

C11

NOV 9, 2015

ATP/RLL

C12

JAN 27, 2017

ATP/LC

Rev No.

Date Inserted

Init

Rev No.

Date Inserted

Init

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Log of Temporary Revisions Note: Insert this Log of Temporary Revisions after the Record of Revisions page. Previous Log of Temporary Revisions may be discarded. Update the Record of Temporary Revisions page(s) as required. Revision No.

Revision Date

12-1

Nov 1/09

Provides additional instructions on the application of de-ice/anti-ice fluids and the removal of thickened residue

C1

05-1

Nov 1/11

Nose Landing Gear Actuator Ultrasonic Inspection

C7

Subject

Revision Incorporated

Temporary Revision 05-1 Nov 1/11

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Record of Temporary Revisions Note: Insert this Record of Temporary Revisions after the Log of Temporary Revisions page(s). When a Temporary Revision is inserted, the temporary revision number, the affected chapter, the date the revision is inserted into the manual, and the initials of the person(s) inserting the revision should be recorded on this page. When a Temporary Revision is removed, enter the manual revision number that incorporated the Temporary Revision and the date the Temporary Revision was removed from the manual. Temporary Revision No.

Affected Chapter

Inserted

Removed

Init

05-1

05-10-00

ATP/RLL

11/16/2011

C7

5/1/2012

12-1

12-30-00

ATP/PC

12/16/2009

C1

2/10/2010

Date

By Revision No.

Date

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

P/N 114-590021-7, Revision C12, Jan 1/17 The chapters which have been revised or added are listed below with the Highlights of each change. Remove the affected pages and insert this Revision in accordance with the attached Instruction Page. Enter the revision number and the date inserted on the Record of Revisions page of this manual. The Highlights Page may be retained with the manual for future reference.

Highlights Chapter/Section

Description of Change

Title Page Volume 1

Revision date and number revised.

Title Page Volume 2

Revision date and number revised.

Chapter 39 39-LOEP

Page dates and page numbers revised.

39-10-00, 201

Added new Caution to top of page block.

C12 Jan 1/17

HIGHLIGHTS

Page 1

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Instruction Sheet for P/N 114-590021-7, Revision C12, Jan 1/17 Remove Page

Ch-Se-Su

Insert Page

Ch-Se-Su

Dated

Title Page (Vol 1)

Title Page (Vol 1)

Jan 1/17

Logo Page

Logo Page

---

“A” Page

“A” Page

C12

“B” Page

“B” Page

C12

Title Page (Vol 2)

Title Page (Vol 2)

Jan 1/17

Logo Page

Logo Page

---

1

39-LOEP

1

39-LOEP

Jan 1/17

201 thru 218

39-10-00

201 thru 218

39-10-00

Jan 1/17

After compliance, this Instruction Sheet may be discarded.

C12 Jan 1/17

PAGE 1 OF 1

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL --Introduction

1. INFORMATION Since a wide variety of avionic components and equipment is available, avionic manufacturers normally supply parts and servicing manuals with each set/component. The manufacturer of the equipment should be contacted when additional parts or servicing information is required. Reissues and revisions are automatically provided to the subscription holders of the manual. Additional publications are listed on the web at http://pubs.beechcraft.com. For more information on these publications, or to check subscription status, contact the Technical Manual Distribution Center (TMDC) at 1.800.796.2665 or 316.676.8238, fax 316.671.2540, E-mail [email protected]. The Interactive Maintenance Library (IML) contains selected Manuals in a digital format. This manual, along with others, is available on CD-ROM and Online. Optional paper copies of the manuals on the CD-ROM are available for purchase. Help Line phone support for IML CD and Online Users is available 8:00 AM to 4:30 PM Central Time (US and Canada). During off-hours, leave a detailed voicemail message. Calls will be returned within one business day. Contact the Help Line at 1.800.240.2959 or 1.316.676.3053, E-mail [email protected]. Comprehensive user guides for the IML CD and Online manuals are available on the Beechcraft Corporation (BC) Technical Publications website http://pubs.beechcraft.com. Illustrated and detailed procedures for using IML Features are included in the downloadable user guides.

A. General This Maintenance Manual applies to the Beechcraft Corporation 1900 Airliner (UA Serials), 1900C Airliner (UB and UC Serials). The wording MODEL 1900/1900C AIRLINER appearing in the masthead at the top of each page is not intended to indicate applicability beyond that noted above. Separate Maintenance Manual coverage is provided for subsequent models in the 1900 Airliner Series. The Model 1900/1900C Airliner Maintenance Manual is prepared in accordance with the ATA (Air Transport Association of America) Specification 2200 format. It meets the requirements with respect to the arrangement and content of the System/Chapters within the designated chapter-numbering system. This manual includes the maintenance information required to be available by 14 CFR Part 23. In addition to this manual and its subsequent revisions, additional maintenance information is published in the form of Beechcraft Corporation service bulletins. The information contained in these service bulletins is an integral part of, and is to be used in conjunction with, the information contained in this manual. This Maintenance Manual is supplemented by the following publications: •

Model 1900 and 1900C Airliner Parts Catalog, P/N 114-590021-5 (UA-1 and After; UB-1 and After)

•

Model 1900C Airliner Parts Catalog, P/N 114-590021-59 (UC-1 and After)

•

Model 1900 Airliner Series Wiring Diagram Manual, P/N 114-590032-3 (UA-1 and After)

•

Model 1900C Airliner Wiring Diagram Manual, P/N 114-590021-13 (UB-1 and After)

•

Model 1900C Airliner Wiring Diagram Manual, P/N 114-590021-61 (UC-1 and After)

•

Model 1900 Airliner Series Component Maintenance Manual, P/N 114-590021-11

•

Model 1900 Airliner Series Structural Repair Manual, P/N 114-590021-9

•

Model 1900/1900C Airliner Structural Inspection Manual, P/N 98-30937

INTRODUCTION

Page 1 Nov 1/13

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL •

Model 1900 Airliner Series Corrosion Control Manual, P/N 114-590021-197

•

Model 1900 Series Airliners Kit Supplement Illustrated Parts Catalog, P/N 129-590000-131

•

Model 1900 Airliner Series Airworthiness Limitations Manual, P/N 129-590000-133

It shall be the responsibility of the owner/operator to ensure that the latest revision of the publications referenced in this manual are utilized during operation, servicing and maintenance of the airplane. Beechcraft Corporation expressly reserves the right to supersede, cancel and/or declare obsolete any parts, part numbers, kits or publications that may be referenced in this manual without prior notice. WARNING: Use only parts obtained from sources approved by Beechcraft Corporation, in connection with the maintenance and repair of Beechcraft Corporation airplanes. Genuine Beechcraft Corporation parts are produced and inspected under rigorous procedures to insure airworthiness and suitability for use in Beechcraft Corporation airplane applications. Parts purchased from sources other than those approved by Beechcraft Corporation, even though outwardly identical in appearance, may not have the required tests and inspections performed, may be different in fabrication techniques and materials, and may be dangerous when installed in an airplane. Salvaged airplane parts, reworked parts obtained from sources not approved by the Beechcraft Corporation or parts, components or structural assemblies, the service history of which is unknown or cannot be authenticated, may have been subjected to unacceptable stresses or temperatures or have other hidden damage, not discernible through routine visual or usual nondestructive testing techniques. This may render the part, component or structural assembly, even though originally manufactured by the Beechcraft Corporation, unsuitable and unsafe for airplane use. Beechcraft Corporation expressly disclaims any responsibility for malfunctions, failures, damage or injury caused by use of parts not approved by the Beechcraft Corporation. Any maintenance requiring the disconnection and connection of flight control cables, plumbing, electrical connectors or wiring requires identification of each side of the component being disconnected to facilitate correct reassembly. At or prior to disassembly, components should be color coded, tagged or properly identified in a way that it will be obvious how to correctly reconnect the components. After connection of any component, remove all identification tags. Check all associated systems for correct function prior to returning the airplane to service.

B. Correspondence If a question should arise concerning the care of your airplane, it is important to include the airplane serial number in any correspondence. The serial number appears on the model designation placard. Refer to Chapter 11 for placard location.

C. Publications Change Request (PCR) If an irregularity or missing information is noted, the user of this manual may access a PCR form at http://pubs.beechcraft.com. Instructions on how to submit a PCR are available on the web page.

Page 2 Nov 1/13

INTRODUCTION

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

D. Normal Revisions Normal Revisions to this manual are issued to provide changes to maintenance information. 1. Paper Revision That portion of text which has been revised by the addition of, or a change in, information is denoted by a solid revision bar adjacent to the text. The date printed on the bottom of each page can be compared to the “A” page to determine the revision number. Each revised page will ONLY show revision bars for text changed by the revision. There will not be a revision bar if text was deleted from the page. Revised illustrations will be identified by a revision bar printed on the side of the page. 2. CD-ROM Revision Normal revised text on the CD-ROM will be highlighted yellow across the revised passage of text. For each revision of this manual, a new CD-ROM will be issued. The CD-ROM may contain revised illustrations. Revisions to the illustrations are not identified.

E. Temporary Revisions Temporary Revisions to this manual are issued to provide maintenance information in the interim between normal revisions. Each temporary revision is issued by the chapter number to which it applies, followed by a sequential number in the order of publication (Temporary Revisions 12-1, 12-2, etc.). If relevant, the information in the temporary revision should be included in the next normal revision of the manual. 1. Paper Temporary Revisions Temporary Revisions are printed on yellow paper and are to be inserted in the maintenance manual in accordance with the instructions provided and adjacent to applicable chapter, section, and subject matter in the manual. 2. CD-ROM Temporary Revisions A new CD-ROM will be issued for each Temporary Revision to this manual. This information is listed in conjunction with the applicable chapter, section, subject on the CD-ROM.

F. Revised Text That portion of text which has been revised by the addition of, or a change in, punctuation and/or information is denoted by a solid revision bar adjacent to the textual column in the margin of this paragraph. Each page may or may not have revision bars. That date printed on the bottom of each page indicates when the information on that page was changed. Each page will ONLY show revision bars for punctuation and/or text changed by the current revision. Revised text in the IML will be denoted by yellow highlighting.

G. Revised Illustrations When an illustration is modified or a new illustration is added, it will be noted by a solid line (revision bar) along the outside margin of the illustration.

INTRODUCTION

Page 3 Nov 1/13

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

H. Warnings, Cautions, and Notes •

WARNING - Brings attention to an operating procedure, inspection or maintenance practice, which if not correctly followed, could result in personal injury or loss of life.

•

CAUTION - Brings attention to an operating procedure, inspection, repair or maintenance condition, which if not strictly observed, could result in damage or destruction of equipment.

•

NOTE - Brings attention to an operating procedure, inspection, repair or maintenance condition, which is essential to highlight.

I. Special Conditions Cautionary Notice Airplanes operated for Air Taxi, or other than normal operation, and airplanes operated in humid tropics, cold and damp climates, etc., may need more frequent inspections for wear, corrosion and/or lack of lubrication. Under these adverse conditions, perform periodic inspections in compliance with this guide at more frequent intervals until the owner or operator can set his own inspection periods based on the contingencies of field experience. CAUTION: The recommended periods do not constitute a guarantee the item will reach the period without malfunction as the aforementioned factors cannot be controlled by the manufacturer.

2. MANUAL LAYOUT A. Title Page A Title page is located at the beginning of the manual and provides the part number of the manual, and lists all aircraft models pertaining to this manual and their respective serial numbers. Information throughout this manual is applicable to all serial numbers listed on the title page except where specifically stated.

B. List of Effective Revisions/List of Effective Pages The printed manual will have a List of Effective Revisions/List of Effective Pages, (“A” page) following the title page of the manual. The List of Effective Revisions page lists the revisions currently effective for the manual. The List of Effective Pages section lists the page effectivity for the Title Page(s), “A” page(s) and Introduction chapter. It will also show the effective pages for an entire manual if the manual does not have individual Chapter List of Effective Pages.

C. Record of Revisions Page The printed manual will have a Record of Revisions page. The Record of Revisions is provided following the List of Effective Revisions/List of Effective Pages (“A” page). When a revision is inserted, the revision number, the date the revision is inserted into the manual, and the initials of the person(s) inserting the revision should be recorded on this page. IML CD and Online Manuals do not include a Record of Revisions page. Revisions standard is available by a link to the catalog Title Page or accessible by clicking Help on the menu bar and selecting ”About This IML Book...”.

Page 4 Nov 1/13

INTRODUCTION

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

D. Log of Temporary Revisions Page The printed manual will have a Log of Temporary Revisions page. The Log of Temporary Revisions Page is located following the Record of Revisions page. The Log of Temporary Revisions page provides a history of each temporary revision, including the revision number which incorporated the temporary revision into the manual.

E. Record of Temporary Revisions Page The printed manual will have a Record of Temporary Revisions page. The Record of Temporary Revisions Page is located following the Log of Temporary Revisions page. When a temporary revision is inserted or removed from this manual, the appropriate information should be recorded on this page.

F. Introduction This section contains general and specific information on how to use this manual.

G. Chapter List of Effective Page The printed manual may have a Chapter List of Effective Pages. The List of Effective Pages follow the Chapter-Divider-Tab and lists the issue date of each page that is effective for that chapter.

H. Chapter Table of Contents Pages The printed manual may have a Chapter Table of Contents Pages. The Chapter Table of Contents Pages follow the Chapter List of Effective Pages and lists the contents of the data for that chapter.

3. HOW TO USE THE MANUAL A. ATA Subject Matter Assignment The contents of this manual are organized into four levels. The four levels are: (1) Group These are the primary divisions of the manual that enable broad separation of content. Typical of this division is the separation between Airframe Systems and the Power Plant. (2) System/Chapter The various groups are broken down into major systems such as Environmental Systems, Electrical Power, Landing Gear, etc. The systems are arranged more or less alphabetically rather than by precedence or importance. They are assigned a number, which becomes the first element of the standardized numbering system. Thus, the element 28 of the number 28-40-01 refers to the chapter FUEL. Everything concerning the fuel system will be covered in this chapter. (3) Subsystem/Section The major systems/chapters of an airplane are broken down into subsystems. These subsystems are identified by the second element of the standard numbering system. The element 40 of the number 28-40-01 concerns itself with the indicating section of the fuel system.

INTRODUCTION

Page 5 Nov 1/13

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

(4) Unit/Subject The individual units within a subsystem/section may be identified by the third element of the standard numbering system. The element 01 of the number 28-40-01 is a subject designator. This element is assigned at the option of the manufacturer and may or may not be used.

B. Application Any publication conforming to the GAMA or ATA format will use the same basic numbering system. Thus, whether the manual is a Model 1900/1900C Airliner Maintenance Manual, or a Model 1900C Airliner Wiring Diagram Manual, the person wishing information concerning the indication portion of the fuel system, would refer to the System/Chapter Tab 28-FUEL. The table of contents in the front of this chapter will provide a list of subsystems covered in this chapter. For example, the fuel system chapter with a full index would contain: 28-00 - General 28-10 - Storage (Tanks, cells, necks, caps, instruments, etc.). 28-20 - Distribution (Fuel lines, pumps, valves, controls, etc.). 28-30 - Dump (If in-flight dumping system is installed, it would appear here). 28-40 - Indicating (Quantity, temperature, pressure, etc., does not include engine fuel flow or pressure).

C. References to Procedures, Figures, Equipment and Materials A system has been developed to provide a method of allowing the user to quickly locate data referred to in the other manuals. This system provides information for both the printed manual as well as a hyperlink in electronic manuals. Here are a few examples: When the user is directed to a procedure in another manual, the text will be as follows: (Ref. Chapter 06-50-00, FUSELAGE ACCESS PANELS). The procedure is found in Chapter 06-50-00, of the Maintenance Manual. When the user is in a procedure and is directed to information about a tool, piece of equipment or material used, the text will be as follows: (4, Table 2, 27-00-00), which contains a listing of all the special tools and equipment used to maintain the 1900/1900C airplane. Item 4 in Table 2 is Corrosion Preventive Compound.

4. CHAPTER/SYSTEM INDEX GUIDE TABLE The following System/Chapter, Subsystem/Section Index Guide is prepared in accordance with ATA Specification No. 2200. This outline in general is used in Maintenance Manuals, Parts Catalogs and Wiring Diagram Manuals. The organization of information will follow this outline, however the Subsystem/Sections may vary slightly to accommodate accurate coverage of the specific aircraft. The following chapters are not applicable to this Maintenance Manual: 29, 31, 38, 49, 60, 65, 70, 75, 81, 82, 83 and 95.

Page 6 Nov 1/13

INTRODUCTION

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

System/Chapter

Subsystem/Section

Title

INTRODUCTION AIRCRAFT GENERAL 04

AIRWORTHINESS LIMITATIONS 00-00

05

06

07

Inspection/Check

TIME LIMITS/MAINTENANCE CHECKS 00-00

General Information

10-00

Time Limited Inspections

11-00

Major Maintenance Schedule

20-00

Scheduled Maintenance Checks

20-01

Routine Inspection

20-02

First 200-Hour Interval Detailed Inspection

20-03

Second 200-Hour Interval Detailed Inspection

20-04

Third 200-Hour Interval Detailed Inspection

20-05

Fourth 200-Hour Interval Detailed Inspection

20-06

Fifth 200-Hour Interval Detailed Inspection

20-07

Sixth 200-Hour Interval Detailed Inspection

50-00

Unscheduled Maintenance Checks

DIMENSIONS AND AREAS 00-00

General Information

10-00

Airplane Dimensions

30-00

Airplane Stations

40-00

Airplane Zones

50-00

Fuselage Access Panels

LIFTING AND SHORING 00-00

General Information

10-00

Jacking

20-00

Shoring

INTRODUCTION

Page 7 Nov 1/13

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL System/Chapter 08

09

Subsystem/Section LEVELING AND WEIGHING 00-00

General Information

10-00

Weighing and Balancing

20-00

Leveling

TOWING AND TAXIING 10-00

10

11

12

Title

Towing

PARKING, MOORING, STORAGE AND RETURN TO SERVICE 10-00

Parking

10-01

Storage

20-00

Mooring

30-00

Return to Service

PLACARDS AND MARKINGS 20-00

Exterior Placards and Markings (UA-1 and After; UB-1and After)

21-00

Exterior Placards and Markings (UC-1 and After)

30-00

Flight Control Rig Pin Fuselage and Empennage Placards

SERVICING 00-00

General Information

10-00

Replenishing

20-00

Scheduled Servicing

30-00

Unscheduled Servicing

AIRFRAME SYSTEMS

20

Page 8 Nov 1/13

STANDARD PRACTICES - AIRFRAME 00-00

General Information

00-01

Electrical Bonding Procedures

INTRODUCTION

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL System/Chapter 20 (Cont’d)

21

Subsystem/Section

Title

00-02

Control Cables and Pulleys

00-03

Wiring

00-04

Electrostatic Discharge Sensitivity

00-05

Tubing, Hose and Fittings

01-00

Torque Wrenches

04-00

Leading Edge Erosion Protection

07-00

Locking Devices

08-00

Airplane Finish Care

09-00

Corrosion

10-00

Airframe Penetration Inspection

15-00

Control of Life-Limited Parts

ENVIRONMENTAL SYSTEMS 00-00

General Information

10-00

(UA-1 and After, UB-1 and After and UC-1 thru UC-45, Except UC-39) Bleed Air Control

10-01

Precooler-Through and Bypass Valves

10-03

Pressure Regulator/Shutoff Valves

10-06

ACM Overpressure Switch

10-07

ACM Overtemperature Switch

11-00

(UC-39, UC-46 and After) Bleed Air Control

11-01

Precooler-Through and Bypass Valves

11-02

Temperature Controller Sense Line Filter

11-03

Pressure Regulator/Shutoff Valves

11-06

ACM Overpressure Switch

11-07

ACM Overtemperature Switch

20-00

Distribution

20-01

Vent Blower

20-02

Air Outlet

30-00

Pressurization Control

30-01

Outflow Valve

30-02

Cabin Pressure Controller

INTRODUCTION

Page 9 Nov 1/13

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL System/Chapter 21 (Cont’d)

22

Subsystem/Section 52-02

Compressor (UA-1 and After; UB-1 and After; UC-1 thru UC-100 Not Modified by Service Bulletin NO. 2345)

30-03

Pneumatic Relay

30-04

Volume Tank

30-07

Cabin Altitude Warning Pressure Switch

40-00

Heating

50-00

Cooling

51-00

Air Cycle System

51-01

Refrigeration Package

51-02

Fog Nozzle and In-line Filter

51-03

Recirculating Ejector

52-00

Vapor Cycle System

52-01

Evaporator

52-03

Condenser and Blower

52-04

Receiver and Dryer

52-05

Compressor (UA-1 and After; UB-1 and After; UC-1 and After Modified by Service Bulletin NO. 2345)

60-00

Temperature Control

60-01

Air Duct Temperature Sensor

60-02

Cabin Temperature Controller

AUTO FLIGHT 10-00

23

24

Page 10 Nov 1/13

Title

Autopilot

COMMUNICATIONS 10-00

Speech Communication

60-00

Static Discharging

ELECTRICAL POWER 00-00

Electrical System

20-00

AC Power and Control

30-00

DC Generation and Control

30-01

Starter-Generator

INTRODUCTION

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL System/Chapter 24 (Cont’d)

25

26

27

Subsystem/Section

Title

30-02

Generator Control Panel

31-00

Battery Power and Control

32-00

Battery Monitor

40-00

External Power and Control

50-00

Electrical Load Distribution

EQUIPMENT/FURNISHINGS 10-00

Flight Compartment - Seats

20-00

Passenger Seats - Seats

20-01

Passenger Compartment - Carpet

20-03

Passenger Compartment Sidewall Upholstery

60-00

Emergency Locator Transmitter (ELT)

FIRE PROTECTION 10-00

Fire Detection System

11-00

Engine Bleed Air Warning System

20-00

Fire Extinguishing System

FLIGHT CONTROLS 00-00

General Information

00-01

Control Column Bearing Support

00-02

Travel Board

10-00

Ailerons

10-01

Control Wheel

10-02

Aileron Cables

10-03

Aileron Control System

10-04

Aileron Trim Tab

10-05

Aileron Trim Tab Actuators and Cables

10-06

Aileron Trim Tab Indicator

10-07

Aileron Trim Tab Control System

10-08

Aileron Balance Weights

20-00

Rudder

20-01

Rudder Cables

20-02

Rudder Control System

INTRODUCTION

Page 11 Nov 1/13

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL System/Chapter 27 (Cont’d)

28

Page 12 Nov 1/13

Subsystem/Section

Title

20-03

Rudder Pedals

20-04

Rudder Trim Tab

20-05

Rudder Trim Tab Cables and Actuators

20-06

Rudder Trim Tab Indicator

20-07

Rudder Trim Tab Control System

21-00

Flight Control Assist Systems

30-00

Elevator

30-01

Elevator Cables

30-02

Elevator Control System

30-03

Elevator Trim Tabs

30-04

Elevator Trim Tab Cables

30-05

Elevator Trim Tab Control System

30-06

Elevator Trim Tab Actuators

30-07

Elevator Electric Trim Tab System

30-08

Elevator Trim Tab Indicator

31-00

Stall Warning System

50-00

Flaps

50-01

Flap Cables

50-02

Flap Tracks

50-03

Flap Motor and Gearbox

50-04

Flap Actuators

50-05

Flap Control System

50-06

Flap Safety System

50-07

Flap Position Switches

70-00

Gust Locks and Dampeners

FUEL 00-00

Fuel System (UA-1 and After; UB-1 and After)

01-00

Fuel System (UC-1 and After)

10-00

Fuel Storage (UA-1 and After; UB-1 and After)

10-01

Antisiphon Valve (UA-1 and After; UB-1 and After)

INTRODUCTION

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL System/Chapter 28 (Cont’d)

30

32

Subsystem/Section

Title

11-00

Fuel Storage (UC-1 and After)

11-01

Antisiphon Valve (UC-1 and After)

20-00

Fuel Distribution (UA-1 and After; UB-1 and After)

20-02

Fuel Filters and Screens (UA-1 and After; UB-1 and After)

20-03

Fuel Pumps (UA-1 and After; UB-1 and After)

20-05

Fuel Valves (UA-1 and After; UB-1 and After)

21-00

Fuel Distribution (UC-1 and After)

21-01

Fuel Fittings (UC-1 and After)

21-02

Fuel Filters and Screens (UC-1 and After)

21-03

Fuel Pumps (UC-1 and After)

21-04

Fuel Manifolds (UC-1 and After)

21-05

Fuel Valves (UC-1 and After)

21-06

LH and RH Fuel Lines (UC-1 and After)

40-00

Fuel Quantity Indicating (UA-1 and After; UB-1 and After)

40-01

Fuel Level Sensors (UA-1 and After; UB-1 and After)

41-00

Fuel Quantity Indicating (UC-1 and After)

41-01

Fuel Level and Low Fuel Quantity Sensors (UC-1 and After)

ICE AND RAIN PROTECTION 00-00

General Information

01-00

Brake Deice System

10-00

Airfoil

20-00

Air Intakes

40-00

Windows and Windshields

60-00

Propeller

LANDING GEAR 00-00

General Information

10-00

Main Landing Gear

20-00

Nose Landing Gear

INTRODUCTION

Page 13 Nov 1/13

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL System/Chapter 32 (Cont’d)

Page 14 Nov 1/13

Subsystem/Section

Title

30-00

Landing Gear Extension and Retraction

30-01

Landing Gear Hydraulic Accumulator

30-02

Landing Gear Power Pack

30-03

Landing Gear Power Pack - Sperry Vickers Valve Housing and Controls

30-04

Landing Gear Power Pack Motor

30-05

Power Pack Gear Up Pressure Switch

30-06

Landing Gear Hydraulic Power Pack Filters

30-07

Landing Gear Power Pack Gear-Up and Gear-Down Port Filters

30-08

Landing Gear Power Pack Fluid Level Sensor

30-09

Main Landing Gear

30-10

Main Landing Gear Actuator

30-11

Main Landing Gear Actuator Orifice

30-12

Main Landing Gear Doors

30-13

Nose Landing Gear

30-14

Nose Landing Gear Actuator

30-15

Nose Landing Gear Door

30-16

Hydraulic Landing Gear Service Valve Assembly

30-17

Emergency Extension Hand Pump Assembly

40-00

Wheels and Brakes

41-00

Antiskid Brakes

42-00

Brake Deice System

50-00

Mechanical Steering

51-00

Power Steering (UA-1 and After; UB-1 and After)

52-00

Power Steering (UC-1 and After)

60-00

Landing Gear Position and Warning

60-01

Nose Gear Down-Position Switch

60-02

Main Gear Down-Position Switch

60-03

Nose Gear Actuator Downlock Switch

60-04

Main Gear Actuator Downlock Switch

60-05

Main Gear Safety Switch

INTRODUCTION

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL System/Chapter

Subsystem/Section

32 (Cont’d)

60-06

33

LIGHTS

34

35

Landing Gear Warning Horn

00-00

General Information

10-00

Flight Compartment

20-00

Passenger Compartment

30-00

Baggage and Cargo Compartment

40-00

Exterior

50-00

Self-Illuminated Signs

50-01

Emergency Exit Lighting System (Optional Installation with External Floodlights)

50-02

Emergency Exit Lighting

NAVIGATION 10-00

Flight Environment Data

20-00

Magnetic Compass

50-00

Radio Magnetic Indicator System

OXYGEN 00-00

36

37

Oxygen System

PNEUMATIC 00-00

Pneumatic System

10-00

Pneumatic Distribution System

VACUUM 00-00

39

Title

Vacuum System

ELECTRIC PANELS, PARTS AND INSTRUMENTS 00-00

Electrical Panels and Components

10-00

Instrument and Control Panels

20-00

Electrical and Electronic Equipment Racks

INTRODUCTION

Page 15 Nov 1/13

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL System/Chapter

Subsystem/Section

Title

STRUCTURE

51

STRUCTURES 00-00

52

53

54

55

56

Page 16 Nov 1/13

General Information

DOORS 10-00

Airstair Door

20-00

Emergency Exit

20-01

Emergency Exit Latch Mechanism

30-00

Cargo/Nose Baggage Compartment Doors

70-00

Cargo and Airstair Door Warning

FUSELAGE 00-00

Fuselage and Floor Access Openings

10-00

Main Frame

40-00

Attach Fittings

NACELLES 00-00

General

10-01

Nacelle Inner Fender

30-00

Nacelle Plates/Skins

STABILIZERS 10-00

Horizontal Stabilizer

10-01

Stabilon

10-02

Tail-Let

20-00

Elevator

30-00

Vertical Stabilizer

40-00

Rudder

WINDOWS 00-00

Windows

10-00

Flight Compartment

INTRODUCTION

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL System/Chapter

Subsystem/Section

56 (Cont’d)

20-00

57

WINGS

Title Cabin

00-00

Wings

10-00

Main Frame

30-00

Plates/Skins

50-00

Aileron

PROPELLER

61

PROPELLER 10-00

General Information

20-00

Propeller Controlling

21-00

Propeller Autofeathering

22-00

Propeller Synchrophaser

40-00

Propeller Indicating

POWER PLANT

71

72

POWER PLANT 00-00

General Information

10-00

Cowling

20-00

Mounts

30-00

Fireseals

50-00

Electrical Harness

70-00

Engine Drains

ENGINE 00-00

73

General Information

ENGINE FUEL SYSTEMS 10-00

Distribution

INTRODUCTION

Page 17 Nov 1/13

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL System/Chapter 74

Subsystem/Section IGNITION 00-00

76

77

00-00

General Information

10-00

Power Control

ENGINE INDICATING

80

00-00

Oil System

10-01

Oil Tank

10-02

Oil Breather (UA-1 thru UA-3, UB-1 thru UB 52 without Kit 114-9006-1)

10-03

Oil Breather (UB-53 and After, UC-1 and After and (UA-1 thru UA-3, UB-1 thru UB 52 with Kit 114-9006-1 Installed))

STARTING Starting System

CHARTS 00-00

Page 18 Nov 1/13

Exhaust System

OIL

00-00

91

Engine Indicating System

EXHAUST 00-00

79

Ignition System

ENGINE CONTROLS

00-00

78

Title

INTRODUCTION

Consumable Materials/Special Tools and Equipment

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

CHAPTER 04 - AIRWORTHINESS LIMITATIONS TABLE OF CONTENTS SUBJECT

PAGE

INSPECTION/CHECK 04-00-00 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1

04-CONTENTS

Page 1 May 1/10

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

List of Effective Pages CH-SE-SU

PAGE

DATE

04-LOEP

1

May 1/10

04-CONTENTS

1

May 1/10

04-00-00

1

May 1/10

C2

04-LOEP

Page 1 May 1/10

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

04-00-00 00

AIRWORTHINESS LIMITATIONS INSPECTION/CHECK GENERAL At revision C2 a new Airworthiness Limitations Manual was created. Refer to Model 1900 Airliner Series Airworthiness Limitations Manual, P/N 129-590000-133.

04-00-00

Page 1 May 1/10

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CHAPTER 05 - TIME LIMITS/MAINTENANCE CHECKS TABLE OF CONTENTS SUBJECT

PAGE

GENERAL INFORMATION 05-00-00 Inspection/Check. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Inspection Program . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Special Conditions Cautionary Notice . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Time-Limited Inspections . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Time-Limited Major Maintenance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Continuous Inspection Program . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Unscheduled Maintenance Checks - Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

1 1 2 2 2 3 3 3

TIME LIMITED INSPECTIONS 05-10-00 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Inspections . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Chapter 21 - Environmental Systems . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Chapter 22 - Auto Flight . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Chapter 23 - Communications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2 Chapter 24 - Electrical Power . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2 Chapter 25 - Equipment / Furnishing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2 Chapter 26 - Fire Protection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2 Chapter 27 - Flight Controls . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3 Chapter 28 - Fuel System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4 Chapter 31 - Indicating/Recording Systems . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5 Chapter 32 - Landing Gear . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5 Chapter 35 - Oxygen . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7 Chapter 56 - Windows . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8 Chapter 57 - Wings . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8 Chapter 61 - Propeller . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9 Chapter 71 - Power Plant . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9 Chapter 72 - Engine . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9 Chapter 73 - Engine Fuel and Control . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .9 Chapter 79 - Oil . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9

MAJOR MAINTENANCE SCHEDULE 05-11-00 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Schedule . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Chapter 21 - Environmental Systems . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Chapter 23 - Communications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Chapter 24 - Electrical Power . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Chapter 28 - Fuel System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Chapter 32 - Landing Gear . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Chapter 34 - Navigation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Chapter 71 - Power Plant . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Chapter 73 - Engine Fuel and Control . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Chapter 77 - Engine Indicating . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Chapter 79 - Oil . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

05-CONTENTS

1 1 1 1 1 2 2 3 3 3 4 4

Page 1 Nov 1/13

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CHAPTER 05 - TIME LIMITS/MAINTENANCE CHECKS TABLE OF CONTENTS (CONTINUED) SUBJECT

PAGE

SCHEDULED MAINTENANCE CHECKS 05-20-00 Continuous Inspection Program. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Special Conditions Cautionary Notice . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Continuous Inspection General Information . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Inspection Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Discrepancies . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Away-From-Station Requirements . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Continuous Inspection Schedule . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General Information . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

1 1 2 3 4 5 5 5 5

ROUTINE INSPECTION 05-20-01 Inspection Procedures. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Forms Required . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Reference Material . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Inspection Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Routine Inspection Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Airframe . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Power Plant . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Emergency and Survival Equipment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

1 1 1 1 1 2 2 4 4

FIRST 200-HOUR-INTERVAL DETAILED INSPECTION 05-20-02 Detailed Inspection Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Forms Required . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Reference Material . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Inspection Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Detailed Inspection Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Wings . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . FWD Right-hand Center Section . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Power Plant . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General Service Items . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Operational Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

1 1 1 1 1 2 3 9 10 10 11

SECOND 200-HOUR-INTERVAL DETAILED INSPECTION 05-20-03 Detailed Inspection Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Forms Required . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Reference Material . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Inspection Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Detailed Inspection Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Power Plant . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Environmental System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General Service Items . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Operational Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Page 2 Nov 1/13

05-CONTENTS

1 1 1 1 1 2 3 7 8 9

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

CHAPTER 05 - TIME LIMITS/MAINTENANCE CHECKS TABLE OF CONTENTS (CONTINUED) SUBJECT

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THIRD 200-HOUR-INTERVAL DETAILED INSPECTION 05-20-04 Detailed Inspection Procedures. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Forms Required . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Reference Material . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Inspection Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Detailed Inspection Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2 Flight Compartment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3 Cabin Section . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6 Power Plant . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10 General Service Items . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10 Operational Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12

FOURTH 200-HOUR-INTERVAL DETAILED INSPECTION 05-20-05 Detailed Inspection Procedures. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Forms Required . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Reference Material . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Inspection Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Detailed Inspection Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Environmental Systems . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Nose Section . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Power Plant . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General Service Items . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Operational Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

1 1 1 1 1 2 3 4 5 5 6

FIFTH 200-HOUR-INTERVAL DETAILED INSPECTION 05-20-06 Detailed Inspection Procedures. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Forms Required . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Reference Material . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Inspection Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Detailed Inspection Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2 Forward Left Hand Center Section . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3 Main Landing Gear . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3 Nose Landing Gear . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5 Landing Gear Retraction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8 Power Plant . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9 General Service Items . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10 Operational Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12

05-CONTENTS

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CHAPTER 05 - TIME LIMITS/MAINTENANCE CHECKS TABLE OF CONTENTS (CONTINUED) SUBJECT

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SIXTH 200-HOUR-INTERVAL DETAILED INSPECTION 05-20-07 Detailed Inspection Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Forms Required . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Reference Material . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Inspection Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Detailed Inspection Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Aft Fuselage and Empennage . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Cabin Section . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Power Plant . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General Service Items . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Operational Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

1 1 1 1 1 2 3 8 9 9 10

UNSCHEDULED MAINTENANCE CHECKS 05-50-00 Inspection Procedures. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Operation in Areas of High Dust Content . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Operating from Very Soft or Unusual Terrain . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Inspection After Hard or Overweight Landing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . First Level . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Second Level . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Inspection After Encountering Turbulent Air . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . First Level . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Second Level . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Inspection After Lightning Strike . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Engine Inspection After Sudden Stoppage . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Inspection After Heavy Equipment Cargo Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Inspection After Deployment of Landing Gear Above Critical Speed Condition . . . . . . . . . . . . . . . . . . . . . . First Level . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Second Level . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Inspection After Deployment of Flaps Above Critical Speed Condition . . . . . . . . . . . . . . . . . . . . . . . . . . . . First Level . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Second Level . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Inspection in the Event of a Bent Nose Steering Stop . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Inspection After Flight in Airspace with a Low Contamination of Volcanic Ash . . . . . . . . . . . . . . . . . . . . . . Eurocontrol Reference Links . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Engine Manufacturer References . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Avionics Manufacturer References . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

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List of Effective Pages CH-SE-SU

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

TIME LIMITS/MAINTENANCE CHECKS GENERAL INFORMATION INSPECTION/CHECK

05-00-00 00

1. GENERAL WARNING: When an airplane has experienced abnormal landing gear procedures of any type, as a safety precaution, place the airplane on jacks prior to performing any inspection or maintenance. Ensure that all three landing gears are down and locked prior to removing the airplane from the jacks. CAUTION: Jacking of an airplane for the purpose of landing gear operation, inspection, servicing or maintenance should be accomplished within and enclosed building or hangar. In the interest of safety, should it become necessary to jack the airplane in the open, wind velocity in any direction and terrain variations, must be compensated for prior to jacking the airplane. This chapter contains TIME-LIMITED INSPECTIONS, a TIME-LIMITED MAJOR MAINTENANCE SCHEDULE, and a CONTINUOUS INSPECTION PROGRAM along with procedures for UNSCHEDULED MAINTENANCE CHECKS. This program has been developed to enable the owner/operator to accomplish inspections and maintenance on a progressive basis in accordance with 14 CFR Part 91.409 (f) (3). The Hawker Beechcraft approved inspection program contained in the chapter is specifically for the Model 1900D Airliner. Any variation to the inspection program must be approved in writing by the FAA Flight Standards District Office (FSDO), or Airworthiness Authority. The inspection program meets the requirement of both 14 CFR Part 91 and 14 CFR Part 135. NOTE: A flight cycle is defined as: Engine start-up and increase to full or partial power (as required during a normal flight), one landing gear retraction and extension and a complete shutdown. The inspection program in this chapter is based on numbers of flight hours, cycles of operation or calendar time. The basis for calendar-time-limited inspections is the date on the “ORIGINAL STANDARD AIRWORTHINESS CERTIFICATE”, FAA Form No. 8100-2, which is issued with a new airplane. Additionally, Hawker Beechcraft Corporation recommends that operators record the number of cycles experienced on individual components for purposes of complying with inspections based on cycle count. Hobbs meter time or airplane log sheets can be used for determining when inspections and maintenance based on flight hours will be due. However, the method chosen for recording flight hours should remain constant throughout the life of the airplane. The times in this inspection program have been established only as a guideline to give the owner/operator a benchmark from which to begin the program. The service history or fleet experience of a particular operation may indicate that departure from the times in this chapter would be advantageous. If, however, changes to a previously approved program are desired, they must be submitted to the FSDO for approval. All inspections listed in this chapter should be accomplished with reference to this Maintenance Manual and the appropriate supplier maintenance publications. Maintenance information on most of the major components of the airplane is contained in the Model 1900 Airliner Series Component Maintenance Manual.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

A. Inspection Program The Model 1900/1900C Airliner inspection program has been developed to enable an owner/operator to accomplish inspections and maintenance on an on-going basis in accordance with 14 CFR 91.409 (f) (3). A complete inspection cycle is 1,200 hours or 24 months. The inspection cycle is divided into six Detailed Inspections and each inspection cycle is done at 200 hours with each consecutive Detailed Inspection 200 hours after the previous inspection. The Detailed Inspections provide a thorough inspection of specific components and systems and occur at 200-hour intervals. For newly added items, as an example; A new inspection requirement added to the second 200-hour-Interval Detailed Inspection need not be accomplished until the next scheduled second 200-hour-Interval Detailed Inspection, unless otherwise stated.

B. Special Conditions Cautionary Notice Extremely high utilization airplanes and/or airplanes operated in extreme climates may need more frequent inspections for wear, corrosion, and lubrication. The periodic inspections in this chapter should be accomplished until the owner/operator can establish his own inspection periods based on experience or another program which has had prior approval. NOTE: The time periods listed in this chapter do not constitute a guarantee the item will reach the period without malfunction as the aforementioned factors cannot be controlled by the manufacturer.

C. Time-Limited Inspections This subchapter lists items that are subject to a thorough inspection based on flight hours, cycles of operation or calendar time. These TIME-LIMITED INSPECTIONS do not meet the criteria established for more detailed and frequent inspections listed in the CONTINUOUS INSPECTION PROGRAM subchapter. The first TIME-LIMITED INSPECTION of an item must be accomplished not later than the period stated in this subchapter unless prior experience indicates otherwise. Discrepancies noted and corrective action taken during these TIME-LIMITED INSPECTIONS should be recorded in the appropriate airplane records. Requirements added to the TIME- LIMITED INSPECTIONS, TIME-LIMITED MAJOR MAINTENANCE SCHEDULE or the CONTINUOUS INSPECTION PROGRAM, need not be complied with immediately. Unless otherwise directed by relevant Communiqué or Service Bulletin. A new requirement added to the TIME-LIMITED INSPECTIONS need not be complied with until one year from the date the new requirement was published, unless otherwise stated. For example a new inspection requirement added to the second 200-hour-Interval Detailed Inspection need not be accomplished until the next scheduled second 200-hour-Interval Detailed Inspection. A new requirement added to the TIME-LIMITED INSPECTIONS that specifies a 12 month inspection interval may be introduced using a reasonable phase in schedule.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

D. Time-Limited Major Maintenance The subchapter under this heading is a MAJOR MAINTENANCE SCHEDULE. This schedule lists components of the 1900/1900C Airliner aircraft which require periodic major maintenance. The first Major Maintenance of an item must be accomplished not later than the period stated in this subchapter unless prior experience indicates otherwise. The components listed may require complete replacement or major repair based on numbers of flight hours, cycles of operation or calendar time applicable to the particular component. If more frequent checks or servicing of one or more of these components are necessary, these additional requirements will be listed in the CONTINUOUS INSPECTION PROGRAM.

E. Continuous Inspection Program The Hawker Beechcraft recommended CONTINUOUS INSPECTION PROGRAM provides a means of inspecting and maintaining the aircraft on a 50- and 200-hour basis. Routine inspections and servicing are conducted every 50 hours of operation. A Detailed inspection of specific areas and systems of the aircraft is conducted every 200 hours for a period of 1,200 hours. Work sheets are provided at the end of the Routine and each Detailed Inspection to record discrepancies and the corrective action taken. At the end of each 1,200-hour cycle, the owner/operator will have performed a complete inspection of the entire airplane. Although the times of Routine and Detailed inspections may be altered, each item should be accomplished as stated in the CONTINUOUS INSPECTION PROGRAM. A detailed preamble to this subchapter is included and should be read and understood before beginning the CONTINUOUS INSPECTION PROGRAM.

F. Unscheduled Maintenance Checks - Maintenance Practices This subchapter is assembled in Table form to allow a technician to perform checks for damage after operating the aircraft in conditions which could require unscheduled maintenance. Specific conditions, such as lightning strikes, turbulent air penetration and hard landings etc., are included. Inspection instructions are included for each of the conditions listed.

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TIME LIMITS/MAINTENANCE CHECKS TIME LIMITED INSPECTIONS GENERAL

05-10-00 00

1. INSPECTIONS A. Chapter 21 - Environmental Systems ITEM

INSPECTION REQUIREMENTS

1. Sensor, Bleed Air Temperature

Every 5,000 hours replace or perform BLEED AIR TEMPERATURE CHECK (Ref. Chapter 21-10-00).

2. Bleed Air Precooler Bypass Valve

Every 5,000 hours replace or perform BLEED AIR TEMPERATURE CHECK (Ref. Chapter 21-10-00).

3. Bleed Air Precooler-Through Valve

Every 5,000 hours replace or perform BLEED AIR TEMPERATURE CHECK (Ref. Chapter 21-10-00).

4. Bleed Air Pressure Regulator Shutoff Valve

Every 5,000 hours replace or perform BLEED AIR PRESSURE CHECK (Ref. Chapter 21-10-00).

5. Cabin Altitude Warning Pressure Switch System

Perform the CABIN ALTITUDE WARNING PRESSURE SWITCH SYSTEM FUNCTIONAL TEST every 24 months (Ref. Chapter 21-30-07).

B. Chapter 22 - Auto Flight ITEM 1. Autopilot

INSPECTION REQUIREMENTS Annually, perform the autopilot preflight or ground check procedures in the appropriate 1900/1900C AFM Autopilot Supplement. Annually, check autopilot servos for loose or worn mounting hardware and verify that the servo mounts are securely mounted to the airframe. Visually inspect for capstan or cable wear, contamination and proper spool-off. With the autopilot disengaged, operate each control system through its entire range and observe the servo mount for any unusual noise, binding, backlash or other mechanical irregularities.

05-10-00

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

C. Chapter 23 - Communications ITEM

INSPECTION REQUIREMENTS

1. Cockpit Voice Recorder (CVR) Underwater Locator Device (ULD) Test

Test the underwater locator device every 24 months (Ref. Chapter 23 of the Model 1900 Airliner Series Component Maintenance Manual).

2. CVR ULD Battery Replacement

Replace the underwater locator device battery, 72 months after installation or by the expiration date as stated on the battery (Ref. Chapter 23 of the Model 1900 Airliner Series Component Maintenance Manual).

D. Chapter 24 - Electrical Power ITEM 1. Starter-Generator

INSPECTION REQUIREMENTS Replace or overhaul every 1,500 hours.

E. Chapter 25 - Equipment / Furnishing ITEM 1. Emergency Locator Transmitter

INSPECTION REQUIREMENTS Annually, inspect for proper installation, battery corrosion, operation of controls and crash sensor and presence of sufficient signal radiated from the antenna as instructed in Chapter 25-60-00. Replace battery at 50% of life, as stated on the battery, or anytime the transmitter is used more than one cumulative hour.

F. Chapter 26 - Fire Protection ITEM 1. Bleed Air Warning Switches

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05-10-00

INSPECTION REQUIREMENTS Every 5,000 hours perform BLEED AIR WARNING SWITCHES CHECK FOR PROPER ELECTRICAL CONNECTION (Ref. Chapter 26-11-00).

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

G. Chapter 27 - Flight Controls ITEM 1. Flight Controls - Gust Lock Inspection/Replacement

INSPECTION REQUIREMENTS Every 12 months check for Gust Lock P/N 101-590016-5 UA-3; UB-1 through UB-74; UC-1 through UC-174. Check condition of gust lock and that it is in the cockpit available to the crew for installation. Refer to Mandatory Service Bulletin (MSB) 27-3459 for detailed information and recurring requirement.

2. Aileron Trim Tab Control

Perform the AILERON TRIM TAB CONTROL INSPECTION every 5,000 hours (Ref. Chapter 27-10-06).

3. Rudder Trim Tab Control

Perform the RUDDER TRIM TAB CONTROL INSPECTION every 5,000 hours (Ref. Chapter 27-20-06).

4. Elevator Trim Tab Indicator

Perform the ELEVATOR TRIM TAB INDICATOR INSPECTION every 5,000 hours (Ref. Chapter 27-30-08).

5. Flap Flexible Shafts

Replace every 15,000 cycles1 (Ref. Chapter 27-50-03).

6. Flap Motor, Gearbox, Actuators and 90° Drives

Replace or inspect every 10,000 cycles1 (Ref. Chapter 27 of the Model 1900 Airliner Series Component Maintenance Manual).

7. Outboard Flap - Airplanes that have complied with Service Bulletin 27-3158.

Remove flaps and inspect flap attach brackets, roller bearings and attachment hardware for wear every 5,000 hours or five years, whichever comes first (Ref. Chapter 27).

8. Outboard Flap - Airplanes that have not complied with Service Bulletin 27-3158.

Remove flaps and inspect flap attach brackets, roller bearings and attachment hardware for wear every 1,200 cycles1 or one year, whichever comes first (Ref. Chapter 27).

9. Inboard Flap

Remove flaps and inspect flap attach brackets, roller bearings and attachment hardware for wear every 5,000 hours or five years, whichever comes first (Ref. Chapter 27).

10. Aileron Yoke Assembly and Aileron Bellcrank Assembly (UC-1 and After)

Perform the AILERON YOKE ASSEMBLY CHECKS and AILERON BELLCRANK ASSEMBLY REMOVAL AND INSPECTION every 3,000 hours (Ref. Chapter 27-10-02).

11. Aileron Balance Weights Clip Inspection

Perform AILERON BALANCE WEIGHTS CLIP INSPECTION procedure every 3,000 hours (Ref. Chapter 27-10-08).

05-10-00

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL ITEM 12. Elevator System Bobweight Link Assembly Inspection

INSPECTION REQUIREMENTS Perform the BOBWEIGHT LINK ASSEMBLY Inspection procedure every 5,000 hours (Ref. Chapter 27-30-02).

H. Chapter 28 - Fuel System ITEM 1. Fuel System Collector Tank (UC-1 and After)

2. Fuel System Main Fuel Tank at WS 124 thru 130 (UC-1 and After)

INSPECTION REQUIREMENTS Every 12 months perform the FUEL SYSTEM TANK INSPECTION procedure outlined in Chapter 28-11-00 of the Model 1900/1900C Airliner Maintenance Manual or the FUEL TANK INTERNAL INSPECTION procedure outlined in Chapter 28-10-01 of the Model 1900 Airliner Series Corrosion Control Manual. NOTE The removal of the sealant from the main spar forward flange, lower cap and the bulkhead at WS 124 thru 130 is required during the initial inspection. But the removal of the sealant during the recurring 12 month inspections may be skipped for up to 36 months if the fuel system is sterilized using BIOBOR JF at concentrations of 270 PPM or Kathon FP 1.5 at concentrations of 100PPM every six months and is documented in the airplane maintenance records. For application of BIOBOR JF (Ref. Chapter 12-10-00). Every 12 months perform the FUEL SYSTEM TANK INSPECTION procedure outlined in Chapter 28-11-00 of the Model 1900/1900C Airliner Maintenance Manual or the FUEL TANK INTERNAL INSPECTION procedure outlined in Chapter 28-10-01 of the Model 1900 Airliner Series Corrosion Control Manual.

3. Fuel Lines (UC-1 and After)

Inspect wiring and fuel lines for chafing behind the LH and RH nacelle inner fender every 2,400 hours or 12 months, whichever occurs first, as instructed in Chapter 28-21-06.

4. Fuel Bays and Bladders

Inspect for microbiological growth every 4,800 hours or 36 months, whichever occurs first. Clean fuel bays and probes thoroughly.

5. Fuel Level Sensor

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Functional test the Fuel Level Sensors every 4,800 hours or 36 months, Whichever occurs first (Ref. Chapter 28-40-01).

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

I. Chapter 31 - Indicating/Recording Systems ITEM

INSPECTION REQUIREMENTS

1. Flight Data Recorder (FDR) Underwater Locator Device (ULD) Test

Test the Underwater Locator Device every 24 months.

2. FDR ULD Battery Replacement

Replace the Underwater Locator Device battery P/N DK100/120, 72 months after installation or by the expiration date as stated on the battery.

J. Chapter 32 - Landing Gear ITEM 1. Main Gear Assembly, Drag Brace Assembly, Axle and Torque Knees

INSPECTION REQUIREMENTS Replace or inspect every 10,000 cycles1 or five years, whichever comes first (Ref. Chapter 32 of the Model 1900 Airliner Series Component Maintenance Manual). Bushing removal for O. D. corrosion check only required at 10 year intervals.

2. Nose Gear Assembly, Drag Brace Assembly, Axle and Torque Knee

Replace or inspect every 10,000 cycles1 or five years, whichever comes first (Ref. Chapter 32 of the Model 1900 Airliner Series Component Maintenance Manual). Bushing removal for O. D. corrosion check only required at 10 year intervals.

3. Landing Gear and Drag Brace Attach Bolts (Hollow “Lube Type” Bolts)

Replace every 10,000 cycles1 or 5 years whichever comes first (Ref. Chapter 32 of the Model 1900 Airliner Series Component Maintenance Manual).

4. Actuator, Main Gear

AIRIGHT/APPH - Overhaul or replace at 10,000 cycles1 or if leakage past the rod seal exceeds one drop per 25 cycles1. TACTAIR/PHOENIX Controls - (Ref. 05-11-00). FRISBY/TRIUMPH ACTUATION SYSTEMS (Ref. 05-11-00). Perform MAIN LANDING GEAR ACTUATOR END CAP INSPECTION every 1,200 cycles (Ref. Chapter 32-30-10). For new or newly overhauled actuators with records that show the end cap has 8,000 cycles or less, perform the LANDING GEAR ACTUATOR END CAP INSPECTION initially at 8,000 cycles and thereafter at every 1,200 cycles.

5. Airight Main Gear Actuator Shuttle Valve

Perform the MAIN GEAR ACTUATOR SHUTTLE VALVE FUNCTIONAL TEST every 5,000 hours (Ref. Chapter 32-30-10).

05-10-00

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INSPECTION REQUIREMENTS

6. Antiskid System

Complete system functional check should be accomplished at a maximum interval of one year/4,000 hours. Tire flat-spotting or other system difficulties warrant a complete system test.

7. Wheel Speed Transducers

Complete system functional check should be accomplished at a maximum interval of one year/4,000 hours. Tire flat-spotting or other system difficulties warrant a complete system test. Overhaul at 10,000 hours.

8. Hydraulic Line Filter

Inspect filter every 3,000 hours (Ref. Chapter 32-30-00 for detailed inspection).

9. Actuator, Nose Gear

AIRIGHT/APPH - Overhaul including NEW upper end cap or replace with new actuator or newly overhauled actuator with NEW upper end cap every 10,000 cycles1. NOTE At overhaul, end caps returned to service have the actuator serial number stamped between the ports. Only New end caps have “APW” and the end cap serial number stamped on the top of the end cap. The end cap cycles1 must be tracked in addition to the actuator serial number cycles1. AIRIGHT/APPH - Upper end caps with more than 10,000 cycles1 or if there are no records that show the total number of cycles1 on the end cap, perform the NOSE LANDING GEAR ACTUATOR ULTRASONIC INSPECTION every 600 cycles1 (Ref. Chapter 32-30-14). The 600 cycle1 repetitive inspection starts with the receipt of this revision. Overhaul or replace if hydraulic leakage is noted anywhere except for the rod seal. The rod seal is allowed one drop per 25 cycles1 or from the vent hole of the lock indicator switch which is allowed two drops per 25 cycles1.

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K. Chapter 35 - Oxygen ITEM 1. AVOX Systems Inc., formerly Scott Aviation, Altitude Compensated Regulator (Passenger) (Flight Compartment Sidewall) NOTE

INSPECTION REQUIREMENTS Return to AVOX Systems Inc., formerly Scott Aviation, for Functional Test or perform FUNCTIONAL TEST procedure in the Model 1900 Airliner Series Component Maintenance Manual (Ref. Chapter 35-20-03) every five years.

Kit 118-5000-3 adds a crew regulator P/N 118-560005-1. 2. AVOX Systems Inc., formerly Scott Aviation, Altitude Compensated Regulator (Flight Compartment Sidewall)

Return to AVOX Systems Inc., formerly Scott Aviation, for Functional Test every five years.

3. Crew Masks Without Kit 118-5000-3 Installed

Considered on condition by AVOX Systems Inc., formerly Scott Aviation. No overhaul requirements. NOTE Old style mask similar to passenger mask. P/N 249-339-1 AVOX Systems Inc. P/N 249-339-1 Smoke Goggles N/A

4. Crew Masks With Kit 118-5000-3 Installed

AVOX Systems Inc., formerly Scott Aviation, recommends a five year overhaul of this crew mask. P/N 129-380020-1 AVOX Systems Inc. P/N 359-61G12 Smoke Goggles 322-70

5. Passenger Oxygen Masks

For overhaul or replacement (Ref. Chapter 35, Model 1900 Airliner Series Component Maintenance Manual).

6. Diluter Demand Oxygen Masks (Ref. Chapter 35-00-00)

For the Puritan-Bennett only, return to factory every three years for Functional Test.

05-10-00

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL ITEM 7. Oxygen System

INSPECTION REQUIREMENTS Check the condition of the following systems annually: Deploy all cabin oxygen masks and check for oxygen flow every 12 months. OXYGEN SYSTEM LOW PRESSURE TEST - CREW AND AUXILIARY SECTION (Ref. Chapter 35-00-00, OXYGEN - MAINTENANCE PRACTICES). NOTE Auxiliary masks not used in airplane serials UB-23 and After, and UC-1 and After. OXYGEN SYSTEM LOW PRESSURE TEST - CABIN SECTION (Ref. Chapter 35-00-00, OXYGEN MAINTENANCE PRACTICES). CREW OXYGEN MASK AND CONTAINER INSPECTION (Ref. Chapter 35-00-00, OXYGEN MAINTENANCE PRACTICES). PASSENGER OXYGEN MASK AND CONTAINER INSPECTION (Ref. Chapter 35-00-00, OXYGEN MAINTENANCE PRACTICES).

L. Chapter 56 - Windows ITEM 1. Window Frames

INSPECTION REQUIREMENTS Inspect the attach frames for attachment at two years and repeat the inspection every 4,500 hours or annually, whichever occurs first (Ref. Chapter 56-10-00, INSPECTION AND REPAIR OF WINDOW ATTACH FRAMES).

M. Chapter 57 - Wings ITEM 1. Internal Wing Structure (UC-1 and After)

INSPECTION REQUIREMENTS Check for cracks, loose rivets, corrosion, and evidence of sealant deterioration or damage inside all wing inspection areas every 4,800 hours or 36 months, whichever occurs first. Check for nicks, chafes, or breaks in the wing fuel quantity wiring harness every 4,800 hours or 36 months, whichever occurs first. (It is not necessary to remove any spiral wrap that has been installed on the harness to perform this inspection).

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N. Chapter 61 - Propeller ITEM

INSPECTION REQUIREMENTS

1. Hub TBO

Refer to Hartzell Propeller Service Letter 61 for TBO.

2. Propeller Governor

At engine TBO.

3. Blades P/N M10877K

Infinite Life. Service or recondition at hub TBO.

4. Propeller Overspeed Governor

Repair or replace if it fails to pass the functional check or leaks are observed (Ref. Chapter 61-20-00).

O. Chapter 71 - Power Plant ITEM 1. Fuel Purge System

INSPECTION REQUIREMENTS Perform the FUEL PURGE TANK CLEANING procedure every 24 months (Ref. Chapter 71-70-00).

P. Chapter 72 - Engine ITEM

INSPECTION REQUIREMENTS NOTE

A TBO (Time Between Overhaul) recommendation is in no way to be construed as a warranty or engine life proportion basis. The TBO recommendation is based on the projected time for most advantageous initial overhaul. The individual operator's experience may indicate a departure in either direction from the recommended TBO for the particular operation. 1. Engine TBO

Refer to Pratt and Whitney Service Bulletin No. 13003 for overhaul time limits.

Q. Chapter 73 - Engine Fuel and Control ITEM 1. Flammable-Liquid-Carrying Hoses

INSPECTION REQUIREMENTS Replace every five years. Replace when cracked, leaking or deteriorated.

R. Chapter 79 - Oil ITEM

INSPECTION REQUIREMENTS

1. Engine Chip Detectors

Perform the MAGNETIC DRAIN PLUG INSPECTION every 100 hours (Ref. Chapter 79-00-00).

2. Standard Engine Oil Hose, P/N 330996F-8-0095, Oil Drain LH Engine

Replace every five years

3. Standard Engine Oil Hose, P/N 330997F-8-0111, Oil Drain RH Engine

Replace every five years

05-10-00

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INSPECTION REQUIREMENTS

4. Standard Engine Oil Hose, P/N 330997F-12-0290, Oil Cooler Inlet

Replace every five years

5. Standard Engine Oil Hose, P/N 330997F-12-0414, Oil Cooler Outlet

Replace every five years

1

A flight cycle is defined as: Engine start-up and increase to full or partial power (as required during normal flight) one landing gear retraction and extension and a complete shutdown.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

TIME LIMITS/MAINTENANCE CHECKS MAJOR MAINTENANCE SCHEDULE GENERAL

05-11-00 00

1. SCHEDULE NOTE: Items not listed are to be repaired or replaced when necessary. If items are worn, inoperative, inaccurate, intermittent and are not repairable through normal maintenance practices, they must be overhauled or replaced.

A. Chapter 21 - Environmental Systems ITEM

INSPECTION REQUIREMENTS

1. Cabin Temperature Controller (Blower Motor)

On condition. Repair or replace if improper operation is observed.

2. Air Cycle Machine Bypass Valve

On condition. Repair or replace if improper operation is observed.

3. Vent Blower Assembly (Evaporator)

On condition. Repair or replace if improper operation is observed.

4. Valve, Ejector Bypass

On condition. Repair or replace if improper operation is observed.

5. Valve, Pneumatic Regulator/Relief

On condition. Repair or replace if improper operation is observed.

6. Air-Conditioner Condenser Coil

On condition. Repair or replace if improper operation is observed.

7. Air-Conditioner Evaporator Coil

On condition. Repair or replace if improper operation is observed.

8. Air-Conditioner Condenser Blower

On condition. Repair or replace if improper operation is observed.

9. Air Cycle Machine - Heat Exchanger

On condition. Clean Heat Exchanger at intervals noted in Hamilton Sundstrand Service Bulletin SB21-2095.

B. Chapter 23 - Communications ITEM 1. Cockpit Voice Recorder

INSPECTION REQUIREMENTS On condition. Repair or replace if improper operation is observed (Ref. Chapter 23 of the Model 1900 Airliner Series Component Maintenance Manual).

C. Chapter 24 - Electrical Power ITEM 1. Airplane Battery

INSPECTION REQUIREMENTS On condition. Inspect and service (Ref. Chapter 24 of the Model 1900 Airliner Series Component Maintenance Manual).

05-11-00

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D. Chapter 28 - Fuel System ITEM

INSPECTION REQUIREMENTS

1. Fuel Cross-Transfer Valve

On condition. Repair or replace if leaks or improper operation is observed.

2. Firewall Shutoff Valve

On condition. Repair or replace if leaks or improper operation is observed.

E. Chapter 32 - Landing Gear ITEM 1. Actuator, Nose Gear

INSPECTION REQUIREMENTS On condition. Repair or replace if hydraulic leakage is noted anywhere except for the rod end seal which is allowed one drop per 25 cycles1 or from the vent hole of the lock indicator switch which is allowed 2 drops per 25 cycles1. AIRIGHT/APPH - (Ref. 05-10-00). TACTAIR/PHOENIX CONTROLS - On condition. Repair or replace if hydraulic leakage is noted anywhere except for leakage past the rod seal exceeding one drop per 25 cycles1 or leakage from the vent hole of the lock indicator switch exceeding two drops per 25 cycles1.

2. Actuator, Main Gear

AIRIGHT/APPH - (Ref. 05-10-00). TACTAIR/PHOENIX CONTROLS - On condition. Repair or replace if hydraulic leakage is noted anywhere except for leakage past the rod seal exceeding one drop per 25 cycles1 or leakage from the vent hole of the lock indicator switch exceeding two drops per 25 cycles1. FRISBY/TRIUMPH - On condition. Repair or replace if hydraulic leakage is noted anywhere except for leakage past the rod seal exceeding one drop per 25 cycles1 or leakage from the vent hole of the lock indicator switch exceeding two drops per 25 cycles1.

3. Main Gear Brake Master Cylinder

On condition. Repair or replace if leaks or improper operation is observed.

4. Hydraulic Accumulator Assembly

On condition. Repair or replace if leaks or improper operation is observed.

5. Hydraulic Landing Gear Service Valve.

On condition. Repair or replace if leaks or improper operation is observed.

6. Hydraulic Power Pack Assembly

On condition. Repair or replace if leaks or improper operation is observed.

7. Hydraulic Landing Gear Hand Pump

On condition. Repair or replace if leaks or improper operation is observed.

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INSPECTION REQUIREMENTS

8. Main Gear Brakes

On condition. (Ref. Chapter 32-40-00) of the Model 1900/1900C Airliner Maintenance Manual for limits.

9. Power Steering Actuator

On condition. Repair or replace if leaks or improper operation is observed.

10. Power Steering Pump and Motor Assembly

On condition. Repair or replace if leaks or improper operation is observed.

11. Main Wheel

For inspection and repair (Ref. Chapter 32 of the Component Maintenance Manual). For Inspection and Overhaul Schedules refer to the Aircraft Braking Systems (ABS) Wheel Service Letter GS-SL-36.

12. Nose Wheel

For inspection and repair (Ref. Chapter 32 of the Component Maintenance Manual).

13. Wheel Bearing

Inspect and lubricate at tire change. (Bearing P/N 13889, should be replaced at 900 hours ONLY when installed on the main wheel assembly).

14. All Landing Gear Hoses.

On condition. Replace when cracked, leaking or deteriorated.

F. Chapter 34 - Navigation ITEM

INSPECTION REQUIREMENTS

1. Electronic Flight Display (EFD-74)

On condition. All exchange and repaired units have display brightness test routinely performed. If brightness seems questionable, refer to maintenance section (523-0772698) of Collins EHSI-74 Instruction book (523-0772693).

2. Electronic Horizontal Situation Indicator (EHSI-74) CRT

On condition

G. Chapter 71 - Power Plant ITEM 1. Engine Vibration Isolators

INSPECTION REQUIREMENTS On condition (Ref. Chapter 71 of the Model 1900 Airliner Series Component Maintenance Manual).

H. Chapter 73 - Engine Fuel and Control ITEM 1. Engine-Driven Fuel Boost Pump

INSPECTION REQUIREMENTS On condition. Repair or replace if leaks or if improper operation is observed.

05-11-00

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

I. Chapter 77 - Engine Indicating ITEM 1. Fuel Flow Transmitter

INSPECTION REQUIREMENTS On condition. Repair or replace if leaks or if improper operation is observed.

J. Chapter 79 - Oil ITEM

INSPECTION REQUIREMENTS

1. Oil Cooler

On condition. Replace when contaminated.

2. Teflon P/N 124J003-8CR-0095, Oil Drain Hose LH Engine

On condition. Replace when cracked, leaking or deteriorated.

3. Teflon P/N 124J002-8CR-0111, Oil Drain Hose RH Engine

On condition. Replace when cracked, leaking or deteriorated.

4. Teflon P/N 124J002-12CR-0290, Oil Inlet Cooler

On condition. Replace when cracked, leaking or deteriorated.

5. Teflon P/N 124J002-12CR-0414, Oil Outlet Cooler

On condition. Replace when cracked, leaking or deteriorated.

1

A flight cycle is defined as: Engine start-up and increase to full or partial power (as required during normal flight) one landing gear retraction and extension and a complete shutdown.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

TIME LIMITS/MAINTENANCE CHECKS SCHEDULED MAINTENANCE CHECKS CONTINUOUS INSPECTION PROGRAM

05-20-00 00

1. GENERAL The owner or operator is ultimately responsible for maintaining the airplane in an airworthy condition, including compliance with all applicable Airworthiness Directives as specified in Title 14 of the Code of Federal Regulations (CFR) Part 39, or as specified by the directives of the national aviation authorities. The owner or operator should select only qualified personnel to maintain the airplane, and ensure that the airframe and power plant mechanic inspecting the airplane has access to all necessary manuals and service information as well as to an approved inspection guide. It is further the responsibility of the owner or operator to ensure that the airplane is inspected in conformity with the requirements covered in 14 CFR Part(s) 91.409 (f) (3), 121.367, 125.247 or 135.419 of the Code of Federal Regulations or as specified by the directives of the national aviation authorities. These CFR Parts cover the requirements concerning approved airplane inspection programs. Hawker Beechcraft Corporation has prepared this Continuous Inspection Program to assist the owner or operator in meeting the foregoing responsibilities. It is the responsibility of the owner or operator to obtain specific FAA (or national aviation authority), approval for the continuous inspection program the owner or operator adopts. NOTE: When warranted by service experience or engineering recommendations, an approved maintenance program, including the inspection intervals, may be changed at any time with prior notification and approval of the local FAA Flight Standards District Office (FSDO) or as required by the national aviation authority. Hawker Beechcraft Corporation publishes recommended inspection requirements and maintenance schedules for the airframe of your airplane. Remember that maintenance requirements and schedules for some supplier furnished components, such as engines, propellers, avionics, cabin heaters, and other airplane equipment, are separately stated in their respective supplier maintenance manuals. Have your maintenance personnel review the equipment installed on your airplane and ensure that current, up-to-date supplier maintenance publications and manuals are available and all required maintenance is scheduled and performed. This Continuous Inspection Program is provided to enable the owner/operator to inspect and maintain the airplane on a continuous basis. Included in the program are a Routine Inspection and six Detailed Inspections. A sequence of conducting the program along with suggested times are discussed later. The times and sequence are recommendations and may be altered to suit a particular operation. While this program may be used as an outline, detailed information of the many systems and components in the airplane will be found in the various chapters/sections of the maintenance manual and the pertinent supplier publications. It is also recommended that reference be made to the applicable maintenance handbooks, service instructions, applicable FAA (or national aviation authority) Regulations, Publications, and supplier's specifications for torque values, clearances, settings, tolerances, and other requirements. This program is not intended to be all-inclusive, for no such program can replace the good judgement of a certified airframe and power plant mechanic in the performance of his duties.

05-20-00

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NOTE: In addition to the inspections prescribed by this schedule, the altimeter instrument and static system and all ATC transponders MUST be tested and inspected at 24 month intervals or anytime the system is opened in compliance with the requirements specified in 14 CFR Part(s) 91.411 and 91.413 or as specified by the directives of the national aviation authority. Information contained herein is applicable to all Model 1900/1900C Airliner airplanes except where differences are indicated by serial number effectivity.

A. Special Conditions Cautionary Notice Extremely high-utilization airplanes and/or airplanes operated in extreme climates may need more frequent inspections for wear, corrosion, and lubrication. Under these conditions, the items listed in this program should be accomplished as outlined until the owner/operator can establish his own inspection periods based on experience or another program which has had prior approval. Engine power and performance runs should be tailored to each operation to achieve reliable, cost-effective maintenance. Depending on the maintenance performed and components replaced, a GROUND PERFORMANCE CHECK may be required in lieu of the normal Inspection Run. Refer to the applicable maintenance procedures. NOTE: The time periods in this schedule do not constitute a guarantee the item will reach the period without malfunction as the aforementioned factors cannot be controlled by the manufacturer. WARNING: Use only genuine Hawker Beechcraft Corporation or Hawker Beechcraft Corporation approved parts obtained from Hawker Beechcraft Corporation approved sources, in connection with the maintenance and repair of Hawker Beechcraft airplanes. Genuine Hawker Beechcraft Corporation parts are produced and inspected under rigorous procedures to ensure airworthiness and suitability for use in Hawker Beechcraft airplane applications. Parts purchased from sources other than Hawker Beechcraft Corporation, even though outwardly identical in appearance, may not have had the required tests and inspections performed, may be different in fabrication techniques and materials, and may be dangerous when installed in an airplane. Salvaged airplane parts, reworked parts obtained from non-Hawker Beechcraft Corporation approved sources, or parts, components, or structural assemblies, the service history of which is unknown or cannot be authenticated, may have been subjected to unacceptable stresses or temperatures or have other hidden damage, not discernible through routine visual or the usual nondestructive testing techniques. This may render the part, component or structural assembly, even though originally manufactured by Hawker Beechcraft Corporation unsuitable and unsafe for airplane use. Hawker Beechcraft Corporation expressly disclaims any responsibility for malfunctions, failures, damage or injury caused by use of non-Hawker Beechcraft Corporation approved parts.

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B. Continuous Inspection General Information (1) Purpose and Use The Hawker Beechcraft Corporation recommended Continuous Inspection Program is provided to maintain the Model 1900/1900C Airliner airplanes that are utilized by the owner/operator on a continuous basis. Should the owner/operator elect to use the Hawker Beechcraft Corporation recommended program, the complete program must be accomplished at least one time every 24 calendar months. The complete inspection program for each airplane is divided into several parts consisting of a Routine Inspection of the airplane every 50 hours of service time and a Detailed Inspection of a portion of the airplane every 200 hours of service time, thus providing a complete inspection of the airplane every 1,200 hours. The 50-hour or 200-hour interval between performance of the procedures here must not be exceeded by more than 10%. Any extension of either the 50-hour or 200-hour interval must be subtracted from the following 50-hour or 200-hour interval as appropriate, with no time extension permitted. This method will provide greater availability of the airplane during normal operating hours without sacrificing the quality desired during maintenance and inspection periods. (2) Definitions The terminology pertaining to the inspection procedures and their use as explained in this manual are in accordance with Code of Federal Regulations, Parts 1, 43 and 91, issued by the Federal Aviation Administration. These terms are defined as follows: Continuous Inspection - A continuous inspection is a continuing airworthiness inspection of an airplane and its various components and systems at scheduled intervals in accordance with procedures prescribed by the Administrator of the Federal Aviation Administration. Detailed Inspection - Detailed inspection consists of a thorough examination of the appliances, the airplane and components and systems with such disassembly as necessary. Flight Time - Flight time shall mean the total time from the moment the airplane first moves under its own power for the purpose of flight until the moment it comes to rest at the next point of landing. (Block-to-block time.) Maintenance - Means inspection, overhaul, repair, preservations, and the replacement of parts, but excludes preventive maintenance. Pilot in Command - Pilot in command shall mean the pilot responsible for the operation and safety of the airplane during the time defined as flight time. Preventive Maintenance - Means simple or minor preservative operations and the replacement of small standard parts not involving complex assembly operations. Routine Inspection - Routine inspection consists of visual examination or check of the appliances, the airplane and its components and systems insofar as is practicable without disassembly. Time in Service - Time in service, as used in computing maintenance and inspection time records, is the time from the moment the airplane leaves the ground until it touches the ground at the end of the flight.

05-20-00

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(3) Forms and Records The forms and records used with the Continuous Inspection Program consist of a Routine Inspection form, six separate 200-Hour Interval Detailed Inspection forms, Continuous Inspection Work Sheets and In-Flight Work Sheets. Routine Inspection Work Sheet - This Work Sheet lists the airplane components which are to be checked at each 50-flight-hour interval of service time. At each 200-Hour Interval Inspection, this form will be completed in addition to the 200-Hour Interval Detailed Inspection form. 200-Hour-Interval Detailed Inspection Form - Six separate and individual 200-Hour Interval Detailed Inspection forms are used with the Continuous Inspection Program. Each form covers only one portion of the airplane and is designated as 1st, 2nd, 3rd, 4th, 5th or 6th 200-Hour Interval Detailed Inspections. Completion of the 6th of the 200-Hour Interval Detailed Inspection will provide a complete airworthiness inspection of the airplane. Continuous Inspection Work Sheet - This form is used in conjunction with each of the inspection forms to provide a list of all discrepancies which are found during the inspection and their corrective action. In-Flight Work Sheet - Copies of this Work Sheet are to be kept in the airplane and will be used by the pilot in command to list any discrepancy which occurs during a flight. When the flight is completed, this form will then be forwarded to the Maintenance Shop for proper disposition. If the Continuous Inspection Program is discontinued, written notification must be sent to the local FAA Flight Standards district office or as specified by the national aviation authority.

C. Inspection Procedures (1) Routine Inspections (50-Hour Intervals) A Routine Inspection of the airplane shall be conducted each 50 hours of time the airplane is in service. This inspection consists mainly of a visual inspection of the major components of the airplane for general condition. This inspection may be conducted by persons qualified to do preventative maintenance. Refer to Chapter 5-20-01 for complete instructions to conduct routine inspections. (2) Detailed Inspections (200-Hour Intervals) Six separate Detailed Inspections of the airplane are required to accomplish one complete inspection. Only a portion of the airplane's components or systems are inspected at each 200-hour interval, thus accomplishing a complete inspection of the airplane once every 1,200 hours of time in service. Items requiring attention at periods of less than 1,200 hours are duplicated on the appropriate Detailed Inspection form. These inspections are to be conducted by the Crew Chief in Charge of Maintenance, or by a properly qualified mechanic under their supervision. Refer to Chapters 5-20-02, 5-20-03, 5-20-04, 5-20-05, 5-20-06, and 5-20-07 for complete instructions to conduct the 1st through 6th Detailed Inspections.

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D. Discrepancies Discrepancies found on the airplane during an inspection must be listed on the Continuous Inspection Worksheet. The discrepancy must be entered on the worksheet, and the corrective action which is taken must be noted. If more than one line is required to state the discrepancy, as many entry spaces as are necessary may be used. The same method should be used for corrective action explanations which require more than one line, except the bottom lines of the extra spaces will be used. Each separate entry on the sheet will be numbered in the ITEM blocks 1, 2, 3, 4, etc. As many worksheets as necessary will be used to list all discrepancies with the entry numbers in the ITEM block continuing in sequence on each of the additional pages. All discrepancies listed must be corrected before the worksheet is routed to the airplane file. Discrepancies that affect the airworthiness of the airplane will require the necessary corrective action to be accomplished before the airplane is returned to service. Discrepancies that do not affect the airworthiness of the airplane may, at the discretion of the maintenance crew chief, be carried over to the next inspection period. All discrepancies thus carried over will be retained in the Shop File until corrected, and will also be reflected on the Shop Status and Scheduling Board. Discrepancies which occur during a flight must be entered on the In-Flight Worksheet by the pilot in command or other responsible person. At the end of the flight this worksheet is then submitted to the Transportation Department Manager.

E. Away-From-Station Requirements Away-From-Station Inspection. If the airplane is to be away from the home location at the time an inspection is due, the pilot in command of the flight will take with him all forms which will be required for the inspection and a copy of this manual. The detailed inspection will be conducted or supervised by one of the following: 1. A certified airframe repair station. 2. An appropriately rated certified mechanic with inspection authorization. The results of the inspection will be noted on the proper forms which are then brought back to the home location. The pilot will be responsible for all inspection forms and work sheet entries with inspectors and/ or mechanics signature and identification. Away-From-Station Discrepancies. Discrepancies affecting the airworthiness of the airplane, when the airplane is away from the local station, will be corrected by one of the following: 1. A certificated airframe repair station. 2. An appropriately rated certified mechanic. The discrepancy and the corrective action taken is to be listed on the In-Flight Worksheet. The pilot will be responsible for all worksheet entries with mechanic's and/or inspector's signatures and identification.

2. CONTINUOUS INSPECTION SCHEDULE A. General Information The Continuous Inspection Schedule Table on the following pages, lists the major components of the airplane which require periodic inspections. For maintenance items, see the Major Maintenance Schedule list in Chapter 5-11-00.

05-20-00

Page 5 May 1/11

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Table 1 Continuous Inspection Schedule Inspection

RTN

1st

2nd

3rd

4th

5th

6th

X

X

X

X

X

X

X

X

X

X

X

X

A. General Service Bulletins, Compliance Check

X

Switches, Knobs and Circuit Breakers

X

Work Sheet Discrepancies

X

B. Chapter 11 - Placards Placards C. Chapter 12 - Lubrication and Servicing Airplane Lubrication Oxygen System Pressure

X

D. Chapter 21 - Environmental Air Cycle Machine, Fog Nozzle and Filter

X

X

Air Cycle Machine Oil Change

X

Bleed Air Overpressure Check

X

X

X

X

X

X

Bleed Air Overtemperature Check

X

X

X

X

X

X

Blower, Air Conditioning Condenser

X

X

X

X

Compressor Drive Belts

X

X

Compressor Drive Quill Shaft

X

X

Compressor, Refrigerant Condenser, Air Conditioning

X X

X

Evaporator Blower

X

Filter, Pressure Equalization Brake Reservoir

X

Filter, Pressurization Controller Filter, Evaporator

X

X X

X

X

X

Lines and Service Valves, Refrigerant

X

X

X

Outflow Valves

X

Pneumatic Relay Filter

X

Pressurization Controller

X

Pressurization System

X

X

X

Pressurization System Drain Valve

X

X

X

Refrigerant Level Refrigerant Lines and Service Valves

Page 6 May 1/11

05-20-00

X X

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Table 1 Continuous Inspection Schedule (Continued) Inspection

RTN

1st

2nd

Temperature Controller and Filter

X

Vapor Cycle System

X

3rd

4th

5th

X

6th

X

E. Chapter 23 - Communications Radio Equipment

X

Static Dischargers

X

X

F. Chapter 24 - Electrical Power Electrical Equipment, Flight Compartment

X

Electrical Wiring and Equipment, Cabin

X

Electrical Wiring and Equipment, Power Plant

X

Electrical Wiring and Equipment, Wing

X

Electrical, Main and Nose Gear

X

External Power

X

X

X

X

X

X

Inverter Power Relays

X

X

X

X

X

X

Starter-Generator, Brushes

X

X

X

Starter-Generator, Operational Check

X

X

X

X

X

G. Chapter 25 - Equipment and Furnishings Cleanliness, Flight Compartment and Cabin

X

Emergency and Survival Equipment

X

Seat Belts

X

Seat Tracks

X

Seats

X

Underwater Locator Device (ULD)

X

H. Chapter 26 - Fire Protection Fire Detectors

X

Fire Extinguisher, Engine

X

I. Chapter 27 - Flight Controls Aileron and Tab

X

Aileron Quadrant

X

Control Column

X

Flaps and Actuators

X

Flaps, Motor and Drives and Flex Shafts Flight Control Components, Cables and Pulleys

X X

X

05-20-00

X

Page 7 May 1/11

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Table 1 Continuous Inspection Schedule (Continued) Inspection Flight Controls

RTN

1st

X

X

2nd

3rd

4th

5th

X

Rudder Pedals and Arms

6th X

X

Visual Damage, Control Surfaces

X

Visual Damage, Fuselage, Empennage, Wings

X

J. Chapter 28 - Fuel Fuel Boost Pumps

X

Fuel Filler Cap

X

X

X

Fuel Filters and Screens

X

Fuel Probes

X

X

Fuel Purge System

X

Fuel Tanks, Vents, Valves and Pumps

X

Fuel Valve, Antisiphon

X

Fuel Valve, Cross-Transfer

X

X

X

Fuel Valve, Firewall Shutoff

X

X

X

Integral Fuel Tank

X

X

K. Chapter 30 - Ice and Rain Protection Deicer, Propeller

X

X

X

X

X

Deicer Boots, Empennage

X

Deicer Boots, Operational Check

X

Deicer Boots, Wing

X

X

X

Deicer Distributor Valve Pitot and Stall Warning Heat

X X

Propeller Deicer

X

Surface Deicers

X

X

X

X

X

X

Actuator, Main Gear

X

Actuator, Nose Gear

X

Brake Deicer

X X

Brakes, Antiskid

X X

X X

Drag Brace

X

Drag Leg, Main Gear

X

Page 8 May 1/11

05-20-00

X X

L. Chapter 32 - Landing Gear

Brakes

X

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Table 1 Continuous Inspection Schedule (Continued) Inspection

RTN

1st

2nd

3rd

4th

5th

Landing Gear, Doors and Linkage

X

Landing Gear, Emergency Extension

X

Landing Gear, Hoses

X

Landing Gear, Hydraulic Power Pack

6th

X

Landing Gear, Hydraulic Power Pack Filters and Screens

X

Landing Gear, Power Pack, Hydraulic Fluid Level Sensor

X

Landing Gear, Position Indicators

X

Landing Gear, Retract Mechanism

X

Landing Gear, Safety Switch

X

Landing Gear, Struts

X

Landing Gear, Warning Horn

X

Filter, Hydraulic System Bleed Air

X

X

Orifice, Pressure Equalization

X

Power Steering System Filter

X

X

Power Steering Actuator, Pump and Motor

X

Shimmy Damper (if installed)

X

X

Shock Struts, Main and Nose Landing Gear

X

X

Steering Linkage

X

Hoses, All

X

X

M. Chapter 33 - Lights Lights, All

X

N. Chapter 34 - Navigation & Pitot Static Electronic Flight Instrument System (EFIS) (if installed)

X

X

X

X

X

X

Pitot and Static System

X

X

X

X

X

X

Plumbing and Wiring, Instrument

X

Static Ports, Aft Fuselage

X

Static Source, Alternate

X

Hoses, All

X

O. Chapter 37 - Vacuum Filter, Instrument Air

X

Filter, Vacuum Regulator Valve

X

Vacuum System

X X

X

X

05-20-00

X X

X

X X

Page 9 May 1/11

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Table 1 Continuous Inspection Schedule (Continued) Inspection

RTN

1st

X

X

2nd

3rd

4th

5th

6th

P. Chapter 52 - Doors Access Panels, Security and Attachment

X

X

Avionics Doors, Fasteners, and Seals

X

Door, Nose

X

Doors, Entrance - Cargo - Emergency

X

Q. Chapter 53 - Fuselage Control Cable Seals, Aft Fuselage

X

Control Cable Seals, Cabin

X

Drain Valves, Belly

X

Instrument Panel Plumbing and Wiring

X

Plumbing, Aft Fuselage

X

Skin, Fuselage

X

Skin, Nose Section

X

Structure, Cabin

X

Structure, FWD L/H Center Section Structure, FWD R/H Center Section

X X

Structure, Nose

X

R. Chapter 55 - Stabilizers Drains, Ventral Fin and Aft Fuselage

X

Skin, Aft Fuselage and Empennage

X

Structure, Aft Fuselage and Empennage

X

S. Chapter 56 - Windows Side Windows

X

Windows, Cabin

X

Windows, Flight Compartment

X

Windshield Tab

X

Windshield Weather Seal

X

X

X

X X

X

T. Chapter 57 - Wings Plumbing and Wiring, FWD R/H Center Section

X

Plumbing, FWD LH Center Section

X

Plumbing, Leading Edge and Nacelle

X

Plumbing, Wing

X

Page 10 May 1/11

05-20-00

X

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Table 1 Continuous Inspection Schedule (Continued) Inspection

RTN

1st

Skin, Wing

X

Structure, Wing

X

2nd

3rd

4th

5th

6th

U. Chapter 61- Propellers Autofeather Relays, Dump Solenoids, and Pressure Switches

X

X

X

Propeller Governor

X

X

X

Propeller Synchrophaser

X

X

X

Propellers

X

X

V. Chapter 71 - Power Plant Accessories, Engine

X

Auto-Ignition

X

Condition Levers

X

X

Drain Plugs, Power Plant

X

Engine Chip Detector Cleaning

X

Engine Control Levers X

X

Engine Cowling

X

X

Engine Fuel Pump Filter and Screens

X

Engine Ground Performance Check

X

Engine Ground Inspection Run

X

Engine Mount Truss X

Engine Oil Tank

X

X

X

X

X

X

X

X

X

X

X

X

Engine Oil Temperature and Pressure

X

Engine Vibration Isolator Mounts

X X

X

Exhaust System

X

X

X X

X

X

X

X

X

X

X X

X

Fuel Nozzle

X

Hoses, All

X

Igniter Plugs

X

X

Engine Oil Cooler

Filter, Engine Oil

X

X

Engine Controls

Engines

X

X

X

Ignition Exciter

X

Induction System

X

X

X

05-20-00

X

X X

X

X

Page 11 May 1/11

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Table 1 Continuous Inspection Schedule (Continued) Inspection

RTN

Inertial Anti-Icer Oil Cooler

3rd

X X

Plumbing, Power Plant

Page 12 May 1/11

2nd X

X

P3 Air Filter

1st

05-20-00

X

4th

5th

6th

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

TIME LIMITS/MAINTENANCE CHECKS ROUTINE INSPECTION INSPECTION PROCEDURES

05-20-01

00

1. GENERAL A. Forms Required (1) Model 1900/1900C Airliner Routine Inspection. (2) Continuous Inspection Worksheet. NOTE: The Model 1900/1900C Airliner Routine Inspection form will be used for all Routine Inspections whether the inspections are conducted at 50-Hour intervals or in conjunction with a 200-Hour Detailed Inspection.

B. Reference Material (1) Model 1900/1900C Airliner Maintenance Manual P/N 114-590021-7. (2) Model 1900 Airliner Series Component Maintenance Manual P/N 114-590021-11. (3) Model 1900 Airliner Wiring Diagram Manual P/N 114-590032-3 (UA-1 and After); 114-590021-13 (UB-1 and After); 114-590021-61 (UC-1 and After).

C. Inspection Procedures (1) Fill out the heading on each form in its entirety. (2) The mechanic checks each item on the inspection form and initials the form in the spaces provided. (3) List all discrepancies found during the inspection on the Continuous Inspection Worksheet. (4) When the inspection is complete, the mechanic, crew chief and the Quality Control Inspector will sign the INSPECTION COMPLETED block at the end of the inspection sheet. NOTE: Any repairs made during the Routine Inspection will be noted on the Continuous Inspection Worksheet and attached to the completed Routine Inspection Form. Maintenance, other than preventive maintenance, will be signed off by the mechanic and Quality Control Inspector. Only the inspection mechanic’s initials will be required on work classified as preventive maintenance. Corrosion detected while performing this routine inspection may be treated in accordance with Chapter 20-09-00.

05-20-01

Page 1 Feb 1/10

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

2. ROUTINE INSPECTION PROCEDURES Inspection Date___________________________________________ Airplane Serial___________________ Airframe______________________ Cycles____________ LH Engine Hrs_____________ Cycles____________ RH Engine Hrs______________ Cycles____________ LH Engine Power Module S/N_______________________________________________ RH Engine Power Module S/N_______________________________________________ LH Engine Gas Generator S/N_______________________________________________ RH Engine Gas Generator S/N_______________________________________________

A. Airframe Model 1900/1900C Airliner Series Routine Inspection Complete Steps 1 thru 17. (1) All in-flight work sheet discrepancies cleared. (2) Check Service Bulletins for compliance as required. (3) Check cabin and flight compartment for cleanliness and visual damage. (4) Check fuselage, empennage and wings for scratches, paint blistering, corrosion, missing fasteners and visual damage. (5) Check all control surfaces for security, scratches, paint blistering, corrosion and visual damage. (6) Check all access plates for presence, security, scratches, paint blistering, corrosion and visual damage. (7) Check main and nose landing gear and shock struts for damage, corrosion, attachment, correct inflation and leaks. Deflate struts and check fluid level if leaks are evident. Check for leaks in hydraulic retraction and power steering system, if installed. (8) Inspect the nose landing gear shimmy damper (on airplanes without power steering only) for leaks, corrosion, correct fluid level and security of attachment. Service if necessary (Ref. Chapter 32-20-00). (9) Inspect brakes for wear (Ref. Chapter 32-40-00). Check brakes for security of mounting and leaks. Check brake reservoir fluid level and fill (Ref. Chapter 12-10-00). (10) Check oxygen system pressure (Ref. Chapter 12-10-00). (11) Check tires for wear, deterioration and correct inflation (Ref. Chapter 12-10-00).

Page 2 Feb 1/10

05-20-01

Mechanic

Inspector

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Model 1900/1900C Airliner Series Routine Inspection (Continued) Complete Steps 1 thru 17.

Mechanic

Inspector

(12) Visually check all lights for illumination. Replace all non-illuminating bulbs, flashtubes and/or broken lenses etc. (13) Check switches, knobs and circuit breakers for looseness and operation. WARNING When testing the heater function of the pitot mast, use extreme caution as the surface can cause burns. Prolonged operation of the anti-ice function in still air could damage the pitot mast; therefore, it is imperative that the pitot heat be turned off immediately after the heat test. (14) Check pitot heaters for operation by watching for a loadmeter needle deflection when the control is switched on with minimum electrical load established and LH and RH annunciators extinguished. Clean as necessary. WARNING When testing the anti-icing function of the transducer vane and mounting plate, use extreme caution as either surface can cause burns. Prolonged operation of the anti-ice function in still air could damage the transducer; therefore, it is imperative that the transducer heat be turned off immediately after the heat test. (15) Check the stall warning heating system operation (Ref. the Pilot’s Operating Handbook). The heat system may be checked by placing the hand near, but not touching, the stall warning vane and mounting plate. (16) Check all flight controls and tabs for freedom of operation, scratches, paint blistering, corrosion and security of attachment. (a) Check for operation of the yaw damper by pressing the Yaw Damp test switch forward or (if installed) press and hold the YD ENG annunciator/test button (Ref. Chapter 27-21-00). (17) Inspect the windshield wiper blades for deterioration, cuts, etc.

05-20-01

Page 3 Feb 1/10

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

B. Power Plant Model 1900/1900C Airliner Series Routine Inspection Complete Steps 1 thru 6

Mechanic LH

Inspector

RH

(1) Visually check propeller for nicks, erosion and damage. If damage exists, (Ref. Chapter 61-00-00 of the Model 1900 Component Maintenance Manual). (2) Check engines through front cowl, cowl doors and access openings for fuel, oil and exhaust leaks and damage. (3) Visually check oil coolers for leaks and obstruction of air flow. CAUTION Do not place the power levers into reverse unless the engine is running. (4) Check all engine controls for freedom of operation. (5) Check oil tank for correct fluid level (Ref. Chapter 12-10-00). (6) Check engine cowling and access plates for security of attachment, scratches, paint blistering, corrosion, missing fasteners and damage.

C. Emergency and Survival Equipment Model 1900/1900C Airliner Series Routine Inspection Complete Steps 1 thru 5

Mechanic

Inspector

(1) Check first aid kit for security of attachment, integrity of seal and expiration date. (2) Check passenger briefing cards for cleanliness and that one is available for each passenger. (3) Check flight compartment and cabin fire extinguishers for integrity of seal, security of mounting and correct charge. (4) Check overwater survival equipment (life vests, rafts, etc.), if required. (5) Check emergency lighting equipment. INSPECTION COMPLETED I certify that a Routine Inspection was performed in accordance with the Continuous Inspection Program and that the airplane is approved for return to service. CREW CHIEF________________________________________________

Page 4 Feb 1/10

05-20-01

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

CONTINUOUS INSPECTION WORKSHEET Inspection Type_________________________________Inspection Number__________________________ Model__________L Eng Time___________Cycles__________R Eng Time___________Cycles___________ A/C Time_______________Cycles________________A/C Serial_________________Date_______________

Item

Description

Mech

05-20-01

Crew Chief

Q.C.

Page 5 Feb 1/10

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

TIME LIMITS/MAINTENANCE CHECKS FIRST 200-HOUR-INTERVAL DETAILED INSPECTION DETAILED INSPECTION PROCEDURES

05-20-02 00

1. GENERAL A. Forms Required (1) First 200-Hour-Interval Detailed Inspection. (2) Model 1900/1900C Airliner Routine Inspection. (3) Continuous Inspection Worksheet. NOTE: A Routine Inspection must be conducted in conjunction with each Detailed Inspection to comply with Continuous Inspection Regulations.

B. Reference Material (1) Model 1900/1900C Airliner Maintenance Manual P/N 114-590021-7. (2) Model 1900 Airliner Series Component Maintenance Manual P/N 114-590021-11. (3) Model 1900 Airliner Series Structural Repair Manual P/N 114-590021-9. (4) Model 1900 Airliner Series Corrosion Control Manual P/N 114-590021-197. (5) Model 1900 Airliner Wiring Diagram Manual P/N 114-590032-3 (UA-1 and After); 114-590021-13 (UB-1 and After); 114-590021-61 (UC-1 and After). (6) Model 1900 Airliner Pilots Operating Handbook/Airplane Flight Manual P/N 114-590021-3. (7) Model 1900C Airliner Pilots Operating Handbook/Airplane Flight Manual P/N 114-590021-57.

C. Inspection Procedures (1) Fill out the heading on each form in its entirety. (2) The mechanic checks each item on the inspection form and initials the form in the space provided. (3) List all discrepancies found during the inspection on the Continuous Inspection Worksheet. NOTE: Check all In-Flight Worksheets turned in since the last inspection for discrepancies that have not yet been worked off. (4) Each discrepancy is to be signed off by the mechanic, crew chief and a Quality Control Inspector when the discrepancy has been corrected. (5) In the spaces provided on the Major Maintenance Worksheet, the mechanic is to list all components which are removed from the airplane for overhaul or replacement, then to add Part Number and Serial Number of the component which is installed. (6) The Quality Control Inspector will stamp off each item on the inspection form to complete the inspection.

05-20-02

Page 1 Nov 1/11

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (7) When the inspection has been completed, the mechanic, crew chief and the Quality Control Inspector must sign the INSPECTION COMPLETED block at the end of the inspection form.

2. DETAILED INSPECTION PROCEDURES Inspection Date___________________________________________ Airplane Serial___________________ Airframe______________________ Cycles____________ LH Engine Hrs_____________ Cycles____________ RH Engine Hrs______________ Cycles____________ LH Engine Power Module S/N_______________________________________________ RH Engine Power Module S/N_______________________________________________ LH Engine Gas Generator S/N_______________________________________________ RH Engine Gas Generator S/N_______________________________________________

NOTE: Corrosion detected while performing this detailed inspection may be treated in accordance with Chapter 20-09-00. To minimize the possibility of foreign object damage to engines, observe the following maintenance practices: •

Ensure all loose materials (rivets, screws, safety wire, etc.) are removed from engine cowling area after maintenance.

•

Maintain clean ramp and taxi areas.

•

Running at maximum power with the airplane stationary should be minimized and done only on a clean ramp.

•

Propeller reverse operation for backing the airplane should be avoided.

•

Avoid operation in dust and sand storms.

•

Do not operate engines in feather, except during external power starts and feather checks.

Prior to beginning this inspection, the following access panels must be removed: UA-1 and After; Wing Access - 1, 2, 3, 4, 5, 6, 8, 10, 11, 12, 13, 15, 16, 17, 18, 19, 20, 21, 22, 24, 25, 26, 27, 28, 29, 30, 35, 36, 37, 38, 43, 44, 45, 46, 47, 50, 51, 53, 54, 55, 56, 57, 58, 59, 60, 61. Floor Access - 7, 11. UB-1 and After; Wing Access - 1, 2, 3, 4, 5, 6, 8, 10, 11, 12, 13, 15, 16, 17, 18, 19, 20, 21, 22, 24, 25, 26, 29, 30, 35, 36, 37, 38, 43, 44, 45, 46, 47, 50, 51, 53, 54, 55, 56, 57, 58, 59, 60, 61. Floor Access - 8, 12. UC-1 and After; Wing Access - 1, 2, 3, 4, 5, 6, 8, 9, 10, 11, 12, 13, 15, 16, 17, 18, 19, 21, 22, 23, 24, 25, 26, 27, 28, 29. Floor Access - 8, 12. For zone and access panel locations (Ref. Chapter 06).

Page 2 Nov 1/11

05-20-02

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

A. Wings First 200-Hour-Interval Detailed Inspection Complete Steps 1 thru 19.

Mechanic LH

Inspector

RH

(1) SKINS Zone inspection areas: 611 thru 650, 511 thru 550. (a) Inspect skins for cracks, scratches, dents, paint blistering, corrosion and loose or missing fasteners. If any damage or corrosion is found, inspect the adjacent structure. (b) Inspect the entire circumference of the aft nacelle panel seals for deterioration, signs of compression and reduced resilience. Compress (squeeze) the seal and release. If the seal does not easily compress or regain its original shape quickly, replace the seal. (c) Inspect for evidence of chafing between the aft nacelle fairing and the upper wing skin (Ref. Chapter 57-90-04 of the 1900 Airliner Series Structural Repair Manual). (2) STRUCTURE Wing panel inspection areas: Inside all areas where panels are removed. (a) Check for cracks, loose rivets, corrosion and concealed damage inside all wing inspection areas where access panels have been removed. Repair any discrepancies (Ref. Model 1900 Airliner Series Structural Repair Manual, P/N 114-590021-9). (3) AILERONS AND TRIM TAB Zone inspection areas: 533, 543, 633 and 643. (a) Inspect skins for cracks, scratches, damage, loose or missing rivets, paint blistering and corrosion.

05-20-02

Page 3 Nov 1/11

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL First 200-Hour-Interval Detailed Inspection (Continued) Complete Steps 1 thru 19.

Mechanic LH

(b) Check ailerons for attachment, freedom of movement, and amount of movement (freeplay). Perform the AILERON FREEPLAY CHECK procedure (Ref. Chapter 27-10-00). Read the aileron acceptable freeplay distance from zero on the travel board scale or dial indicator. Right Aileron Freeplay:_______ (UA and UB Limits 0.12 Inch Maximum) Left Aileron Freeplay: _______(UA and UB Limits 0.12 Inch Maximum) Right Aileron Freeplay: ________ (UC Limits 0.06 Inch Maximum) Left Aileron Freeplay: ________ (UC Limits 0.06 Inch Maximum) (c) Check aileron trim tab actuator for correct direction of travel, smoothness of travel, corrosion and attachment. (d) Perform AILERON TRIM TAB FREEPLAY CHECK procedures and Table (Ref. Chapter 27-10-04). AILERON TRIM TAB FREEPLAY: ________(LIMIT: 0.053 INCH MAXIMUM) (e) Perform the AILERON TRIM TAB FUNCTIONAL CHECK procedure (Ref. Chapter 27-10-07). Rotate the cockpit pedestal aileron trim control knob counter clockwise to the full left position and verify the aileron (left side) trim tab moves in the trailing edge down direction. Check that the system moves smoothly with no unusual noise or binding. Aileron Trim Tab Travel Full Down:__________(LIMIT: 13.5° to 16.5° trim tab trailing edge down) Turn the cockpit pedestal aileron trim control knob clockwise to the full right position and verify the aileron (left side) trim tab moves in the trailing edge up direction. Check that the system moves smoothly with no unusual noise or binding. Aileron Trim Tab Travel Full Up:___________(LIMIT: 13.5° to 16.5° trim tab trailing edge up)

Page 4 Nov 1/11

05-20-02

RH

Inspector

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL First 200-Hour-Interval Detailed Inspection (Continued) Complete Steps 1 thru 19.

Mechanic LH

Inspector

RH

(f) Perform the AILERON FUNCTIONAL CHECK procedure (Ref. Chapter 27-10-03). Check that the system moves smoothly with no unusual noise or binding. Move the pilot’s control wheel counter clockwise and make sure that the left aileron moves in the trailing edge up direction. Left Aileron Travel:________ (Limit: 23° to 26° trailing edge up) Right Aileron Travel:_______ (Limit: 16° to 19° trailing edge down) (Limit: 15° to 18° for UA and UB airplanes) Move the pilot’s control wheel clockwise and make sure that the left aileron moves in the trailing edge down direction. Left Aileron Travel:________ (Limit: 16° to 19° trailing edge down) (Limit: 15° to 18° for UA and UB airplanes) Right Aileron Travel:________ (Limit: 23° to 26° trailing edge up) (4) FLIGHT CONTROL COMPONENTS, CABLES AND PULLEYS Wing panel inspection areas: 27, 28, 29, 30, 37, 38 (UA-1 and After; UB-1 and After), 19 (UC-1 and After). (a) Inspect control system components (pushrods, turnbuckles, end fittings, castings, etc.) for bulges, splits, bends, cracks and corrosion. Replace any damaged component. (b) Check control cables, pulleys and associated equipment for condition, attachment, alignment, clearance, corrosion, cleanliness and correct direction of travel. Check areas around cables for evidence of chafing. Replace any damaged component. Replace cables that have broken strands or evidence of corrosion (Ref. Chapter 20-00-02). (c) Inspect pulleys under evaporator ducts. (d) Perform the AILERON WING CABLE TENSION CHECK procedures (Ref. Chapter 27-10-03). Record aileron cable and aileron trim cable tensions: Temperature:________ °F 3/16 inch Aileron Cable Tension LH: _______ RH: _______ 1/16 inch Aileron Trim Tab Cable Tension: ________

05-20-02

Page 5 Nov 1/11

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL First 200-Hour-Interval Detailed Inspection (Continued) Complete Steps 1 thru 19.

Mechanic LH

(e) Inspect all flight control rigging pin placards for legibility and security of attachment (if installed) (Ref. Chapter 11-30-00). (5) FLAPS AND ACTUATORS Zones inspection areas: 513, 533, 613 and 633. Wing panel inspection areas: 35, 36 (UA-1 and After; UB-1 and After), 17 (UC-1 and After). (a) Inspect flap drive cables and actuators for wear, corrosion and attachment. (b) Inspect flap skins and structure for cracks, scratches, dents, paint blistering, corrosion and loose or missing rivets. In addition, inspect flap roller brackets, roller bearings and attachment hardware for damage, corrosion and attachment. (c) Perform the LEFT AND RIGHT FLAP SAFETY SWITCH TEST procedure (Ref. Chapter 27-50-06). (6) FUEL TANKS, VENTS, VALVES AND PUMPS Zone inspection areas: 550, 650, 730 and 740. Wing panel inspection areas: 17, 18, 25, 26, 50, 51 (UA-1 and After; UB-1 and After), 9, 21 (UC-1 and After). (a) Inspect fuel tanks for corrosion, leaks and plugged or obstructed vents. (b) Check under wing vent and wing tip fuel vents for obstructions and erosion. (c) Check the heated vent for operation (warm to the touch). (d) Check pumps, drain valves, firewall shutoff valves and fuel low pressure switch for evidence of leaks and security of attachment. (7) FUEL FILLER CAP Wing panel inspection areas: 53 (UA-1 and After; UB-1 and After), 3 (UC-1 and After). (a) Check for damage, scratches, paint blistering, corrosion and leaks. (b) Check locking mechanism for ease of operation. Check integrity of cap lanyard. (8) FUEL VALVE, ANTISIPHON Wing panel inspection areas: 53 (UA-1 and After; UB-1 and After), 3 (UC-1 and After). (a) Check for condition.

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(9) INTEGRAL FUEL TANK Zone Inspection areas: 542 and 642. (a) Check the integral tank access plates for leaks and condition. (10) FUEL PROBES Wing panel inspection areas: 2, 3, 4, 5, 6, 19, 20, 21, and 22 (UA-1 and After; UB-1 and After), 2 and 5 (UC-1 and After). (a) Check for leaks at points of attachment. (b) Visually inspect for damage, security and corrosion. (11) ELECTRICAL WIRING AND EQUIPMENT Wing panel inspection areas: 1, 4, 13 (UA-1 and After; UB-1 and After); 21 and 22 (UC-1 and After). (a) Check for security, chafing, damage and attachment. (12) BLEED AIR PRESSURE REGULATOR/SHUTOFF VALVE Zone inspection area: 511 and 611 (a) Inspect the filter (if installed) for condition and cleanliness. (Ref. Model 1900 Airliner Series Component Maintenance Manual P/N 114-590021-11). (b) Inspect and clean the filter (if installed) in the servo air line to the shutoff valve. (13) LEADING EDGE AND NACELLE PLUMBING AND WIRING Zone inspection areas: 511, 521, 522, 531, 541, 611, 621, 622, 631 and 641. Wing panel inspection areas: 54, 55, 56, 57, 58, 59, 60, 61 (UA-1 and After; UB-1 and After). 4, 23, 24, 25, 26, 27, 28 and 29 (UC-1 and After). (a) Visually check plumbing for damage, security, leaks and corrosion. (b) Check the wiring for chafing and security of attachment. (c) Clean the power distribution bay and inspect for corrosion (Ref. Chapter 24-50-00 of the Model 1900 Airliner Series Corrosion Control Manual). (d) Inspect the power distribution panels and electrical connectors for corrosion and security. (e) Clean the inverter bay and inspect for corrosion (Ref. Chapter 24-50-00 of the Model 1900 Airliner Series Corrosion Control Manual). (f) Inspect the inverters, inverter cooling fans, ac circuit breakers and electrical connectors for corrosion and security.

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(14) PLUMBING Zone inspection areas: 531, 532, 631 and 632. Wing panel inspection areas: 4, 17, 18 (UA-1 and After; UB-1 and After). 8, 9, 11, 12, 15, 18, 21, 23, 25 and 29 (UC-1 and After). (a) Visually check for leaks, chafing, corrosion or damage and proper attachment. (15) DEICER BOOTS Zone inspection areas: 511, 531, 541, 611, 631 and 641. (a) Visually check deicer boots for cracks, gaps, tears, damage and attachment. Inspect deicer boot edge sealant for looseness, gaps and damage. Reapply as necessary. (16) POWER PACK HYDRAULIC SYSTEM BLEED AIR FILTER Zone inspection area: 511. Wing panel inspection area: 10. (a) Clean filter (Ref. Chapter 32-30-06). (17) AIR CYCLE MACHINE Zone inspection area: 511. Wing panel inspection area: 15 and 57 (UA-1 and After; UB-1 and After). 15 and 18 (UC-1 and After). (a) Change air cycle machine oil. (Ref. Hamilton Standard B-1900 Refrigeration Package Maintenance Manual and Parts List in the Model 1900 Airliner Series Component Maintenance Manual, Chapter 21). (b) Visually and aurally check the bypass valves for operation of the actuator. (c) Check for security of mounting and obstructions of ambient air flow across the heat exchanger (Ref. Chapter 21-51-00). (18) AIR CYCLE MACHINE FOG NOZZLE AND FILTER Zone inspection area: 511. Wing panel inspection area: 11 and 15. (a) Clean the air cycle machine fog nozzle and filter (if installed) (Ref. Chapter 21-51-02). (19) LANDING GEAR HYDRAULIC POWER PACK Zone inspection area: 511. Wing panel inspection area: 10 and 12. (a) Visually inspect area around the Hydraulic Power Pack for leaks, corrosion, damage and security of attachment.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

B. FWD Right-hand Center Section First 200-Hour-Interval Detailed Inspection Complete Steps 1 thru 3.

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(1) STRUCTURE Wing panel inspection areas: 54 and 55 (UA-1 and After; UB-1 and After), 23 and 24 (UC-1 and After). (a) Check structure for cracks, scratches, corrosion, loose rivets and damage. (b) Repair any damage. (Ref. Model 1900 Airliner Series Structural Repair Manual P/N 114-590021-9). (2) PLUMBING AND WIRING (a) Inspect plumbing and wiring for chafing, leaks, corrosion and security. (3) AIR-CONDITIONING CONDENSER AND BLOWER Zone inspection area: 611. Wing Panel inspection area: 16. (a) Check the condenser, blower and associated plumbing for leaks, corrosion, damage and security of attachment. (b) Inspect the inlet guard for security of attachment and for broken strands. (c) Inspect the impeller for security to the shaft and ease of rotation. (d) Inspect the standoffs for security and tightness. (e) Inspect the guide vanes in the blower housing assembly for cracking, security of attachment and corrosion.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

C. Power Plant First 200-Hour-Interval Detailed Inspection Complete Steps 1 and 2.

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(1) IGNITER PLUGS Zone inspection areas: 400. (a) Inspect the igniter plugs for condition and erosion (Ref. Chapter 74-00-00 and the Pratt and Whitney PT6A-65B Engine Maintenance Manual). (2) OIL FILTER Zone inspection areas: 400. (a) Inspect the oil filter for contamination (Ref. Pratt and Whitney Engine Maintenance Manual PT6A-65B, Chapter 72-00-00, Table 601 Periodic Inspection).

D. General Service Items First 200-Hour-Interval Detailed Inspection Complete Steps 1 thru 9. (1) AIRPLANE LUBRICATION (a) Lubricate as necessary in accordance with the LUBRICATION SCHEDULE (Ref. Chapter 12-20-00). (2) INSTRUMENT AIR FILTER Zone inspection area: 212. (a) Replace the filter (Ref. Chapter 37-00-00, INSTRUMENT AIR FILTER SERVICING). (3) EVAPORATOR FILTERS Zone inspection areas: 153 and 173. Floor panel areas: 7 and 11 (UA-1 and After); 8 and 12 (UB-1 and After; UC-1 and After). (a) Clean or replace the evaporator filters (Ref. Chapter 21-52-01). (4) PITOT AND STATIC SYSTEM Zone inspection area: 221 (UC-1 and After), 246, 247 and 248. (a) Drain the system (Ref. Flight Manual Supplements 114-590021-41 and 114-590021-87 as applicable). Close the drains when completed.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL First 200-Hour-Interval Detailed Inspection (Continued) Complete Steps 1 thru 9.

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(5) ELECTRONIC FLIGHT INSTRUMENT SYSTEM (EFIS) Zone inspection area: 248. (a) Verify operation of Electronic Attitude Director Indicator (EADI) and Electronic Horizontal Situation Indicator (EHSI) tube fans by listening for fan operation. EADI and EHSI tubes must be on (Ref. the appropriate Pilot's Operating Handbook/ Airplane Flight Manual). (6) VACUUM REGULATOR VALVE FILTER Zone inspection area: 812. (a) Clean or replace the filter (Ref. Chapter 37-00-00). (7) PLACARDS (a) Verify all placards are in place and legible (Ref. Pilots Operating Handbook, Airplane Flight Manual, Chapter 11-20-00 (UA-1 and After, UB-1 and After) and Chapter 11-21-00 (UC-1 and After)). (8) ACCESS PANELS (a) Check all panels removed for this inspection for fit, attachment, scratches, paint blistering, corrosion and visual damage. (9) WINDSHIELDS (a) Inspect windshield weather seal for debonding, cracks and wear (Ref. Chapter 56-10-00).

E. Operational Inspection First 200-Hour-Interval Detailed Inspection Complete Steps 1 thru 6. (1) ENVIRONMENTAL TEST (a) Perform the BLEED AIR TEMPERATURE OPERATIONAL TEST procedure (Ref. Chapter 21-10-00 or Chapter 21-11-00). (b) Perform the BLEED AIR PRESSURE OPERATIONAL TEST procedure (Ref. Chapter 21-10-00 or Chapter 21-11-00). (2) PROPELLER DEICER (a) Perform the PROPELLER DEICER SYSTEM INSPECTION procedure (Ref. Chapter 30-60-00).

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(3) GROUND INSPECTION RUN Depending on the maintenance performed and components replaced, a GROUND PERFORMANCE CHECK PROCEDURE (Ref. Chapter 76-00-00) may be required in lieu of this Inspection Run. Refer to the applicable maintenance procedures. (a) Start engines and allow the oil temperature to increase into the operating range. (b) Run engines at a minimum of 80% N1 long enough for engine indicators to stabilize. (c) Shut down the engines and inspect for attachment and security of all components and for oil and fuel leaks. (4) INVERTER POWER (a) Perform the INVERTER POWER SELECT RELAY CHECK procedure (Ref. Chapter 24-20-00). (b) Perform the INVERTER BLOWER FAN OPERATIONAL CHECK procedure (Ref. Chapter 24-20-00). (5) STATIC DISCHARGER (a) Perform the STATIC DISCHARGER INSPECTION procedure (Ref. Chapter 23-60-00) on all wing (aileron and wing tip) mounted static dischargers. (6) EXTERNAL POWER Zone inspection area: 522. (a) Check the external power relay for operation (rotate the voltmeter select switch to the EXT PWR position and check for external power voltage).

INSPECTION COMPLETED I certify that a Detailed Inspection was performed in accordance with the Continuous Inspection Program and that the airplane is approved for return to service. CREW CHIEF________________________________________________

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

CONTINUOUS INSPECTION WORKSHEET Inspection Type_________________________________Inspection Number__________________________ Model__________L Eng Time___________Cycles__________R Eng Time___________Cycles___________ A/C Time_______________Cycles________________A/C Serial_________________Date_______________

Item

Description

Mech

05-20-02

Crew Chief

Q.C.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Inflight Worksheet (Page 1 of 2) S/N________________N Number__________________Date___________________NO._________________ Flt. No.

Pilot

Copilot

TO

Time

LND

Time

No. LND

Flt. Time

Today’s Total Previous Total Total Fuel Type

Engine Oil Gallon

Left

Cruise Condition

Right

Data O.A.T. PA IAS ITT Torque Prop. RPM N1 RPM Fuel Flow Fuel Pressure Oil Pressure Oil Temp.

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Right

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Inflight Worksheet (Page 2 of 2) Discrepancy Worksheet S/N________________N Number__________________Date___________________NO._________________

A/C Hrs

Date

Discrepancy

Corrective Action

Date

Mechanic Inspector

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Major Maintenance Worksheet Component

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Date

Reason for Replacement

05-20-02

Replacement Part Number Serial Number

Next Overhaul A/C Hours or Cycles Date

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

TIME LIMITS/MAINTENANCE CHECKS SECOND 200-HOUR-INTERVAL DETAILED INSPECTION DETAILED INSPECTION PROCEDURES

05-20-03 00

1. GENERAL A. Forms Required (1) Second 200-Hour-Interval Detailed Inspection. (2) Model 1900/1900C Airliner Routine Inspection. (3) Continuous Inspection Worksheet. NOTE: A Routine Inspection must be conducted in conjunction with each Detailed Inspection to comply with Continuous Inspection Regulations.

B. Reference Material (1) Model 1900/1900C Airliner Maintenance Manual, P/N 114-590021-7. (2) Model 1900 Airliner Series Component Maintenance Manual, P/N 114-590021-11. (3) Model 1900/1900C Airliner Structural Inspection Manual, P/N 98-30973. (4) Model 1900 Airliner Series Structural Repair Manual, P/N 114-590021-9. (5) Model 1900 Airliner Wiring Diagram Manual P/N 114-590032-3 (UA-1 and After), 114-590021-13 (UB-1 and After), 114-590021-61 (UC-1 and After). (6) Model 1900 Airliner Pilot’s Operating Handbook/Airplane Flight Manual, P/N 114-590021-3. (7) Model 1900C Airliner Pilot’s Operating Handbook/Airplane Flight Manual, P/N 114-590021-57.

C. Inspection Procedures (1) Fill out the heading on each form in its entirety. (2) The mechanic checks each item on the inspection form and initials the form in the space provided. (3) List all discrepancies found during the inspection on the Continuous Inspection Worksheet. NOTE: Check all In-Flight Worksheets turned in since the last inspection for discrepancies that have not yet been worked off. (4) Each discrepancy is to be signed off by the mechanic, crew chief and a Quality Control Inspector when the discrepancy has been corrected. (5) In the spaces provided on the Major Maintenance Worksheet, the mechanic is to list all components which are removed from the airplane for overhaul or replacement, then to add the Part Number and Serial Number of the component which is installed. (6) The Quality Control Inspector will stamp off each item on the inspection form to complete the inspection.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (7) When the detailed inspection has been completed, the mechanic, crew chief and the Quality Control Inspector must sign the INSPECTION COMPLETED block at the end of the inspection form.

2. DETAILED INSPECTION PROCEDURES Inspection Date___________________________________________ Airplane Serial___________________ Airframe______________________ Cycles____________ LH Engine Hrs_____________ Cycles____________ RH Engine Hrs______________ Cycles____________ LH Engine Power Module S/N_______________________________________________ RH Engine Power Module S/N_______________________________________________ LH Engine Gas Generator S/N_______________________________________________ RH Engine Gas Generator S/N_______________________________________________

NOTE: Corrosion detected while performing this detailed inspection may be treated in accordance with Chapter 20-09-00. To minimize the possibility of foreign object damage to engines, observe the following maintenance practices: •

Ensure all loose materials (rivets, screws, safety wire, etc.) are removed from engine cowling area after maintenance.

•

Maintain clean ramp and taxi areas.

•

Running at maximum power with the airplane stationary should be minimized and done only on a clean ramp.

•

Propeller reverse operation for backing the airplane should be avoided.

•

Avoid operation in dust and sand storms.

•

Do not operate engines in feather, except during external power starts and feather checks.

Prior to beginning this inspection, the engine cowling and the following access panels must be removed: UA-1 and After; Wing Access - 12 and 15 Floor Access - 7 and 11 UB-1 and After; Wing Access - 12 and 15 Floor Access - 8 and 12 UC-1 and After; Wing Access - 12 and 13 Floor Access - 8 and 12 For zone and access panel location (Ref. Chapter 06).

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A. Power Plant Second 200-Hour-Interval Detailed Inspection Complete Steps 1 thru 25.

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(1) COWLING Zone inspection areas: 400. (a) Check adjustment of latches. Inspect for cracks, dents, paint blistering, corrosion and loose or missing fasteners. (b) Inspect aft cowling door latches for excessive wear, distortion, or deterioration of latch pawl tips. Replace if required. (2) ENGINE OIL FILTER Zone inspection areas: 400. (a) Inspect the oil filter for contamination (Ref. Pratt and Whitney PT6A-65B Engine Maintenance Manual P/N 3032842, Chapter 72-00-00, Table 601 Periodic Inspection). (3) MAGNETIC CHIP DETECTOR CLEANING Zone inspection areas: 400. (a) Clean the magnetic chip detector (Ref. Pratt and Whitney PT6A-65B Engine Maintenance Manual P/N 3032842). (4) DRAIN PLUGS Zone inspection areas: 400. (a) Check all drain plugs for security. (5) IGNITER PLUGS Zone inspection areas: 400. (a) Inspect the igniter plugs for condition and erosion (Ref. Chapter 74-00-00 and the Pratt and Whitney PT6A-65B Engine Maintenance Manual P/N 3032842). (6) P3 FILTER Zone inspection areas: 400. (a) Clean or Replace (Ref. Pratt and Whitney PT6A-65B Engine Maintenance Manual P/N 3032842). (7) FUEL NOZZLES Zone inspection areas: 400. (a) Inspect and clean the fuel nozzles using either the “in situ” method or by bench check (Ref. Pratt and Whitney PT6A-65B Engine Maintenance Manual P/N 3032842). (b) Borescope inspect the engine hot section at this time (Ref. Pratt and Whitney PT6A-65B Engine Maintenance Manual P/N 3032842.

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(8) EXHAUST SYSTEM Zone inspection areas: 400. (a) Inspect the exhaust system and visible portions of the power turbine for burning, distortion, dents, corrosion and cracks (Ref. Pratt and Whitney PT6A-65B Engine Maintenance Manual for corrective action). (b) Inspect exhaust stacks for cracking. (c) Inspect engine inlet heat tubes and ducts for cracks. (9)

INDUCTION SYSTEM Zone inspection areas: 400. (a) Inspect the air intake duct and engine inlet screen for obstruction, cracks, scratches, paint blistering, corrosion and security. (b) Remove the air inlet screen and inspect the compressor inlet area, struts and first stage blades for dirt deposits, corrosion, erosion and foreign object damage.

(10) OIL COOLER Zone inspection areas: 400. (a) Inspect the cooler and plumbing for leakage, corrosion and attachment. (11) FIRE DETECTORS Zone inspection areas: 400. (a) Check forward and aft fire detector loops for condition and security. Correct any evidence of chaffing or interference with other components or structure by repositioning the fire detector loop clamps as required. Check fire detector loop clamps and connectors for condition and security (Ref. Chapter 26-10-00). (12) ENGINE FIRE EXTINGUISHERS Zone inspection areas: 730 and 740. (a) Check pressure of supply cylinders using the pressure gage mounted on the cylinder (Ref. Chapter 26-20-00, EXTINGUISHER CARTRIDGE AND SUPPLY CYLINDER SERVICE LIFE). (b) Visually check plumbing for leakage, corrosion and security of attachment. (c) Check for presence of activation voltage to the squibs (Ref. Chapter 26-20-00, EXTINGUISHER ACTIVATION CHECK).

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(13) ENGINE ACCESSORIES Zone inspection areas: 400. (a) Inspect all accessories, plumbing and associated equipment for damage, attachment, corrosion and leakage. (14) ENGINE VIBRATION ISOLATOR MOUNTS Zone inspection area: 400. CAUTION If any isolator mounts have dislodged from their bracket positioning pins or have a gap or change in relative position, the airplane may have experienced a hard landing or encountered severe or extreme turbulent air. If so, perform additional inspections in Chapter 05-50-00. All mounts on an engine must be of the same manufacturer and carry the same part numbers. (a) Inspect for damage, scratches, corrosion and attachment (Ref. Chapter 71 of the Model 1900 Airliner Series Component Maintenance Manual, P/N 114-590021-11). Replace or repair as required. The mount may be repaired by replacement of the rubber cushion. (15) ENGINE MOUNT TRUSS Zone inspection area: 400. (a) Visually inspect the engine mount truss assembly for cracks, scratches, corrosion, condition of paint, chafing and dents with a 10x magnifying glass (Ref. Model 1900/1900C Airliner Series Structural Inspection Manual P/N 98-30973). (b) Perform the ENGINE TRUSS BOLT TORQUE CHECK procedure (Ref. Chapter 71-20-00). (16) ELECTRICAL WIRING AND EQUIPMENT Zone inspection area: 400. (a) Inspect wiring and associated equipment and accessories for damage, chafing and attachment. (17) INERTIAL ANTI-ICER Zone inspection areas: 400. (a) Check the inertial vane and bypass door for freedom of movement and correct travel with the electrical actuator. (b) Inspect bypass door shock links for condition and security. (18)

PROPELLERS (a) Inspect propellers for evidence of leakage, damage, erosion and attachment.

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Mechanic LH

(b) Inspect the mechanical feedback ring, stop rods and springs for damage and corrosion. (c) Inspect the carbon block pin for freedom of movement. (d) Check the clearance between the beta ring and the metal retaining clip of the carbon block assembly. If at any point this clearance is 0.005 inch or less replace the carbon block. (e) Inspect the reversing linkage for setting, operation, evidence of binding and security of attachment. (f) Check operation of all pedestal controls and switches. (19) AUTOFEATHER/AUTO-IGNITION PRESSURE SWITCHES Zone Inspection Area: 400. (a) Check for Operation, security of attachment and correct electrical connections. (20) STARTER-GENERATOR (a) Inspect brushes for indication of excessive wear or damage (determine wear by observing diagonal groove on brush). (Ref. Chapter 24-30-01 and to the Starter-Generator Manufacturer’s Maintenance Manual in Chapter 24 of the Model 1900 Airliner Series Component Maintenance Manual). (b) Inspect inlet duct and cooling cap for cracks, corrosion and obstruction. (21) IGNITION EXCITER (a) Check exciter and electrical harness for damage, chafing and security of attachment. (22) FUEL PURGE SYSTEM (a) Check plumbing and tank for leaks and security of attachment. (b) Perform the FUEL PURGE SYSTEM AIR FILTER CLEANING procedure (Ref. Chapter 71-70-00). (c) Perform the FUEL PURGE SYSTEM CHECK VALVE INSPECTION, CLEANING AND LEAKAGE TEST procedure (Ref. Chapter 71-70-00). (d) Perform the FUEL FLOW DIVIDER/PURGE VALVE INSPECTION procedure (Ref. Chapter 71-70-00). (23) COMPRESSOR DRIVE QUILL SHAFT Zone inspection area: 420. (a) Check for wear and damage.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Second 200-Hour-Interval Detailed Inspection (Continued) Complete Steps 1 thru 25.

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(b) Lubricate the spline on the pulley end of the shaft (Ref. Chapter 21-52-02). (24) COMPRESSOR DRIVE BELTS Zone inspection area: 420. (a) Check for cracks, shredding and wear. Check the adjustment (Ref. Chapter 21-52-02). (25) ENGINE FUEL PUMP, FILTERS AND SCREENS Zone inspection areas: 400. (a) Inspect the filters and screens for contamination (Ref. Pratt and Whitney PT6A-65B Engine Maintenance Manual P/N 3032842).

B. Environmental System Second 200-Hour-Interval Detailed Inspection Complete Step 1.

Mechanic

Inspector

(1) TEMPERATURE CONTROLLER (QUARTZ ROD) AND FILTER, PRECOOLER VALVES AND ASSOCIATED PLUMBING Zone inspection areas: 521 and 621. Panel Inspection areas: Lower aft cowling. (a) Inspect and clean the filters in each end of the sense line to the temperature controller. (Ref. Chapter 21-10-00 for UA-1 and After; UB-1 and After; UC-1 thru UC-45, except UC-39. For UC-39 and UC-46 and After Ref. Chapter 21-11-00). (b) Perform the PRECOOLER VALVES FUNCTIONAL CHECK procedure (Ref. Chapter 21-10-00 for UA-1 and After; UB-1 and After; UC-1 thru UC-45, except UC-39. For UC-39 and UC-46 and After Ref. Chapter 21-11-00).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

C. General Service Items Second 200-Hour-Interval Detailed Inspection Complete Steps 1 thru 12. (1) AIRPLANE LUBRICATION (a) Lubricate as necessary in accordance with the LUBRICATION SCHEDULE (Ref. Chapter 12-20-00). (2) ALL HOSE ASSEMBLIES Zone inspection areas: 410, 420, 730 and 740. (a) Visually check all hose assemblies and tubes in the engine compartment and wheel wells for cracks, chafing, damage and leaks. (3) ANTISKID BRAKES (IF INSTALLED) Zone inspection area: 730 and 740. (a) Check operation, charge accumulator as required and replace filter (Ref. Chapter 32-40-00). (b) Remove antiskid hydraulic line/antiskid accumulator cover and inspect for damage, cracks, leaks, deterioration and security. (4) POWER STEERING (IF INSTALLED) Zone inspection area: 710. (a) Replace the filter (Ref. Chapter 32-52-00). (5) EVAPORATOR FILTERS Zone inspection areas: 153 and 173. Floor panel inspection areas: 7 and 11 (UA-1 and After); 8 and 12 (UB-1 and After; UC-1 and After). (a) Replace the evaporator filters (Ref. Chapter 21-52-01). (6) PITOT AND STATIC SYSTEM Zone inspection areas: 221 (UC-1 and After), 246, 247 and 248. (a) Drain system (Ref. Flight Manual Supplements 114-590021-41 and 114-590021-87 as applicable). Close the drains when completed. (7) ELECTRONIC FLIGHT INSTRUMENT SYSTEM (EFIS) (IF INSTALLED) Zone inspection area: 248. (a) Verify operation of Electronic Attitude Director Indicator (EADI) and Electronic Horizontal Situation Indicator (EHSI) tube fans by listening for fan operation. EADI and EHSI tubes must be on (Ref. the appropriate Pilot's Operating Handbook/ Airplane Flight Manual).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Second 200-Hour-Interval Detailed Inspection (Continued) Complete Steps 1 thru 12.

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Inspector

(8) VACUUM REGULATOR VALVE FILTER Zone inspection area: 812. (a) Clean or replace filter (Ref. Chapter 37-00-00). (9) PRESSURIZATION SYSTEM DRAIN VALVE Panel inspection area: 311 and 312. (a) On airplanes with the valve installed, open the drain valve to remove condensation in pressure lines. Close drains upon completion. (10) PLACARDS (a) Verify all placards are in place and legible (Ref. Pilots Operating Handbook, Airplane Flight Manual, Chapter 11-20-00 (UA-1 and After, UB-1 and After) and Chapter 11-21-00 (UC-1 and After)). (11) FUEL FILTERS AND SCREENS Zone inspection areas: 400, 410, 420, 511, 611, 730 and 740. (a) Inspect the filters and screens for contamination. Clean filters and screens (Ref. Chapter 28-21-02). (12) WINDSHIELDS (a) Inspect windshield weather seal for debonding, cracks and wear.

D. Operational Inspection Second 200-Hour-Interval Detailed Inspection Complete Steps 1 thru 18.

Mechanic

Inspector

NOTE The following Operational Inspection procedures are to be applied during start and run of the engine. CAUTION Do not place the power levers into reverse unless engines are running. (1) ENGINE CONTROLS (a) Check for freedom of movement, full travel and friction lock. (2) STARTER-GENERATOR (a) Check for output of 28.25 ± 0.25 vdc (Ref. Chapter 24-30-01). (3) OIL (a) Check that the pressure and temperature are within limits (Ref. the applicable Pilot's Operating Handbook/Airplane Flight Manual).

05-20-03

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Second 200-Hour-Interval Detailed Inspection (Continued) Complete Steps 1 thru 18. (4) PROPELLER GOVERNOR (a) Check operation and manual feathering (Ref. the applicable Pilot's Operating Handbook/Airplane Flight Manual). (5) PROPELLER SYNCHROPHASER (a) Check operation (Ref. Chapter 61-22-00). (6) PROPELLER DEICER (a) Perform the PROPELLER DEICER SYSTEM INSPECTION procedure (Ref. Chapter 30-60-00). (7) AUTOFEATHERING SYSTEM (a) Perform autofeathering operation (Ref. Chapter 61-21-00). (8) GROUND PERFORMANCE CHECK (a) Perform the GROUND PERFORMANCE CHECK PROCEDURE procedure (Ref. Chapter 76-00-00). Ensure any airframe components that extracts power from the engine (such as the generator, ice vane, bleed air, air conditioner, etc.) are turned off during this procedure. If only one engine performance parameter is found to be outside expected limits, confirm the accuracy of the appropriate indicating system before making any engine adjustments. (b) After engine shutdown, inspect for attachment and security of all components and for oil and fuel leaks. Check for clean shutdown at IDLE CUTOFF. (9) FUEL BOOST PUMPS (a) Check the operation of the electric pumps (Ref. the Applicable Pilot's Operating Handbook/Airplane Flight Manual). (10) FUEL CROSS-TRANSFER VALVES (a) Check the operation (Ref. the applicable Pilot's Operating Handbook/Airplane Flight Manual). (11) FIREWALL FUEL SHUTOFF VALVES (a) Perform the FIREWALL FUEL SHUTOFF VALVE FUNCTIONAL CHECK procedure (UA-1 and After; UB-1 and After (Ref. Chapter 28-20-05)) or (UC-1 and After (Ref. Chapter 28-21-05)). (12) VACUUM SYSTEMS (a) Perform vacuum regulator valve adjustment (Ref. Chapter 37-00-00).

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Mechanic

Inspector

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Second 200-Hour-Interval Detailed Inspection (Continued) Complete Steps 1 thru 18.

Mechanic

Inspector

(13) ENVIRONMENTAL TEST (a) Perform the BLEED AIR TEMPERATURE OPERATIONAL CHECK procedure (Ref. Chapter 21-10-00 or Chapter 21-11-00). (b) Perform the BLEED AIR PRESSURE OPERATIONAL CHECK procedure (Ref. Chapter 21-10-00 or Chapter 21-11-00). (14) SURFACE DEICERS (a) Check for inflation and cycling. (b) Perform the SURFACE DEICER OPERATIONAL CHECK procedure (Ref. Chapter 30-10-00). (15) PRESSURIZATION SYSTEM (a) Check operation (Ref. Chapter 21-30-00). (16) ENVIRONMENTAL VAPOR CYCLE SYSTEM AND AIR CYCLE MACHINE (a) Check operation when the switch is in the AUTO or MANUAL position. Ambient temperature must be above 50°F. (b) Check operation of all outlets and ease of operation of all controls (Ref. Chapters 21-50-00 and 21-52-00). (17) INVERTER SYSTEM (a) Perform the INVERTER POWER SELECT RELAY CHECK procedure (Ref. Chapter 24-20-00). (b) Perform the INVERTER BLOWER FAN OPERATIONAL CHECK procedure (Ref. Chapter 24-20-00). (18) EXTERNAL POWER Zone inspection area: 522. (a) Check the external power relay for operation (rotate the voltmeter select switch to the EXT PWR position and check for external power voltage).

INSPECTION COMPLETED I certify that a Detailed Inspection was performed in accordance with the Continuous Inspection Program and that the airplane is approved for return to service. CREW CHIEF________________________________________________

05-20-03

Page 11 Nov 1/11

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

CONTINUOUS INSPECTION WORKSHEET Inspection Type_________________________________Inspection Number__________________________ Model__________L Eng Time___________Cycles__________R Eng Time___________Cycles___________ A/C Time_______________Cycles________________A/C Serial_________________Date_______________

Item

Page 12 Nov 1/11

Description

05-20-03

Mech

Crew Chief

Q.C.

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Inflight Worksheet (Page 1 of 2) S/N________________N Number__________________Date___________________NO._________________ Flt. No.

Pilot

Copilot

TO

Time

LND

Time

No. LND

Flt. Time

Today’s Total Previous Total Total Fuel Type

Engine Oil Gallon

Left

Cruise Condition

Right

Data

Left

Right

O.A.T. PA IAS ITT Torque Prop. RPM N1 RPM Fuel Flow Fuel Pressure Oil Pressure Oil Temp.

05-20-03

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Inflight Worksheet (Page 2 of 2) Discrepancy Worksheet S/N________________N Number__________________Date___________________NO._________________

A/C Hrs

Page 14 Nov 1/11

Date

Discrepancy

05-20-03

Corrective Action

Date

Mechanic Inspector

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Major Maintenance Worksheet Component

Date

Reason for Replacement

Replacement Part Number Serial Number

05-20-03

Next Overhaul A/C Hours or Cycles Date

Page 15 Nov 1/11

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

TIME LIMITS/MAINTENANCE CHECKS THIRD 200-HOUR-INTERVAL DETAILED INSPECTION DETAILED INSPECTION PROCEDURES

05-20-04 00

1. GENERAL A. Forms Required (1) Third 200-Hour Interval Detailed Inspection. (2) Model 1900/1900C Airliner Routine Inspection. (3) Continuous Inspection Worksheet. NOTE: A Routine Inspection must be conducted in conjunction with each Detailed Inspection to comply with Continuous Inspection Regulations.

B. Reference Material (1) Model 1900/1900C Airliner Maintenance Manual P/N 114-590021-7. (2) Model 1900 Airliner Series Component Maintenance Manual P/N 114-590021-11. (3) Model 1900 Airliner Series Structural Repair Manual P/N 114-590021-9. (4) Model 1900 Airliner Wiring Diagram Manual P/N 114-590032-3 (UA-1 and After), 114-590021-13 (UB-1 and After), 114-590021-61 (UC-1 and After). (5) Model 1900 Airliner Pilot’s Operating Handbook/Airplane Flight Manual, P/N 114-590021-3. (6) Model 1900C Airliner Pilot’s Operating Handbook/Airplane Flight Manual, P/N 114-590021-57.

C. Inspection Procedures (1) Fill out the heading on each form in its entirety. (2) The mechanic checks each item on the inspection form and initials the form in the space provided. (3) List all discrepancies found during the inspection on the Continuous Inspection Worksheet. NOTE: Check all In-Flight Worksheets turned in since the last inspection for discrepancies that have not yet been worked off. (4) Each discrepancy is to be signed off by the mechanic, crew chief and a Quality Control Inspector when the discrepancy has been corrected. (5) In the spaces provided on the Major Maintenance Worksheet, the mechanic is to list all components which are removed from the airplane for overhaul or replacement, then add the Part Number and Serial Number of the component which is installed. (6) The Quality Control Inspector will stamp off each item on the inspection form to complete the inspection.

05-20-04

Page 1 May 1/12

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (7) When the inspection has been completed, the crew chief will sign the “INSPECTION COMPLETED” block at the end of the inspection sheet.

2. DETAILED INSPECTION PROCEDURES Inspection Date___________________________________________ Airplane Serial___________________ Airframe______________________ Cycles____________ LH Engine Hrs_____________ Cycles____________ RH Engine Hrs______________ Cycles____________ LH Engine Power Module S/N_______________________________________________ RH Engine Power Module S/N_______________________________________________ LH Engine Gas Generator S/N_______________________________________________ RH Engine Gas Generator S/N_______________________________________________

NOTE: Corrosion detected while performing this detailed inspection may be treated in accordance with Chapter 20-09-00. To minimize the possibility of foreign object damage to engines, observe the following maintenance practices: •

Ensure all loose materials (rivets, screws, safety wire, etc.) are removed from engine cowling area after maintenance.

•

Maintain clean ramp and taxi areas.

•

Running at maximum power with the airplane stationary should be minimized and done only on a clean ramp.

•

Propeller reverse operation for backing the airplane should be avoided.

•

Avoid operation in dust and sand storms.

•

Do not operate engines in feather, except during external power starts and feather checks.

Prior to beginning this inspection, the following access panels should be removed: UA-1 and After; Wing Access - 12, 13, 15 and 57. Floor Access - 1, 2, 3, 4, 5, 6, 7, 8, 9, 10, 11, 12, 13, 14, 15, 16, 17 and 18. UB-1 and After; Wing Access - 12, 13, 15 and 57. Floor Access - 1, 2, 3, 4, 5, 6, 7, 8, 9, 10, 11, 12, 13, 14, 15, 16, 17 and 18. UC-1 and After; Wing Access - 12, 13, 15 and 18. Floor Access - 1, 2, 3, 4, 5, 6, 7, 8, 9, 10, 11, 12, 13, 14, 15, 16, 17 and 18. For zone and access panel locations (Ref. Chapter 06).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

A. Flight Compartment Third 200-Hour-Interval Detailed Inspection Complete Steps 1 thru 15.

Mechanic

Inspector

(1) FLIGHT CONTROL COMPONENTS, CABLES AND PULLEYS Zone inspection areas: 121, 122. (a) Inspect the control system components (pushrods, turnbuckles, end fittings, castings, etc.) for bulges, splits, bends, cracks and corrosion. (b) Inspect the control cables, pulleys and associated equipment for cracks, breaks, wear, attachment, alignment, clearance, corrosion and correct direction of travel. (c) Inspect the control cables for broken strands or evidence of corrosion (Ref. Chapter 20-00-02). (d) Check the aileron autopilot servo bridle cable tension, if installed (Ref. Chapter 22-10-00). Temperature:___________°F. Aileron Servo Cable Tension:_____________. (e) Perform the AILERON FUSELAGE CABLE TENSION CHECK procedure (Ref. Chapter 27-10-03). Temperature:_________°F. 1/8 Inch Aileron Cable Tension: Left________Right_______. (f) Inspect all flight control rigging pin placards for legibility and security of attachment, if installed (Ref. Chapter 11-30-00). (g) Inspect underfloor areas for foreign objects, structural damage, loose or missing rivets, cracks and corrosion. (2) BRAKE SYSTEM Zone inspection areas: 121, 122, 730 and 740 Panel inspection area: 6. (a) Inspect brake system components for damage, leakage, cracks and corrosion. (b) Inspect brake master cylinder for operation. (c) Inspect the brake line plumbing for damage, corrosion, leakage and attachment. (d) Inspect the brake pedals and linkage for travel, wear, damage, cracks, corrosion, attachment and operation. (e) Check parking brake for correct release. (3) BRAKE RESERVOIR PRESSURE EQUALIZATION LINE ORIFICE AND FILTER Zone inspection area: 120. Panel inspection area: 1. (a) Clean filter (Ref. Chapter 32-40-00).

05-20-04

Page 3 May 1/12

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Third 200-Hour-Interval Detailed Inspection (Continued) Complete Steps 1 thru 15. (4) RUDDER PEDALS Zone inspection areas: 121 and 122. Panel inspection area: 1 and 18. (a) Inspect the rudder pedals for wear, damage, cracks, corrosion, clearance and attachment. (b) Inspect the rudder pedal arm (Ref. Chapter 27-20-03). (c) Inspect nose landing gear aft steering link rod end (Mechanical Steering Installed). (d) Inspect the pilot’s and copilot’s rudder pedal bellcrank support attach bolts for security. (5) INSTRUMENT PANEL PLUMBING AND WIRING Zone inspection areas: 243, 244, 245, 246, 247, 248 and 249. (a) Inspect instrument panel for damage and attachment. (b) Inspect subpanels for damage and attachment. (c) Inspect placards for proper location and legibility. Inspect for scratches and damage. (d) Inspect shock mounts for damage and attachment. (e) Inspect instruments for damage and attachment. (f) Inspect the instrument plumbing and wiring for damage, attachment, chafing, etc. (6) CONTROL COLUMN Zone inspection areas: 254 and 255. (a) Check control column for damage, cracks, corrosion, attachment and freedom of movement. (b) Perform the CONTROL COLUMN BEARING SUPPORT INSPECTION procedure (Ref. Chapter 27-00-01). (c) Inspect the flight control gust lock to determine if it is the correct part number. Perform the CONTROL LOCK INSPECTION procedures (Ref. Chapter 27-70-00). Check flight control gust lock for positive locking and alignment. (d) Inspect the extension cord for deterioration or damage. (e) Inspect the control wheel switches for operation and damage. (f) Perform the AILERON CONTROL COLUMN INTERCONNECT CABLE TENSION CHECK procedure (Ref. Chapter 27-10-03). Inspect and record the aileron control column interconnect cable tension. Temperature:________ °F. Aileron Control Column Interconnect Cable Tension:____________. (g) Perform the CONTROL COLUMN CLEARANCE INSPECTION procedure (Ref. Chapter 20-00-03).

Page 4 May 1/12

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Inspector

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Third 200-Hour-Interval Detailed Inspection (Continued) Complete Steps 1 thru 15.

Mechanic

Inspector

(h) Perform the BOBWEIGHT AND STOP Inspection procedure (Ref. Chapter 27-30-02). (7) ELECTRICAL EQUIPMENT Zone inspection areas: 221, 222 and 253. (a) On airplanes equipped with optional electric pitch trim, inspect wires for chafing, correct separation and protection from other wires. (b) Inspect all inverter system wiring for connection and chafing. (c) Inspect all wiring for chafing, security, etc. (8) ALTERNATE STATIC AIR SOURCE Zone inspection area: 240. (a) Visually inspect tubing and hardware for security. (9) SEATS (a) Check seats for wear, deterioration, damage, cracks, corrosion and attachment. (b) Check seats for proper operation, engagement and adequate lubrication. (10) SEAT BELTS (a) Check seat belts for wear, cuts, fraying, damage, deterioration and attachment. (11) SEAT TRACKS (a) Check seat tracks for damage, cracks, wear and corrosion (Ref. Chapter 53-40-00). (12) WINDOWS AND WINDOW FRAMES (a) Inspect the windows and window seals for deterioration. (b) Clean the side windows interior side. (c) Inspect the outer windows for chips, excessive crazing, out of contour and other damage. (d) Inspect side window attach frames for security (Ref. Chapter 56-10-00, INSPECTION AND REPAIR OF WINDOW ATTACH FRAMES). (13) PITOT AND STATIC SYSTEM Zone inspection areas: 221. (UC-1 and After), 241, 242, 246, 247 and 248. (a) Visually inspect the pitot/static masts for damage, cracks, corrosion and obstructions.

05-20-04

Page 5 May 1/12

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Third 200-Hour-Interval Detailed Inspection (Continued) Complete Steps 1 thru 15.

Mechanic

Inspector

Mechanic

Inspector

(14) WINDSHIELDS (a) Inspect windshields for antistatic coating and antistatic tab bonding (Ref. Chapter 56-10-00). (b) Inspect windshield weather seal for debonding, cracks and wear. (15) ENGINE CONTROL LEVERS (ALL) (a) Inspect the pedestal power lever stop pin for wear. A groove of 0.03 inch or less is acceptable (Ref. Chapter 76-00-00). (b) Inspect the forward and aft edges of the levers to ensure that the wear does not exceed 0.25 inch into the material. (c) Check the sides of the condition levers for wear. A groove of 0.03 inch or less is acceptable. (d) Inspect the condition control catch (condition lever low idle detent) for wear by checking for positive engagement with the condition levers. If positive engagement with the condition levers does not exist, replace the condition control catch (Ref. Chapter 76-10-00).

B. Cabin Section Third 200-Hour-Interval Detailed Inspection Complete Steps 1 thru 19. (1) FUSELAGE SKINS Zone inspection areas: 261, 262, 271 and 272. (a) Inspect exterior fuselage skins for cracks, scratches, paint blistering, corrosion, damage and loose or missing rivets. If damage is found, check adjacent structure. (2) STRUCTURE (a) Inspect structure under removed floorboards for cracks, corrosion, loose or missing rivets and concealed damage. (3) FLIGHT CONTROL COMPONENTS, CABLES AND PULLEYS Zone inspection areas: 131, 141, 151, 161 and 171. Floor panel inspection area: 16. (a) Inspect the control system components (pushrods, turnbuckles, end fittings, castings, etc.) for bulges, splits, bends, cracks and corrosion. Replace any damaged component. (b) Inspect the control cable pulleys and associated equipment for cracks, corrosion, breaks, wear, attachment, alignment, clearance and operation.

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05-20-04

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Third 200-Hour-Interval Detailed Inspection (Continued) Complete Steps 1 thru 19.

Mechanic

Inspector

(c) Inspect the control cables for clearance, alignment, broken strands and evidence of corrosion. Replace cables that have broken strands or evidence of corrosion. (4) UNDERFLOOR AREAS Zone inspection areas: 131, 132, 133, 141, 142, 143, 151, 152, 153, 161, 162, 163, 171, 172 and 173. Floor panel inspection areas: 6 thru 17. (a) Inspect underfloor areas for foreign objects, structural damage, loose or missing rivets, cracks and corrosion. Clean as required. (5) FLAP MOTOR AND FLEXIBLE DRIVE SHAFTS Zone inspection area: 163. Panel inspection area: 9 (UA-1 and After); 10 (UB-1 and After; UC-1 and After). (a) Inspect for condition and attachment. (b) Lubricate as instructed (Ref. Chapter 12-20-00). (c) Check flap flexible drive shafts for nicks, cuts, cracks and corrosion. (6) AILERON QUADRANT Zone inspection area: 163. Panel inspection area: 16 (between 4th and 5th windows). (a) Check for security, attachment, damage, cracks, corrosion and correct travel. (b) Inspect flight control rigging pin placard for legibility and security of attachment, if installed (Ref. Chapter 11-30-00). (7) BELLY DRAIN VALVES Zone inspection areas: 143, 163, 173, 181, 182, 294, 335 and 436. (a) Inspect for possible obstructions and operation. (8) WINDOWS AND WINDOW FRAMES Zone inspection areas: 271 and 272. (a) Inspect the windows and window seals for deterioration. (b) Inspect the outer windows for chips, excessive crazing and other damage (Ref. Chapter 56-00-00, WINDOWS). (c) Inspect cabin window attach frames, including the escape hatch windows for security (Ref. Chapter 56-10-00, INSPECTION AND REPAIR OF WINDOW ATTACH FRAMES). (9) CABIN ENTRANCE DOOR Zone inspection area: 820 forward, 850 aft. (a) Inspect door, seal, counterbalance mechanism, handrail and handrail cables for damage, broken strands, cracks, scratches, paint blistering, corrosion, condition and security.

05-20-04

Page 7 May 1/12

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Third 200-Hour-Interval Detailed Inspection (Continued) Complete Steps 1 thru 19. (b) Remove the interior door handle and inspect the square shaft for damage and corrosion. (c) Remove upholstery panels and inspect latching mechanism and cables for clearances, damage, broken strands, cracks, corrosion, condition and security of attachment. (d) Measure latching mechanism cable tension (Ref. ADJUSTMENT OF AIRSTAIR DOOR LATCHING MECHANISM procedure and the Airstair Door Cable Tension Graph Chapter 52-10-00). (e) Move inside and outside door handles and check operation. (f) Use the outside door handle and measure the torque required to operate the latching mechanism while closing the door (Ref. ADJUSTMENT OF AIRSTAIR DOOR LATCHING MECHANISM procedure Chapter 52-10-00). (g) Use the outside door handle and measure the torque to operate the latching mechanism while opening the door. (h) Perform the AIRSTAIR DOOR HANDRAIL FUNCTIONAL CHECK procedure (Ref. Chapter 52-10-00). (10) CABIN CARGO DOOR Zone inspection area: 850. (a) Inspect doors, seals and latching mechanism for damage, cracks, corrosion, condition and security of attachment. (b) Check the inside and outside door handles and latch mechanism for proper operation with the door in open and closed position. Ensure the door handles and latch mechanism operate and unlatch with normal effort. (11) EMERGENCY EXIT DOORS Zone inspection areas: 830 and 840. (a) Check inside and outside emergency release handles and latch mechanism for proper operation with the door in the open and closed position. Ensure the door handles and latch mechanism operate and unlatch with normal effort. (b) Check that door opens and closes freely from the inside and outside. (c) Check door for condition, scratches, blistered paint and corrosion and all moving parts for operation. (d) Inspect the hatch seal for condition. Replace the seal if the seal has become hard. (e) Check latching and seal of closed hatch. (f) Remove upholstery, inspect latching mechanism for damage, cracks, corrosion, condition and security of attachment.

Page 8 May 1/12

05-20-04

Mechanic

Inspector

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Third 200-Hour-Interval Detailed Inspection (Continued) Complete Steps 1 thru 19.

Mechanic

Inspector

(12) ELECTRICAL WIRING AND EQUIPMENT Zone inspection areas: 133, 143, 153 and 163. (a) Inspect wiring for chafing and electrical equipment for damage, condition and attachment. (b) Perform the TRIPLE FED BUS DIODES, OPERATIONAL CHECK procedure (Ref. Chapter 24-50-00). (c) Check cabin and compartment lights for condition and attachment. Replace bulbs as necessary. (13) EJECTOR AND DEICER DISTRIBUTOR VALVE Zone inspection area: 153. Panel inspection area: 8 (UA-1 and After; UB-1 and After). 9 (UC-1 and After). (a) Check equipment and plumbing for security, damage, cracks, corrosion and condition. (14) SEATS (a) Check seats for wear, damage, deterioration, cracks, corrosion and attachment. (b) Check seats for proper operation, engagement and adequate lubrication. (15) SEAT BELTS (a) Check seat belts for wear, cuts, fraying, damage, deterioration and attachment. (16) SEAT TRACKS (a) Check seat tracks for damage, cracks, wear and corrosion (Ref. Chapter 25-20-00, PASSENGER SEATS). (17) CONTROL CABLE SEALS Zone inspection areas: 161 and 162. (a) Check for damage, security, cleanliness and lubrication. (18) EVAPORATOR BLOWER MOTORS Zone inspection areas: 153 and 173. (a) Inspect brushes for wear. Replace if required. (19) AUTOPILOT (IF INSTALLED) Zone inspection areas: 163, 171, 311 and 312. Panel inspection areas: 11, 16F, 7 and 8. (a) Check aileron and elevator trim tab autopilot servos for loose or worn bearings and mounting hardware. (b) Verify that the servo mounts are securely mounted to the airframe. (c) Visually inspect the capstan and cable for wear, contamination and proper spool-off.

05-20-04

Page 9 May 1/12

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Third 200-Hour-Interval Detailed Inspection (Continued) Complete Steps 1 thru 19.

Mechanic

Inspector

Mechanic

Inspector

(d) With the autopilot disengaged, operate each control system through its entire range. Observe the servo mounts for any unusual noise, binding, backlash or other mechanical irregularities.

C. Power Plant Third 200-Hour-Interval Detailed Inspection Complete Steps 1 thru 3.

LH

RH

(1) IGNITER PLUGS Zone inspection areas: 400. (a) Inspect the igniter plugs for condition and erosion (Ref. Chapter 74-00-00 and the Pratt and Whitney PT6A-65B Engine Maintenance Manual). (2) ENGINE OIL FILTER Zone inspection area: 400. (a) Inspect the oil filter for contamination (Ref. Pratt and Whitney PT6A-65B Engine Maintenance Manual P/N 3032842, Chapter 72-00-00, Table 601 Periodic Inspection). (3) PLUMBING (a) Inspect plumbing and associated equipment for condition, damage, cracks, corrosion and attachment.

D. General Service Items Third 200-Hour-Interval Detailed Inspection Complete Steps 1 thru 9. (1) AIRPLANE LUBRICATION (a) Lubricate as necessary (Ref. Chapter 12-20-00, SCHEDULED SERVICING). (2) ELECTRONIC FLIGHT INSTRUMENT SYSTEM (EFIS) (IF INSTALLED) Zone inspection area: 248. (a) Verify operation of Electronic Attitude Director Indicator (EADI) and Electronic Horizontal Situation Indicator (EHSI) tube fans by listening for fan operation. EADI and EHSI tubes must be on (Ref. the appropriate Pilot's Operating Handbook/Airplane Flight Manual).

Page 10 May 1/12

05-20-04

Mechanic

Inspector

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Third 200-Hour-Interval Detailed Inspection (Continued) Complete Steps 1 thru 9.

Mechanic

Inspector

(3) EVAPORATOR FILTERS Zone inspection area: 153 and 173. Panel inspection area: 7 and 11 (UA-1 and After). 8 and 12 (UB-1 and After; UC-1 and After). (a) Replace the evaporator filters (Ref. Chapter 21-52-01, EVAPORATOR FILTER REPLACEMENT). (4) VACUUM REGULATOR VALVE FILTER Zone inspection area: 812. (a) Clean or replace the filter (Ref. Chapter 37-00-00, VACUUM REGULATOR VALVE FILTER SERVICING). (5) PITOT, STATIC AND PRESSURIZATION SYSTEMS Zone inspection areas: 221 (UC-1 and After), 246 and 247. (a) Drain system. Close the drains when complete (Ref. Model 1900/1900C Flight Manual Supplements 114-590021-41 and 114-590021-87 as applicable). Close the drains when completed. (6) ACCESS PANELS (a) Check all panels removed for this inspection for fit, attachment, scratches, paint blistering and corrosion. (7) PLACARDS (a) Verify all placards are in place and legible (Ref. Pilots Operating Handbook, Airplane Flight Manual, Chapter 11-20-00 (UA-1 and After, UB-1 and After) and Chapter 11-21-00 (UC-1 and After)). (8) DEICER BOOTS Zone inspection areas: 351, 352, 511, 531, 541, 611, 631 and 641. (a) Visually check deicer boots for cracks, gaps, damage and attachment. (b) Inspect deicer boot edge sealant for looseness, gaps and damage. Apply as necessary. (9) AIR CYCLE MACHINE Zone inspection areas: 511. Panel inspection area: 57 (UA-1 and After; UB-1 and After). 18 (UC-1 and After). (a) Change air cycle machine oil (Ref. Hamilton Standard B-1900 Refrigeration Package Maintenance Manual and Parts List in the Model 1900 Airliner Series Component Maintenance Manual, Chapter 21).

05-20-04

Page 11 May 1/12

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

E. Operational Inspection Third 200-Hour-Interval Detailed Inspection Complete Steps 1 thru 5.

Mechanic

Inspector

(1) ENVIRONMENTAL TEST (a) Perform the BLEED AIR TEMPERATURE OPERATIONAL CHECK procedure (Ref. Chapter 21-10-00 or Chapter 21-11-00). (b) Perform the BLEED AIR PRESSURE OPERATIONAL CHECK procedure (Ref. Chapter 21-10-00 or Chapter 21-11-00). (2) PROPELLER DEICER (a) Perform the PROPELLER DEICER SYSTEM INSPECTION procedure (Ref. Chapter 30-60-00). (3) GROUND INSPECTION RUN Depending on the maintenance performed and components replaced, a GROUND PERFORMANCE CHECK PROCEDURE may be required in lieu of this Inspection Run. Refer to the applicable maintenance procedures. (a) Start engines and allow the oil temperature to increase into the operating range. (b) Run engines at a minimum of 80% N1 long enough for engine indicators to stabilize. (c) Shut down the engines and inspect for attachment and security of all components and for oil and fuel leaks. (4) INVERTER SYSTEM (a) Perform the INVERTER POWER SELECT RELAY CHECK procedure (Ref. Chapter 24-20-00). (b) Perform the INVERTER BLOWER FAN OPERATIONAL CHECK procedure (Ref. Chapter 24-20-00). (5) EXTERNAL POWER Zone inspection area: 522. (a) Check the external power relay for operation (rotate the voltmeter select switch to the EXT PWR position and check for external power voltage).

INSPECTION COMPLETED I certify that a Detailed Inspection was performed in accordance with the Continuous Inspection Program and that the airplane is approved for return to service. CREW CHIEF________________________________________________ Page 12 May 1/12

05-20-04

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

CONTINUOUS INSPECTION WORKSHEET Inspection Type_________________________________Inspection Number__________________________ Model__________L Eng Time___________Cycles__________R Eng Time___________Cycles___________ A/C Time_______________Cycles________________A/C Serial_________________Date_______________

Item

Description

Mech

05-20-04

Crew Chief

Q.C.

Page 13 May 1/12

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Inflight Worksheet (Page 1 of 2) S/N________________N Number__________________Date___________________NO._________________ Flt. No.

Pilot

Copilot

TO

Time

LND

Time

No. LND

Flt. Time

Today’s Total Previous Total Total Fuel Type

Engine Oil Gallon

Left

Cruise Condition

Right

Data O.A.T. PA IAS ITT Torque Prop. RPM N1 RPM Fuel Flow Fuel Pressure Oil Pressure Oil Temp.

Page 14 May 1/12

05-20-04

Left

Right

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Inflight Worksheet (Page 2 of 2) Discrepancy Worksheet S/N________________N Number__________________Date___________________NO._________________

A/C Hrs

Date

Discrepancy

Corrective Action

Date

Mechanic Inspector

05-20-04

Page 15 May 1/12

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Major Maintenance Worksheet Component

Page 16 May 1/12

Date

Reason for Replacement

05-20-04

Replacement Part Number Serial Number

Next Overhaul A/C Hours or Cycles Date

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

TIME LIMITS/MAINTENANCE CHECKS FOURTH 200-HOUR-INTERVAL DETAILED INSPECTION DETAILED INSPECTION PROCEDURES

05-20-05

00

1. GENERAL A. Forms Required (1) Fourth 200-Hour-Interval Detailed Inspection. (2) Model 1900/1900C Airliner Routine Inspection. (3) Continuous Inspection Worksheet. NOTE: A Routine Inspection must be conducted in conjunction with each Detailed Inspection to comply with Continuous Inspection Regulations.

B. Reference Material (1) Model 1900/1900C Airliner Maintenance Manual P/N 114-590021-7. (2) Model 1900 Airliner Series Component Maintenance Manual P/N 114-590021-11. (3) Model 1900 Airliner Series Structural Repair Manual, P/N 114-590021-9. (4) Model 1900 Airliner Wiring Diagram Manual P/N 114-590032-3 (UA-1 and After), 114-590021-13 (UB-1 and After) and 114-590021-61 (UC-1 and After). (5) Model 1900 Airliner Pilot’s Operating Handbook/Airplane Flight Manual, P/N 114-590021-3. (6) Model 1900C Airliner Pilot’s Operating Handbook/Airplane Flight Manual, P/N 114-590021-57.

C. Inspection Procedures (1) Fill out the heading on each form in its entirety. (2) The mechanic checks each item on the inspection form and initials the form in the space provided. (3) List all discrepancies found during the inspection on the Continuous Inspection Worksheet. NOTE: Check all In-Flight Worksheets turned in since the last inspection for discrepancies that have not yet been worked off. (4) Each discrepancy is to be signed off by the mechanic, crew chief and a Quality Control Inspector when the discrepancy has been corrected. (5) In the spaces provided on the Major Maintenance Worksheet, the mechanic is to list all components which are removed from the airplane for overhaul or replacement, then add the Part Number and Serial Number of the component which is installed. (6) The Quality Control Inspector will stamp off each item on the inspection form to complete the inspection.

05-20-05

Page 1 May 1/10

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (7) When the inspection has been completed, the crew chief will sign the “INSPECTION COMPLETED” block at the end of the inspection sheet.

2. DETAILED INSPECTION PROCEDURES Inspection Date___________________________________________ Airplane Serial___________________ Airframe______________________ Cycles____________ LH Engine Hrs_____________ Cycles____________ RH Engine Hrs______________ Cycles____________ LH Engine Power Module S/N_______________________________________________ RH Engine Power Module S/N_______________________________________________ LH Engine Gas Generator S/N_______________________________________________ RH Engine Gas Generator S/N_______________________________________________

NOTE: Corrosion detected while performing this detailed inspection may be treated in accordance with Chapter 20-09-00. To minimize the possibility of foreign object damage to engines, observe the following maintenance practices: •

Ensure all loose materials (rivets, screws, safety wire, etc.) are removed from engine cowling area after maintenance.

•

Maintain clean ramp and taxi areas.

•

Running at maximum power with the airplane stationary should be minimized and done only on a clean ramp.

•

Propeller reverse operation for backing the airplane should be avoided.

•

Avoid operation in dust and sand storms.

•

Do not operate engines in feather, except during external power starts and feather checks.

Prior to beginning this inspection, the following access panels must be removed: UA-1 and After; Wing Access - 10, 11, 12, 15, 56 and 57. Fuselage Access - 4, 13, 14 and 15. Floor Access - 7 and 11. UB-1 and After; Wing Access - 10, 11, 12, 15, 56 and 57. Fuselage Access - 4, 13, 14 and 15. Floor Access - 8 and 12. UC-1 and After; Wing Access - 10, 11, 12, 15, 18 and 25. Fuselage Access - 4, 13, 14 and 15. Floor Access - 8 and 12. For zone and access panel locations (Ref. Chapter 06).

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05-20-05

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

A. Environmental Systems Fourth 200-Hour-Interval Detailed Inspection Complete Steps 1 thru 8.

Mechanic

Inspector

(1) EVAPORATOR FILTER Zone inspection areas: 153 and 173. Floor panel inspection areas: 7, 11 (UA-1 and After). 8 and 12 (UB-1 and After, UC-1 and After). (a) Replace the evaporator filters (Ref. Chapter 21-52-01). (2) VACUUM REGULATOR VALVE FILTER Zone inspection area: 812. (a) Clean or replace filter (Ref. Chapter 37-00-00). (3) PRESSURIZATION CONTROLLER Zone inspection area: 243. (a) Inspect for security of attachment and damage. (b) Check wiring for damage, chafing and security. (c) Check plumbing for leaks, damage and attachment. (4) PRESSURIZATION CONTROLLER FILTER Zone inspection area: 243. (a) Inspect and clean filter (Ref. Chapter 21-30-02). (5) PRESSURIZATION SYSTEM DRAIN VALVE Zone inspection area: 242. (a) On airplanes with the valve installed, open drain valve to remove condensation in pressure lines. Close drain when completed. (6) PNEUMATIC RELAY FILTER Zone inspection area: 243. (a) Inspect and clean filter (Ref. Chapter 21-30-03). (7) AIR CYCLE MACHINE FOG NOZZLE AND FILTER Zone inspection area: 512. Wing panel inspection areas: 11 and 15. (a) Clean the air cycle machine fog nozzle and filter (if installed) (Ref. Chapter 21-51-02). (8) AIR-CONDITIONING CONDENSER AND BLOWER Zone inspection area: 611. Wing panel inspection area: 56 (UA-1 and After, UB-1 and After). 25 (UC-1 and After). (a) Check the condenser, blower and associated plumbing for leaks, cracks, corrosion, damage and security of attachment.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Fourth 200-Hour-Interval Detailed Inspection (Continued) Complete Steps 1 thru 8.

Mechanic

Inspector

Mechanic

Inspector

(b) Inspect the inlet guard for security of attachment and for broken strands. (c) Inspect the impeller for security to the shaft and ease of rotation. (d) Inspect the standoffs for security and tightness. (e) Inspect the guide vanes in the blower housing assembly for cracking, corrosion and security of attachment.

B. Nose Section Fourth 200-Hour-Interval Detailed Inspection Complete Steps 1 thru 6. (1) SKINS (a) Inspect skins for cracks, scratches, paint blistering, corrosion, damage and loose or missing rivets. (2) STRUCTURES (a) Inspect for cracks, corrosion, loose rivets and concealed damage where access panels have been removed. (3) PLACARDS (a) Verify that all nose interior and exterior placards are in place. Inspect for scratches, damage and legibility. (4) RADIO EQUIPMENT Zone inspection areas: 211 and 212. Fuselage panel inspection areas: 13, 14 and 15. (a) Inspect radio rack structure; check security of units in their mounts. (5) INSTRUMENT AIR FILTER Zone inspection area: 212. (a) Replace the filter (Ref. Chapter 37-00-00). (6) AVIONICS COMPARTMENT DOOR, FASTENERS AND SEAL Zone inspection area: 811 and 812. (a) Inspect seals for deterioration. (b) Inspect doors for damage, scratches, paint blistering, corrosion, operation and deterioration. (c) Inspect for loose or missing fasteners.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

C. Power Plant Fourth 200-Hour-Interval Detailed Inspection Complete Steps 1 thru 4.

Mechanic LH

Inspector

RH

(1) IGNITER PLUGS Zone inspection areas: 400. (a) Inspect the igniter plugs for condition and erosion (Ref. Pratt and Whitney PT6A-65B Engine Maintenance Manual or Chapter 74-00-00 of this manual). (2) ENGINE OIL FILTER Zone inspection areas: 400. (a) Inspect the oil filter for contamination (Ref. Pratt and Whitney PT6A-65B Engine Maintenance Manual P/N 3032842, Chapter 72-00-00, Table 601 Periodic Inspection). (3) FUEL NOZZLES Zone inspection areas: 400. (a) Inspect and clean the fuel nozzles using either the “in situ” method or by bench check (Ref. Pratt and Whitney PT6A-65B Engine Maintenance Manual P/N 3032842). (b) Borescope inspect the engine hot section at this time (Ref. Pratt and Whitney PT6A-65B Engine Maintenance Manual P/N 3032842). (4) STARTER-GENERATOR (a) Inspect brushes for indication of excessive wear or damage (determine wear by observing diagonal groove on brush) (Ref. Chapter 24-30-01 and to the Starter-Generator Manufacturer’s Maintenance Manual in Chapter 24 of the Model 1900 Airliner Series Component Maintenance Manual). (b) Inspect inlet duct and cooling cap for cracks, corrosion or obstruction.

D. General Service Items Fourth 200-Hour-Interval Detailed Inspection Complete Steps 1 thru 6.

Mechanic

Inspector

(1) PITOT AND STATIC SYSTEM Zone inspection areas: 221 (UC-1 and After), 246 and 247. (a) Drain System (Ref. Flight Manual Supplements 114-590021-41 and 114-590021-87 as applicable). Close drains when completed.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Fourth 200-Hour-Interval Detailed Inspection (Continued) Complete Steps 1 thru 6.

Mechanic

Inspector

Mechanic

Inspector

(2) ELECTRONIC FLIGHT INSTRUMENT SYSTEM (EFIS) (IF INSTALLED) Zone inspection area: 248. (a) Verify operation of Electronic Attitude Director Indicator (EADI) and Electronic Horizontal Situation Indicator (EHSI) tube fans by listening for fan operation. EADI and EHSI tubes must be on (Ref. the appropriate Pilot's Operating Handbook/ Airplane Flight Manual). (3) AIRPLANE LUBRICATION (a) Lubricate as necessary in accordance with the LUBRICATION SCHEDULE (Ref. Chapter 12-20-00). (4) PLACARDS (a) Verify all placards are in place and legible (Ref. Pilots Operating Handbook, Airplane Flight Manual, Chapter 11-20-00 (UA-1 and after, UB-1 and after) and Chapter 11-21-00 (UC-1 and after)). (5) WINDSHIELDS (a) Inspect windshield weather seal for damage, debonding, cracks and wear. (6) POWER PACK BLEED AIR FILTER Zone inspection area: 511. Wing panel inspection area: 10. (a) Clean filter (Ref. Chapter 32-30-06).

E. Operational Inspection Fourth 200-Hour-Interval Detailed Inspection Complete Steps 1 thru 21. NOTE The following Operational Inspection procedures are to be applied during start and run of the engine. CAUTION Do not place the power levers into reverse unless engines are running. (1) ENGINE CONTROLS (a) Check for freedom of movement, full travel and friction lock. (2) OIL (a) Check that the pressure and temperature are within limits (Ref. the appropriate Pilot's Operating Handbook/Airplane Flight Manual).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Fourth 200-Hour-Interval Detailed Inspection (Continued) Complete Steps 1 thru 21.

Mechanic

Inspector

(3) PROPELLER GOVERNOR (a) Check the operation and feathering (Ref. the appropriate Pilot's Operating Handbook/Airplane Flight Manual). (4) PROPELLER SYNCHROPHASER (a) Check the operation (Ref. Chapter 61-22-00). (5) PROPELLER DEICER (a) Perform the PROPELLER DEICER SYSTEM INSPECTION procedure (Ref. Chapter 30-60-00). (6) AUTOFEATHERING SYSTEM (a) Perform the AUTOFEATHER OPERATIONAL CHECK procedure (Ref. Chapter 61-21-00). (7) GROUND INSPECTION RUN Depending on the maintenance performed and components replaced, a GROUND PERFORMANCE CHECK PROCEDURE may be required in lieu of this Inspection Run. Refer to the applicable maintenance procedures. (a) Start engines and allow the oil temperature to increase into the operating range. (b) Run engines at a minimum of 80% N1 long enough for engine indicators to stabilize. (c) Shut down the engines and inspect for attachment and security of all components and for oil and fuel leaks. (8) STARTER-GENERATOR (a) Check for output of 28.25 ± 0.25 VDC (Ref. Chapter 24-30-00). (9) FUEL BOOST PUMPS (a) Check operation of the electric pumps (Ref. the appropriate Pilot's Operating Handbook/Airplane Flight Manual). (10) FUEL CROSS-TRANSFER VALVES (a) Check operation (Ref. the appropriate Pilot's Operating Handbook/Airplane Flight Manual). (11) FUEL FIREWALL SHUTOFF VALVES (a) Perform the FIREWALL FUEL SHUTOFF VALVE FUNCTIONAL CHECK procedure (UA-1 and After; UB-1 and After (Ref. Chapter 28-20-05)) or (UC-1 and After (Ref. Chapter 28-21-05)).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Fourth 200-Hour-Interval Detailed Inspection (Continued) Complete Steps 1 thru 21. (12) VACUUM SYSTEM (a) Perform vacuum regulator valve adjustment (Ref. Chapter 37-00-00). (13) ENVIRONMENTAL TEST (a) Perform the BLEED AIR TEMPERATURE OPERATIONAL CHECK procedure (Ref. Chapter 21-10-00 or Chapter 21-11-00). (b) Perform the BLEED AIR PRESSURE OPERATIONAL CHECK procedure (Ref. Chapter 21-10-00 or Chapter 21-11-00). (14) SURFACE DEICERS (a) Check for inflation and cycling. (b) Perform the SURFACE DEICER OPERATIONAL CHECKS procedure (Ref. Chapter 30-10-00). (15) PRESSURIZATION SYSTEM (a) Check operation (Ref. Chapter 21-30-00) of this manual. (16) PRESSURIZATION SYSTEM DRAIN VALVE (a) Open drain valve to remove condensation in pressure lines. (17) REFRIGERANT LEVEL (a) For aircraft with R12 refrigerant ONLY: Check level through the sight glass or check refrigerant pressure, with the LH engine shut down, the RH engine running above 62% N1 and the air-conditioner ON in either the AUTO or MANUAL mode. Ambient temperature must be above 50°F (Ref. Chapter 21-52-00). (b) For aircraft with R134A refrigerant: check refrigerant pressure, with the LH engine shut down, the RH engine running above 62% N1 and the air-conditioner ON in either the AUTO or MANUAL mode. Ambient temperature must be above 50°F (Ref. Chapter 21-52-00). (18) ENVIRONMENTAL VAPOR CYCLE SYSTEM AND AIR CYCLE MACHINE (a) Check operation when the switch is in the AUTO or MANUAL position. Ambient temperature must be above 50°F. (b) Check operation of all outlets and ease of operation of all controls (Ref. Chapter 21-50-00) for vapor cycle system and (Ref. Chapter 21-52-00) for air cycle. (19) CONDITION LEVERS (a) Check for clean shutdown at IDLE-CUTOFF.

Page 8 May 1/10

05-20-05

Mechanic

Inspector

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Fourth 200-Hour-Interval Detailed Inspection (Continued) Complete Steps 1 thru 21.

Mechanic

Inspector

(20) INVERTER SYSTEM (a) Perform the INVERTER POWER SELECT RELAY CHECK procedure (Ref. Chapter 24-20-00). (b) Perform the INVERTER BLOWER FAN OPERATIONAL CHECK procedure (Ref. Chapter 24-20-00). (21) EXTERNAL POWER Zone inspection area: 253. (a) Check the external power relay for operation (rotate the voltmeter select switch to the EXT PWR position and check for external power voltage).

INSPECTION COMPLETED I certify that a Detailed Inspection was performed in accordance with the Continuous Inspection Program and that the airplane is approved for return to service. CREW CHIEF________________________________________________

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

CONTINUOUS INSPECTION WORKSHEET Inspection Type_________________________________Inspection Number__________________________ Model__________L Eng Time___________Cycles__________R Eng Time___________Cycles___________ A/C Time_______________Cycles________________A/C Serial_________________Date_______________

Item

Page 10 May 1/10

Description

05-20-05

Mech

Crew Chief

Q.C.

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Inflight Worksheet (Page 1 of 2) S/N________________N Number__________________Date___________________NO._________________ Flt. No.

Pilot

Copilot

TO

Time

LND

Time

No. LND

Flt. Time

Today’s Total Previous Total Total Fuel Type

Engine Oil Gallon

Left

Cruise Condition

Right

Data

Left

Right

O.A.T. PA IAS ITT Torque Prop. RPM N1 RPM Fuel Flow Fuel Pressure Oil Pressure Oil Temp.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Inflight Worksheet (Page 2 of 2) Discrepancy Worksheet S/N________________N Number__________________Date___________________NO._________________

A/C Hrs

Page 12 May 1/10

Date

Discrepancy

05-20-05

Corrective Action

Date

Mechanic Inspector

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Major Maintenance Worksheet Component

Date

Reason for Replacement

Replacement Part Number Serial Number

05-20-05

Next Overhaul A/C Hours or Cycles Date

Page 13 May 1/10

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

TIME LIMITS/MAINTENANCE CHECKS FIFTH 200-HOUR-INTERVAL DETAILED INSPECTION DETAILED INSPECTION PROCEDURES

05-20-06 00

1. GENERAL A. Forms Required (1) Fifth 200-Hour Interval Detailed Inspection. (2) Model 1900/1900C Airliner Routine Inspection. (3) Continuous Inspection Worksheet. NOTE: A Routine Inspection must be conducted in conjunction with each Detailed Inspection to comply with Continuous Inspection Regulations.

B. Reference Material (1) Model 1900/1900C Airliner Maintenance Manual P/N 114-590021-7B. (2) Model 1900 Airliner Series Component Maintenance Manual P/N 114-590021-11. (3) Model 1900 Airliner Series Structural Repair Manual P/N 114-590021-9. (4) Model 1900 Airliner Wiring Diagram Manual P/N 114-590032-3 (UA-1 and After), 114-590021-13 (UB-1 and After), 114-590021-61 (UC-1 and After). (5) Model 1900 Airliner Pilot’s Operating Handbook/Airplane Flight Manual, P/N 114-590021-3. (6) Model 1900C Airliner Pilot’s Operating Handbook/Airplane Flight Manual, P/N 114-590021-57. (7) Model 1900/1900C Flight Manual Supplements, P/N 114-590021-41 and -87.

C. Inspection Procedures (1) Fill out the heading on each form in its entirety. (2) The mechanic checks each item on the inspection form and initials the form in the space provided. (3) List all discrepancies found during the inspection on the Continuous Inspection Worksheet. NOTE: Check all In-Flight Worksheets turned in since the last inspection for discrepancies that have not yet been worked off. (4) Each discrepancy is to be signed off by the mechanic, crew chief and a Quality Control Inspector when the discrepancy has been corrected. (5) In the spaces provided on the Major Maintenance Worksheet, the mechanic is to list all components which are removed from the airplane for overhaul or replacement and the Part Number and Serial Number of the component which is installed. (6) The Quality Control Inspector will stamp off each item on the inspection form to complete the inspection.

05-20-06

Page 1 Aug 1/12

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (7) When the inspection has been completed, the crew chief will sign the “INSPECTION COMPLETED” block at the end of the inspection sheet.

2. DETAILED INSPECTION PROCEDURES Inspection Date___________________________________________ Airplane Serial___________________ Airframe______________________ Cycles____________ LH Engine Hrs_____________ Cycles____________ RH Engine Hrs______________ Cycles____________ LH Engine Power Module S/N_______________________________________________ RH Engine Power Module S/N_______________________________________________ LH Engine Gas Generator S/N_______________________________________________ RH Engine Gas Generator S/N_______________________________________________

NOTE: Corrosion detected while performing this detailed inspection may be treated in accordance with Chapter 20-09-00. To minimize the possibility of foreign object damage to engines, observe the following maintenance practices: •

Ensure all loose materials (rivets, screws, safety wire, etc.) are removed from engine cowling area after maintenance.

•

Maintain clean ramp and taxi areas.

•

Running at maximum power with the airplane stationary should be minimized and done only on a clean ramp.

•

Propeller reverse operation for backing the airplane should be avoided.

•

Avoid operation in dust and sand storms.

•

Do not operate engines in feather, except during external power starts and feather checks.

Prior to beginning this inspection, the following access panels must be removed: UA-1 and After: Wing Access - 12, 13, 15, 55 and 57. Floor Access - 7 and 11. UB-1 and After: Wing Access - 12, 13, 15, 55 and 57. Floor Access - 8 and 12. UC-1 and After: Wing Access - 12, 13, 15, 18 and 24. Floor Access - 8 and 12. For zone and access panel locations (Ref. Chapter 06).

Page 2 Aug 1/12

05-20-06

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

A. Forward Left Hand Center Section Fifth 200-Hour-Interval Detailed Inspection Complete Steps 1 thru 3.

Mechanic

Inspector

Mechanic

Inspector

(1) STRUCTURE Zone inspection areas: 511. (a) Inspect structure for cracks, scratches, corrosion, loose rivets and damage. (2) PLUMBING AND WIRING (a) Inspect plumbing and wiring for chafing, leaks and security. (3) AIR CYCLE MACHINE Zone inspection area: 511. Wing panel inspection area: 57 (UA-1 and After; UB-1 and After). 18 (UC-1 and After). (a) Change air cycle machine oil (Ref. Hamilton Standard B-1900 Refrigeration Package Maintenance Manual and Parts List in the Model 1900 Airliner Series Component Maintenance Manual, Chapter 21).

B. Main Landing Gear Fifth 200-Hour-Interval Detailed Inspection Complete Steps 1 thru 10.

LH

RH

(1) BRAKES Zone inspection areas: 730 and 740. (a) Check brake discs and pads for wear, damage, corrosion and security (Ref. Chapter 32-40-00). (b) Inspect lines for damage, leaks, cracks, corrosion and security. (2) BRAKE DEICER (IF INSTALLED) Zone inspection areas: 730 and 740. (a) Check lines, hose and connections for cracks, leaks, corrosion and security of attachment. (b) Check manifold for cleanliness and unobstructed orifices.

05-20-06

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Fifth 200-Hour-Interval Detailed Inspection (Continued) Complete Steps 1 thru 10.

Mechanic LH

(3) ACTUATOR Zone inspection areas: 730 and 740. (a) Check actuators for cracks, corrosion, damage and leaks. (b) Inspect support brackets for cracks, corrosion, damage and loose or missing rivets. (4) MAIN LANDING GEAR SHOCK ABSORBER (STRUT ASSEMBLY) Zone inspection areas: 730 and 740. (a) Inspect shock absorber (strut assembly) and components for damage, cracks, leaks, corrosion and attachment (Ref. Chapter 32-30-00). (b) Inspect shock absorber (strut assembly) for correct inflation and leakage. If signs of leakage are apparent perform the MAIN LANDING GEAR SHOCK ABSORBER SERVICING procedure (Ref. Chapter 32-10-00). (c) Check gland nut at base of the main strut upper brace assembly for possible looseness and abnormal wear. (d) Inspect the sealant location at the main landing gear piston and socket interface for corrosion or rust. Sealer should cover any non-chromed area of the piston. If any non-chromed portion of the piston is exposed, or if seal is damaged, worn or deteriorated, or if corrosion or rust is present, perform the MAIN LANDING GEAR SOCKET/ PISTON SEAL REPAIR procedure (Ref. 1900 Airliner Series Component Maintenance Manual, Chapter 32-10-00). (5) TRUNNION BOLTS Zone inspection areas: 730 and 740. (a) Inspect for proper security and condition of bolts, nuts and cotter pins. If a cotter pin is missing, torque the trunnion bolt nut and install a new cotter pin by performing the applicable Steps of the MAIN LANDING GEAR INSTALLATION procedure (Ref. Chapter 32-10-00). (b) Visually inspect for wear, cracks and corrosion. (6) LANDING GEAR KEEL Zone inspection areas: 700. (a) Perform LANDING GEAR KEEL INSPECTION procedure (Ref. Chapter 57-10-00).

Page 4 Aug 1/12

05-20-06

RH

Inspector

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Fifth 200-Hour-Interval Detailed Inspection (Continued) Complete Steps 1 thru 10.

Mechanic LH

Inspector

RH

(7) DRAG LEG Zone inspection areas: 730 and 740. (a) Visually inspect for wear, distortion, cracks and corrosion. (b) Check security of attach fittings. (8) ELECTRICAL Zone inspection areas: 730 and 740. (a) Check attachment of switches. Clean dirt from terminals and connectors as required. (b) Check wiring for damage, chafing and security. (9) LANDING GEAR HOSES Zone inspection areas: 730 and 740. (a) Check for damage, cracks, leaks, deterioration and security. Replace as necessary.

C. Nose Landing Gear Fifth 200-Hour-Interval Detailed Inspection Complete Steps 1 thru 11.

Mechanic

Inspector

(1) SHIMMY DAMPER (WITHOUT POWER STEERING INSTALLED) Zone inspection area: 710. (a) Inspect for damage, cracks, corrosion, leakage and attachment. Refill if necessary (Ref. Chapter 32-20-00 SHIMMY DAMPER FLUID CHECK (MANUAL STEERING ONLY) procedure and/or SHIMMY DAMPER SERVICING (MANUAL STEERING ONLY) procedure). (2) ACTUATOR Zone inspection area: 710. (a) Visually check actuator for damage, cracks, corrosion and leakage. (b) Inspect support bracket for damage, cracks, corrosion and loose or missing rivets.

05-20-06

Page 5 Aug 1/12

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Fifth 200-Hour-Interval Detailed Inspection (Continued) Complete Steps 1 thru 11. (3) STEERING LINKAGE (WITH MECHANICAL STEERING ONLY) Zone inspection area: 710. (a) Check nose steering mechanism for damage, cracks, corrosion, attachment and correct adjustment (Ref. Chapter 32-50-00). (b) Remove tie-wrap and inspect aft steering linkage and boot for wear, damage and chafing under boot (Ref. Chapter 32-50-00). (c) Inspect forward steering link (with boot) and boot for wear, damage and chafing under boot. Perform the FORWARD STEERING LINK (WITH BOOT) INSPECTION procedure (Ref. Chapter 32-50-00). (d) Disconnect and inspect nose steering disconnect actuator wiring receptacle plug located in left wheel well keel for corrosion. (e) Visually inspect nose landing gear steering disconnect actuator attaching hardware for evidence of looseness, corrosion, or missing fasteners (Ref. Chapter 32-50-00). (4) NOSE GEAR BRACE STEERING STOP LUGS Zone inspection area: 710. (a) Inspect for cracks, damage or distortion. Should cracks be suspect, perform FLUORESCENT LIQUID PENETRANT INSPECTION procedure (Ref. Chapter 20-10-00). (b) Inspect for proper lubrication of bolts (Ref. NOSE LANDING GEAR LUBRICATION procedure Chapter 12-20-00). (5) NOSE GEAR STEERING STOP Zone inspection area: 710. (a) Inspect steering stop for damage and distortion. (b) Inspect steering stop bolts for proper torque. Bolts should be able to rotate with finger pressure. (6) NOSE GEAR SHOCK ABSORBER (STRUT ASSEMBLY) Zone inspection area: 710. (a) Inspect shock absorber (strut assembly) and components for damage and attachment (Ref. Chapter 32-20-00). (b) Inspect shock absorber (strut assembly) for correct inflation and leakage. If signs of leakage are apparent perform the NOSE GEAR SHOCK ABSORBER SERVICING procedure (Ref. Chapter 32-20-00).

Page 6 Aug 1/12

05-20-06

Mechanic

Inspector

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Fifth 200-Hour-Interval Detailed Inspection (Continued) Complete Steps 1 thru 11.

Mechanic

Inspector

(7) ELECTRICAL Zone inspection area: 710. (a) Check attachment of switches; clean dirt from terminals and connectors, as required; check wiring for damage and chafing. (8) TRUNNION BOLT Zone inspection area: 710. (a) Inspect for proper security and condition of bolts, nuts and cotter pins. If a cotter pin is missing, torque the trunnion bolt nut and install a new cotter pin (Ref. the applicable Steps of the NOSE LANDING GEAR INSTALLATION procedure Chapter 32-20-00). (9) DRAG BRACE Zone inspection area: 710. (a) Check for wear in the lower drag leg attach lug hole on the nose gear brace (Ref. Chapter 32 in the Model 1900 Airliner Series Component Maintenance Manual, P/N 114-590021-11). Visually inspect for damage, cracks, corrosion and security of attachment. (b) Inspect bolts for freedom of movement (Ref. NOSE LANDING GEAR DRAG BRACE BOLT INSPECTION procedure Chapter 32-20-00). (10) POWER STEERING (IF INSTALLED) Zone inspection area: 710. (a) Replace filter (Ref. SYSTEM FILTER REPLACEMENT (UA-1 AND AFTER; UB-1 AND AFTER) procedure Chapter 32-51-00 or HYDRAULIC FILTER SERVICING (UC-1 AND AFTER) procedure Chapter 32-52-00). (b) Perform ACTUATOR MAINTENANCE CHECKS (UA-1 AND AFTER; UB-1 AND AFTER) procedure (Ref. Chapter 32-51-00). (11) LANDING GEAR HOSES (a) Check for damage, cracks, leaks, deterioration and security. Replace as necessary. (12) ANTISKID BRAKES (IF INSTALLED) Zone inspection area: 710. (a) Check operation, charge accumulator as required and replace filter (Ref. Chapter 32-41-00). (b) Remove antiskid hydraulic line/antiskid accumulator cover and inspect for damage, cracks, leaks, deterioration and security.

05-20-06

Page 7 Aug 1/12

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

D. Landing Gear Retraction Fifth 200-Hour-Interval Detailed Inspection Complete Steps 1 thru 8.

Mechanic

Inspector

NOTE Battery voltage is not sufficient to properly cycle the landing gear, use only an external power source capable of delivering and maintaining 28.25 ± 0.25 volts throughout the extension and retraction cycles when performing the landing gear retraction inspection. (1) LANDING GEAR HYDRAULIC POWER PACK (a) Perform the POWER PACK FLUID LEVEL SENSOR FUNCTIONAL TEST procedure (Ref. Chapter 32-30-08). (b) Clean power pack filter screens and replace power pack filter (Ref. Chapter 32-30-06). (2) RETRACT MECHANISM (a) Check retraction system for operation of all components through at least two complete cycles (Ref. Chapter 32-30-00). (b) Check for unusual noises and evidence of binding. (3) DOORS AND LINKAGE Zone inspection areas: 710, 730 and 740. (a) Check door for proper operation, fit and rigging. Inspect for damage, cracks, paint blistering, corrosion and attachment. (4) POSITION INDICATORS Zone inspection areas: 245, 710, 730 and 740. (a) Check for security and adjustment of switches, loose or chafing wires and correct indication. (5) WARNING HORN (a) Perform the LANDING GEAR WARNING HORN CHECK procedure (Ref. Chapter 32-60-06). (6) SAFETY SWITCH Zone inspection areas: 730 and 740. (a) Check for security of attachment. (7) ACTUATORS Zone inspection areas: 710, 730 and 740. (a) Check for noise, damage, cracks, visual leaks, corrosion, binding and correct rigging (Ref. Chapter 32-30-00).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Fifth 200-Hour-Interval Detailed Inspection (Continued) Complete Steps 1 thru 8.

Mechanic

Inspector

Mechanic

Inspector

(8) EMERGENCY EXTENSION Zone inspection areas: 121, 710, 730 and 740. (a) Perform LANDING GEAR HAND PUMP CYCLING procedure (Ref. Chapter 32-30-00).

E. Power Plant Fifth 200-Hour-Interval Detailed Inspection Complete Steps 1 thru 10.

LH

RH

(1) IGNITER PLUGS Zone inspection areas: 400. (a) Inspect the igniter plugs for condition and erosion (Ref. Pratt and Whitney PT6A-65B Engine Maintenance Manual P/N 3032842 or Chapter 74-00-00 of this manual). (2) ENGINE OIL FILTER Zone inspection areas: 400. (a) Inspect the oil filter for contamination (Ref. Pratt and Whitney PT6A-65B Engine Maintenance Manual P/N 3032842, Chapter 72-00-00, Table 601 Periodic Inspection). (3) MAGNETIC CHIP DETECTOR CLEANING Zone inspection area: 400. (a) Clean the magnetic chip detector (Ref. Pratt and Whitney PT6A-65B Engine Maintenance Manual P/N 3032842). (4) ENGINE VIBRATION ISOLATOR MOUNT Zone inspection area: 400. CAUTION If any isolator mounts have dislodged from their bracket positioning pins or have a gap or change in relative position, the airplane may have experienced a hard landing or encountered severe or extreme turbulent air. If so, perform additional inspections in Chapter 5-50-00. All mounts on an engine must be of the same manufacturer and carry the same part numbers. (a) Inspect for damage and attachment (Ref. Chapter 71 of the Model 1900 Airliner Series Component Maintenance Manual, P/N 114-590021-11). Replace or repair as required. The mount may be repaired by replacement of the rubber cushion.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Fifth 200-Hour-Interval Detailed Inspection (Continued) Complete Steps 1 thru 10.

Mechanic LH

Inspector

RH

(5) COMPRESSOR DRIVE QUILL SHAFT Zone inspection area: 621. (a) Check for wear and damage. (b) Lubricate the spline on the pulley end of the shaft (Ref. COMPRESSOR QUILL SHAFT LUBRICATION procedure Chapter 21-52-02). (6) COMPRESSOR DRIVE BELTS Zone inspection area: 420. (a) Check for cracks, shredding, fraying and wear. Check adjustment (Ref. COMPRESSOR BELT TENSION procedure Chapter 21-52-02). (7) ENGINE ACCESSORIES Zone inspection areas: 400. (a) Inspect all accessories, plumbing and associated equipment for damage, corrosion, attachment and leakage. (8) REFRIGERANT LINES AND SERVICE VALVES Zone inspection areas: 163, 173, 420 and 611. (a) Inspect refrigerant lines in the right engine cowling, nacelle, right wing and cabin for leaks, damage, cracks, corrosion and attachment. (9) REFRIGERANT COMPRESSOR Zone inspection area: 420. (a) Check for damage, attachment and oil leaks. (10) ENGINE FUEL PUMP, FILTERS AND SCREENS Zone inspection areas: 400, 730 and 740. (a) Inspect the filters and screens for contamination (Ref. Pratt and Whitney PT6A-65B Engine Maintenance Manual P/N 3032842).

F. General Service Items Fifth 200-Hour-Interval Detailed Inspection Complete Steps 1 thru 9. (1) PITOT AND STATIC SYSTEM Zone inspection areas: 221 (UC-1 and After), 246 and 247. (a) Drain system (Ref. Model 1900/1900C Flight Manual Supplements 114-590021-41 and 114-590021-87 as applicable). Close drains when completed.

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Mechanic

Inspector

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Fifth 200-Hour-Interval Detailed Inspection (Continued) Complete Steps 1 thru 9.

Mechanic

Inspector

(2) ELECTRONIC FLIGHT INSTRUMENT SYSTEM (EFIS) (IF INSTALLED) Zone inspection area: 248. (a) Verify operation of Electronic Attitude Director Indicator (EADI) and Electronic Horizontal Situation Indicator (EHSI) tube fans by listening for fan operation. EADI and EHSI tubes must be on (Ref. the applicable Model 1900/1900C Pilot's Operating Handbook/Airplane Flight Manual). (3) AIRPLANE LUBRICATION (a) Lubricate as necessary (Ref. Chapter 12-20-00). (4) EVAPORATOR FILTERS Zone inspection areas: 153 and 173. Floor panel inspection areas: 7 and 11 (UA-1 and After). 8 and 12 (UB-1 and After; UC-1 and After). (a) Replace the evaporator filters (Ref. Chapter 21-52-01). (5) VACUUM REGULATOR VALVE FILTER Zone inspection area: 812. (a) Clean or replace filter (Ref. Chapter 37-00-00). (6) ACCESS PANELS (a) Check panels removed during this inspection for fit, attachment, scratches, paint blistering and corrosion. (7) PLACARDS (a) Verify all placards are in place and legible (Ref. Pilots Operating Handbook, Airplane Flight Manual, Chapter 11-20-00 (UA-1 and After, UB-1 and After) and Chapter 11-21-00 (UC-1 and After)). (8) FUEL FILTERS AND SCREENS Zone inspection areas: 400, 410, 420, 511, 611, 730 and 740. (a) Inspect the filters and screens for microbiological growth. Clean filters and screens (Ref. SERVICING ENGINE FUEL FILTERS AND SCREENS procedure Chapter 28-20-02). (9) WINDSHIELDS (a) Inspect windshield weather seal for damage, debonding, cracks and wear.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

G. Operational Inspection Fifth 200-Hour-Interval Detailed Inspection Complete Steps 1 thru 5. (1) ENVIRONMENTAL TEST (a) Perform the BLEED AIR TEMPERATURE OPERATIONAL CHECK procedure (Ref. Chapter 21-10-00 or Chapter 21-11-00). (b) Perform the BLEED AIR PRESSURE OPERATIONAL CHECK procedure (Ref. Chapter 21-10-00 or Chapter 21-11-00). (2) PROPELLER DEICER (a) Perform the PROPELLER DEICER SYSTEM INSPECTIONS procedure (Ref. Chapter 30-60-00). (3) GROUND INSPECTION RUN Depending on the maintenance performed and components replaced, a GROUND PERFORMANCE CHECK PROCEDURE may be required in lieu of this Inspection Run. Refer to the applicable maintenance procedures. (a) Start engines and allow the oil temperature to increase into the operating range. (b) Run engines at a minimum of 80% N1 long enough for engine indicators to stabilize. (c) Shut down the engines and inspect for attachment and security of all components and for oil and fuel leaks. (4) INVERTER POWER (a) Perform the INVERTER POWER SELECT RELAY CHECK procedure (Ref. Chapter 24-20-00). (b) Perform the INVERTER BLOWER FAN OPERATIONAL CHECK procedure (Ref. Chapter 24-20-00).

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Mechanic

Inspector

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Fifth 200-Hour-Interval Detailed Inspection (Continued) Complete Steps 1 thru 5.

Mechanic

Inspector

(5) EXTERNAL POWER Zone inspection area: 522. (a) Check the external power relay for operation (rotate the voltmeter select switch to the EXT PWR position and check for external power voltage).

INSPECTION COMPLETED I certify that a Detailed Inspection was performed in accordance with the Continuous Inspection Program and that the airplane is approved for return to service. CREW CHIEF________________________________________________

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

CONTINUOUS INSPECTION WORKSHEET Inspection Type_________________________________Inspection Number__________________________ Model__________L Eng Time___________Cycles__________R Eng Time___________Cycles___________ A/C Time_______________Cycles________________A/C Serial_________________Date_______________

Item

Page 14 Aug 1/12

Description

05-20-06

Mech

Crew Chief

Q.C.

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Inflight Worksheet (Page 1 of 2) S/N________________N Number__________________Date___________________NO._________________ Flt. No.

Pilot

Copilot

TO

Time

LND

Time

No. LND

Flt. Time

Today’s Total Previous Total Total Fuel Type

Engine Oil Gallon

Left

Cruise Condition

Right

Data

Left

Right

O.A.T. PA IAS ITT Torque Prop. RPM N1 RPM Fuel Flow Fuel Pressure Oil Pressure Oil Temp.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Inflight Worksheet (Page 2 of 2) Discrepancy Worksheet S/N________________N Number__________________Date___________________NO._________________

A/C Hrs

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Date

Discrepancy

05-20-06

Corrective Action

Date

Mechanic Inspector

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Major Maintenance Worksheet Component

Date

Reason for Replacement

Replacement Part Number Serial Number

05-20-06

Next Overhaul A/C Hours or Cycles Date

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

TIME LIMITS/MAINTENANCE CHECKS SIXTH 200-HOUR-INTERVAL DETAILED INSPECTION DETAILED INSPECTION PROCEDURES

05-20-07 00

1. GENERAL A. Forms Required (1) Sixth 200-Hour Interval Detailed Inspection. (2) Model 1900/1900C Airliner Series Routine Inspection. (3) Continuous Inspection Worksheet. NOTE: A Routine Inspection must be conducted in conjunction with each Detailed Inspection to comply with Continuous Inspection Regulations.

B. Reference Material (1) Model 1900/1900C Airliner Maintenance Manual, P/N 114-590021-7. (2) Model 1900 Airliner Series Component Maintenance Manual, P/N 114-590021-11. (3) Model 1900 Airliner Series Structural Repair Manual, P/N 114-590021-9. (4) Model 1900 Airliner Series Corrosion Control Manual, P/N 114-590021-197. (5) Model 1900 Airliner Wiring Diagram Manual, P/N 114-590032-3 (UA-1 and After), 114-590021-13 (UB-1 and After), 114-590021-61 (UC-1 and After). (6) Model 1900 Airliner Pilot’s Operating Handbook/Airplane Flight Manual, P/N 114-590021-3. (7) Model 1900C Airliner Pilot’s Operating Handbook/Airplane Flight Manual, P/N 114-590021-57.

C. Inspection Procedures (1) Fill out the heading on each form in its entirety. (2) The mechanic checks each item on the inspection form and initials the form in the space provided. (3) List all discrepancies found during the inspection on the Continuous Inspection Worksheet. NOTE: Check all In-Flight Worksheets turned in since the last inspection for discrepancies that have not yet been worked off. (4) Each discrepancy is to be signed off by the mechanic, crew chief and a Quality Control Inspector when the discrepancy has been corrected. (5) In the spaces provided on the Major Maintenance Worksheet, the mechanic is to list all components which are removed from the airplane for overhaul or replacement, then to add the Part Number and Serial Number of the component which is installed. (6) The Quality Control Inspector will stamp off each item on the inspection form to complete the inspection.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (7) When the inspection has been completed, the crew chief will sign the “INSPECTION COMPLETED” block at the end of the inspection sheet.

2. DETAILED INSPECTION PROCEDURES Inspection Date___________________________________________ Airplane Serial___________________ Airframe______________________ Cycles____________ LH Engine Hrs_____________ Cycles____________ RH Engine Hrs______________ Cycles____________ LH Engine Power Module S/N_______________________________________________ RH Engine Power Module S/N_______________________________________________ LH Engine Gas Generator S/N_______________________________________________ RH Engine Gas Generator S/N_______________________________________________

NOTE: Corrosion detected while performing this detailed inspection may be treated in accordance with Chapter 20-09-00. To minimize the possibility of foreign object damage to engines, observe the following maintenance practices: •

Ensure all loose materials (rivets, screws, safety wire, etc.) are removed from engine cowling area after maintenance.

•

Maintain clean ramp and taxi areas.

•

Running at maximum power with the airplane stationary should be minimized and done only on a clean ramp.

•

Propeller reverse operation for backing the airplane should be avoided.

•

Avoid operation in dust and sand storms.

•

Do not operate engines in feather, except during external power starts and feather checks.

Prior to beginning this inspection, the following access panels must be removed: UA-1 and After: Wing Access - 12 and 15. Floor Access - 5, 6, 7, 8, 10, 12, 15, 16 and 18. UB-1 and After: Wing Access - 12 and 15. Floor Access - 5, 6, 7, 8, 9, 10, 12, 13, 15, 16 and 18. UC-1 and After: Wing Access - 12 and 15. Floor Access - 5, 6, 7, 8, 9, 10, 12, 13, 15, 16 and 18. For zone and access panel locations (Ref. Chapter 06).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

A. Aft Fuselage and Empennage Sixth 200-Hour-Interval Detailed Inspection Complete Steps 1 thru 12.

Mechanic

Inspector

(1) SKINS Zone inspection areas: 281, 282, 311, 312, 320, 331 and 340. (a) Inspect skins for dents, cracks, scratches, blistered paint, corrosion and loose or missing rivets. If damage is found, check adjacent structure. (2) STRUCTURE Zone inspection areas: 280, 281, 311, 312, 320, 331 and 340. (a) Check for cracks, corrosion, loose rivets, concealed damage and blistered paint. (b) Perform HORIZONTAL STABILIZER Attach Bolts Torque Check procedure (Ref. Chapter 55-10-00). (c) Between HSS 5.00 thru 99.197 both left and right hand sides using a borescope, inspect the area between the cove and the horizontal stabilizer rear spar. The borescope can be inserted into this area from the edges of the cove. Inspect the horizontal stabilizer rear spar, skins and fasteners for cracks, corrosion and loose and/or missing rivets, paying particular attention to elevator hinge attach points. (d) Check aft spar of vertical stabilizer for cracks, corrosion, loose and working fasteners, concealed damage and blistering paint. (Ref. Chapter 55-30-01) of the Model 1900 Airliner Series Corrosion Control Manual. (3) FLIGHT CONTROL COMPONENTS, CABLES AND PULLEYS Zone inspection areas: 181, 311, 312 and 331. Floor Panel inspection areas: 15. Aft fuselage inspection panels: 5, 6, 7, and 8. (a) Inspect the control system components (pushrods, turnbuckles, end fittings, castings, etc.) for bulges, splits, bends, cracks and corrosion. Replace any damaged component. (b) Check control cables, pulleys and associated equipment for condition, attachment, alignment, clearance, corrosion and correct direction of travel. Replace cables that have broken strands or evidence of corrosion. Perform CONTROL CABLES AND PULLEYS - MAINTENANCE PRACTICES procedure (Ref. Chapter 20-00-02).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Sixth 200-Hour-Interval Detailed Inspection (Continued) Complete Steps 1 thru 12. (c) Check the elevator and rudder autopilot or yaw damp bridle cable attaching clamp hardware for 55 ± 5 inch-pounds of torque and a minimum gap of 0.005 inch remaining between the clamp halves. If the minimum gap is below 0.005 inch, replace the worn parts as needed. (d) Check the elevator and rudder autopilot or yaw damp servo bridle cable tension, if installed. If the elevator or rudder servo cable tension is out of limits, the ELEVATOR SERVO CABLE TENSIONING and/or RUDDER SERVO RIGGING procedure must be performed (Ref. Chapter 22-10-00). Inspect and record elevator servo bridle cable and rudder servo bridle cable tensions: Temperature:________°F Elevator servo bridle cable tension: ______________ Rudder servo bridle cable tension: ______________ (e) Check cable tension (Ref. Chapter 27). If the elevator cable tension is not within limits, the ELEVATOR CONTROL SYSTEM RIGGING procedure must be performed (Ref. Chapter 27-30-02). Changing cable tension may affect other portions of the elevator system. Inspect and record elevator, elevator tab, rudder and rudder tab cable tensions: Temperature:________°F 3/16 in. Elevator Cable Tension: Up_______Down________ 1/16 in. Elevator Tab Cable Tension: _________ 3/16 in. Rudder Cable Tension: Left________Right________ 1/16 in. Rudder Tab Cable Tension: ________ (4) UNDERFLOOR AREAS Zone inspection areas: 181 and 182. (a) Inspect underfloor areas for foreign objects, structural damage, loose or missing rivets, cracks, corrosion and blistered paint. (5) FLIGHT CONTROLS Zone inspection areas: 331, 340, 351, 352, 361 and 362. Stabilizer access panels areas: 15, 18, 20, 22 and 24. (a) Inspect skin for cracks, dents, damage, paint blistering, corrosion and loose or missing rivets. (b) Check surfaces for attachment and freedom of movement. (c) Check optional elevator electric trim actuator and motor for attachment.

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05-20-07

Mechanic

Inspector

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Sixth 200-Hour-Interval Detailed Inspection (Continued) Complete Steps 1 thru 12.

Mechanic

Inspector

(d) Inspect elevator hinge brackets and their spar attach areas for cracks, corrosion, paint blistering, evidence of interference and security. Inspect the rivets attaching the hinge supports to the elevator. Perform the ELEVATOR INSPECTION procedure (Ref. Chapter 27-30-00). (e) Inspect rudder hinge brackets and their spar attach areas for cracks, corrosion, paint blistering, evidence of interference and security. (f) Remove tailcone. Using the rudder pedals in the cockpit, move the rudder from left to right. Check for smooth movement with no evidence of looseness, binding or warping. Check the bellcrank at the base of the rudder (three rudder horn attach bolts and pivot). There should be no evidence of fretting or looseness between the bellcrank and the rudder torque tube. Inspect the bolts that attach the top of the rudder torque tube to the bottom of the rudder for security. Check the torque tube to rudder attach point for evidence of cracking or corrosion. (g) Inspect and record Rudder Freeplay and Rudder Trim Tab Freeplay (Ref. Chapter 27-20-00). Rudder Freeplay: ________ (Limits: 0.12 Inch Maximum) Rudder Trim Tab Freeplay: ________ (Limits: 0.026 Inch Maximum) (h) Inspect and record Elevator Freeplay and Elevator Trim Tab Freeplay (Ref. Chapter 27-30-00). Elevator Freeplay: ________ (Limits: 0.12 Inch Maximum) Elevator Trim Tab Freeplay: ________ (Limits: 0.006 Inch Maximum)

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Sixth 200-Hour-Interval Detailed Inspection (Continued) Complete Steps 1 thru 12. (i) Perform the RUDDER TRIM TAB FUNCTIONAL CHECK procedure (Ref. Chapter 27-20-07). Turn the cockpit pedestal rudder trim control knob counter clockwise to the full nose left position and verify the rudder trim tab moves to the trailing edge right direction. Check that the system moves smoothly with no unusual noise or binding. Rudder Trim Tab Travel Full Nose Left: ____________ (Limit: 15° to 16.5° tab trailing edge right) Turn the cockpit pedestal rudder trim control knob clockwise to the full nose right position and verify the rudder trim tab moves to the trailing edge left direction. Check that the system moves smoothly with no unusual noise or binding. Rudder Trim Tab Travel Full Nose Right: ___________ (Limit: 15° to 16.5° tab trailing edge left) (j) Perform the RUDDER FUNCTIONAL CHECK procedure (Ref. Chapter 27-20-02). Check that the system moves smoothly with no unusual noise or binding. Push forward on the pilot’s left rudder pedal and make sure that the rudder moves to the trailing edge left direction. Rudder Travel Full Nose Left: _______________ (Limit: 25° +1°/ -0° rudder trailing edge left) Push forward on the pilot’s right rudder pedal and make sure that the rudder moves to the trailing edge right direction. Rudder Travel Full Nose Right: ______________ (Limit: 25° +1°/ -0° rudder trailing edge right)

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05-20-07

Mechanic

Inspector

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Sixth 200-Hour-Interval Detailed Inspection (Continued) Complete Steps 1 thru 12.

Mechanic

Inspector

(k) Perform the ELEVATOR TRIM TAB FUNCTIONAL CHECK procedure (Ref. Chapter 27-30-05). Install rig pin in aft elevator bellcrank to position the elevator in neutral. Rotate the elevator trim wheel on the cockpit pedestal to align the trailing edge of the elevator trim tabs with the trailing edge of the elevator. Elevator Trim Tab Wheel Neutral Check. The 0 mark on the trim position dial must be aligned with the trim indicator mark on the pedestal edgelighted panel: ___________________ (enter Yes or No) Rotate the cockpit pedestal elevator trim control wheel counter clockwise to the full nose down position and verify the elevator trim tab moves to the trailing edge up direction. Check that the system moves smoothly with no unusual noise or binding. Elevator Trim Tab Travel Full Nose Down: ___________ (Limit: 5° to 5.5° tab trailing edge up) Rotate the cockpit pedestal elevator trim control wheel clockwise to the full nose up position and verify the elevator trim tab moves to the trailing edge down direction. Check that the system moves smoothly with no unusual noise or binding. Elevator Trim Tab Travel Full Nose Up: _____________ (Limit: 15° to 16° tab trailing edge down) (l) Perform the ELEVATOR FUNCTIONAL CHECK procedure (Ref. Chapter 27-30-02). Perform this check after completing all work affecting the elevator flight control system. Check that the system moves smoothly with no unusual noise or binding. Move a cockpit control wheel aft and measure the elevator trailing edge up travel. Move the control wheel forward and measure the elevator trailing edge down travel. Elevator Up Travel: ________________ (Limit: 20° +1°/ -0°) Elevator Down Travel: ______________ (Limit: 14° +1°/ -0°) (6) VENTRAL FIN AND AFT FUSELAGE DRAINS Zone inspection area: 312. (a) Inspect the drain holes in the ventral fin of the aft fuselage for obstructions at the point of juncture with the aft pressure bulkhead. (b) Check bayonet-type drain for damage and obstruction. (7) PLUMBING Zone inspection areas: 311 and 312. (a) Inspect plumbing for leaks, cracks, damage and attachment.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Sixth 200-Hour-Interval Detailed Inspection (Continued) Complete Steps 1 thru 12.

Mechanic

Inspector

Mechanic

Inspector

(8) DEICER BOOTS Zone inspection areas: 331, 351 and 352. (a) Visually check deicer boots for cracks, gaps, tears, damage and attachment. (9) STATIC PORTS Zone inspection areas: 311 and 312. (a) Check and clean as necessary. (10) OUTFLOW VALVES Zone inspection area: 281and 282. (a) Check for operation, cleanliness and attachment. Clean valves (Ref. Chapter 21-30-01). (11) CONTROL CABLE SEALS Zone inspection area: 281 and 282. (a) Check for damage, security, cleanliness and lubrication. (12) AUTOPILOT (IF INSTALLED) Zone inspection areas: 311 and 312. Panel inspection areas: 7 and 8. (a) Check rudder and elevator autopilot servos for loose or worn bearings and mounting hardware. (b) Verify that the servo mounts are securely mounted to the airframe. (c) Visually inspect the capstan and cable for wear, contamination and proper spool-off. (d) With the autopilot disengaged, operate each control system through its entire range. Observe the servo mounts for any unusual noise, binding, backlash or other mechanical irregularities.

B. Cabin Section Sixth 200-Hour-Interval Detailed Inspection Complete Step 1. (1) EVAPORATOR BLOWER MOTOR Panel inspection areas: 153 and 173. (a) Inspect brushes for wear.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

C. Power Plant Sixth 200-Hour-Interval Detailed Inspection Complete Steps 1 thru 4.

Mechanic LH

Inspector

RH

(1) IGNITER PLUGS Zone inspection areas: 400. (a) Inspect and clean the igniter plugs (Ref. Pratt and Whitney PT6A-65B Engine Maintenance Manual P/N 3032842 or Chapter 74-00-00 of this manual). (2) ENGINE OIL FILTER Zone inspection areas: 400. (a) Inspect the oil filter for contamination (Ref. Pratt and Whitney PT6A-65B Engine Maintenance Manual P/N 3032842, Chapter 72-00-00, Table 601 Periodic Inspection). (3) STARTER-GENERATOR (a) Inspect brushes for indication of excessive wear and damage (determine wear by observing diagonal groove on brush). Replace as necessary. (b) Inspect inlet duct and blast cap for cracks, corrosion and obstruction. (4) FUEL NOZZLES Zone inspection areas: 400. (a) Inspect nozzles (Ref. Pratt and Whitney PT6A-65B Engine Maintenance Manual P/N 3032842). Borescope inspect at this time.

D. General Service Items Sixth 200-Hour-Interval Detailed Inspection Complete Steps 1 thru 7.

Mechanic

Inspector

(1) PITOT AND STATIC SYSTEM Zone inspection areas: 221 (UC-1 and After), 246 and 247. (a) Drain System (Ref. Flight Manual Supplements 114-590021-41 and 114-590021-87 as applicable). (2) ELECTRONIC FLIGHT INSTRUMENT SYSTEM (EFIS) Zone inspection area: 248. (a) Verify operation of the Electronic Attitude Directional Indicator (EADI) and Electronic Horizontal Situation Indicator (EHSI) tube fans (Ref. the appropriate Pilot's Operating Handbook/Airplane Flight Manual).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Sixth 200-Hour-Interval Detailed Inspection (Continued) Complete Steps 1 thru 7.

Mechanic

Inspector

Mechanic

Inspector

(3) AIRPLANE LUBRICATION (a) Lubricate as necessary (Ref. Chapter 12-20-00). (4) EVAPORATOR FILTER Zone inspection areas: 153 and 173. Floor panel inspection areas: 7 and 11 (UA-1 and After). 8 and 12 (UB-1 and After; UC-1 and After). (a) Perform the EVAPORATOR FILTER REPLACEMENT procedure (Ref. Chapter 21-52-01). (5) VACUUM REGULATOR VALVE FILTER Zone inspection area: 812. (a) Clean filter. Perform the VACUUM REGULATOR VALVE FILTER SERVICING procedure (Ref. Chapter 37-00-00). (6) PLACARDS (a) Verify that all placards are in place. Inspect placards for scratches, damage and legibility. (7) WINDSHIELDS (a) Perform WINDSHIELD ANTISTATIC COATING AND TAB INSPECTION procedures (Ref. Chapter 56-10-00). (b) Inspect windshield weather seal for debonding, cracks and wear.

E. Operational Inspection Sixth 200-Hour-Interval Detailed Inspection Complete Steps 1 thru 21. NOTE The following Operational Inspection procedures are to be applied during start and run of the engine: (1) ENGINE CONTROLS (a) Check for freedom of movement, full travel and friction lock. (2) STARTER-GENERATOR (a) Check for output of 28.25 ± 0.25 vdc using the test jack on the RH inboard subpanel (Ref. Chapter 24-30-01). (3) OIL (a) Check that pressure and temperature are within operating limits (Ref. the appropriate Pilot's Operating Handbook/ Airplane Flight Manual).

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05-20-07

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Sixth 200-Hour-Interval Detailed Inspection (Continued) Complete Steps 1 thru 21.

Mechanic

Inspector

(4) PROPELLER GOVERNOR (a) Check operation and feathering (Ref. the appropriate Pilot's Operating Handbook/Airplane Flight Manual). (5) PROPELLER SYNCHROPHASER (a) Check operation (Ref. Chapter 61-22-00). (6) PROPELLER DEICER (a) Perform the PROPELLER DEICER SYSTEM INSPECTION procedure (Ref. Chapter 30-60-00). (7) AUTOFEATHER RELAYS, DUMP SOLENOIDS AND PRESSURE SWITCHES (a) Check autofeathering (Ref. Chapter 61-21-00). (8) GROUND INSPECTION RUN Depending on the maintenance performed and components replaced, a GROUND PERFORMANCE CHECK PROCEDURE may be required in lieu of this Inspection Run. Refer to the applicable maintenance procedures. (a) Start engines and allow the oil temperature to increase into the operating range. (b) Run engines at a minimum of 80% N1 long enough for engine indicators to stabilize. (c) Shut down the engines and inspect for attachment and security of all components and for oil and fuel leaks. (9) FUEL BOOST PUMPS (a) Check operation (Ref. the appropriate Pilot's Operating Handbook/Airplane Flight Manual). (10) FUEL CROSS-TRANSFER VALVES (a) Check operation (Ref. the appropriate Pilot's Operating Handbook/Airplane Flight Manual). (11) FUEL FIREWALL SHUTOFF VALVES (a) Perform the FIREWALL FUEL SHUTOFF VALVE FUNCTIONAL CHECK procedure (Ref. Chapter 28-20-05) (UA-1 and After; UB-1 and After) or (Ref. Chapter 28-21-05) (UC-1 and After). (12) VACUUM SYSTEM (a) Perform the VACUUM REGULATOR ADJUSTMENT procedure (Ref. Chapter 37-00-00).

05-20-07

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Sixth 200-Hour-Interval Detailed Inspection (Continued) Complete Steps 1 thru 21. (13) ENVIRONMENTAL TEST (a) Perform the BLEED AIR TEMPERATURE OPERATIONAL CHECK procedure (Ref. Chapter 21-10-00 or Chapter 21-11-00). (b) Perform the BLEED AIR PRESSURE OPERATIONAL CHECK procedure (Ref. Chapter 21-10-00 or Chapter 21-11-00). (14) SURFACE DEICERS (a) Check for inflation and cycling. (b) Perform the SURFACE DEICER OPERATIONAL CHECK procedure (Ref. Chapter 30-10-00). (15) PRESSURIZATION SYSTEM (a) Check operation (Ref. Chapter 21-30-00). (16) PRESSURIZATION SYSTEM DRAIN VALVE (a) Open drain valves until all moisture is drained. (17) ENVIRONMENTAL VAPOR CYCLE SYSTEM AND AIR CYCLE MACHINE (a) Check operation when the switch is in the AUTO or MANUAL position. Ambient temperature must be above 50°F. (b) Check operation of all outlets and ease of operation of all controls. (18) CONDITION LEVER (a) Check for clean shut down at IDLE CUT-OFF. (19) INVERTER POWER (a) Perform the INVERTER POWER SELECT RELAY CHECK procedure (Ref. Chapter 24-20-00). (b) Perform the INVERTER BLOWER FAN OPERATIONAL CHECK procedure (Ref. Chapter 24-20-00). (20) STATIC DISCHARGER (a) Perform the STATIC DISCHARGER INSPECTION procedure (Ref. Chapter 23-60-00) on all tail (stabilon, if installed, tailet, if installed, elevator, rudder, and strake if installed) mounted static dischargers. (21) EXTERNAL POWER Zone inspection area: 522. (a) Check the external power relay for operation (rotate the voltmeter select switch to the EXT PWR position and check for external power voltage).

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05-20-07

Mechanic

Inspector

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

INSPECTION COMPLETED I certify that a Detailed Inspection was performed in accordance with the Continuous Inspection Program and that the airplane is approved for return to service. CREW CHIEF________________________________________________

05-20-07

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

CONTINUOUS INSPECTION WORKSHEET Inspection Type_________________________________Inspection Number__________________________ Model__________L Eng Time___________Cycles__________R Eng Time___________Cycles___________ A/C Time_______________Cycles________________A/C Serial_________________Date_______________

Item

Page 14 Nov 1/12

Description

05-20-07

Mech

Crew Chief

Q.C.

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Inflight Worksheet (Page 1 of 2) S/N________________N Number__________________Date___________________NO._________________ Flt. No.

Pilot

Copilot

TO

Time

LND

Time

No. LND

Flt. Time

Today’s Total Previous Total Total Fuel Type

Engine Oil Gallon

Left

Cruise Condition

Right

Data

Left

Right

O.A.T. PA IAS ITT Torque Prop. RPM N1 RPM Fuel Flow Fuel Pressure Oil Pressure Oil Temp.

05-20-07

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Inflight Worksheet (Page 2 of 2) Discrepancy Worksheet S/N________________N Number__________________Date___________________NO._________________

A/C Hrs

Page 16 Nov 1/12

Date

Discrepancy

05-20-07

Corrective Action

Date

Mechanic Inspector

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Major Maintenance Worksheet Component

Date

Reason for Replacement

Replacement Part Number Serial Number

05-20-07

Next Overhaul A/C Hours or Cycles Date

Page 17 Nov 1/12

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

TIME LIMITS/MAINTENANCE CHECKS UNSCHEDULED MAINTENANCE CHECKS INSPECTION PROCEDURES

05-50-00 00

1. OPERATION IN AREAS OF HIGH DUST CONTENT Item

Inspection Requirement

Inspection Interval

(1) Nose Landing Gear Shock Strut

Clean off and wipe dry exposed polished surfaces.

Routine.

(2) Instrument Air Filters

Replace instrument line supply filters at or before 150 hours under extremely dusty conditions.

As noted.

CAUTION Disconnect the autopilot barometric altitude sensor line before applying reverse air pressure to pitot and static lines to prevent damage to the barometric altitude sensor. (3) Pitot Static Lines

Check for obstruction by applying reverse air pressure (not to exceed 20 psi.) to the ends of the pitot and static lines disconnected from the instruments.

300 Hours or as requested.

(4) Environmental Air Filter

Inspect for obstruction of air flow. Replace if necessary. As required.

2. OPERATING FROM VERY SOFT OR UNUSUAL TERRAIN

Item (1) Tires

Inspection Interval

Inspection Requirement Visually check for cuts, wear, deterioration and inflation.

Routine.

Check shock strut inflation. Perform METHOD 1 of MAIN LANDING GEAR SHOCK ABSORBER SERVICING - INFLATION (Ref. Chapter 32-10-00). Service as necessary.

Routine.

(2) Main Landing Gear (a) Shock Struts

Thoroughly clean and inspect for leaks, damage and security. Clean exposed surface of shock strut piston with clean cloth moistened with hydraulic fluid (39, Table 1, Chapter 91-00-00).

(b) Wheels

Check fluid level. Perform MAIN LANDING GEAR SHOCK ABSORBER SERVICING (Ref. Chapter 32-10-00).

Every 150 hours.

Remove and clean. Inspect for abrasions, cracks and chipped rims, bearings for wear, corrosion, fretting and bluing. Check seals for distortion, deterioration, proper fit, security and obvious damage.

Every 150 hours.

05-50-00

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Item (c) Brake Units

(d) Wheel wells

Inspection Requirement

Inspection Interval

Check cylinders and associated lines for obvious damage and leaks.

Routine.

Check for evidence of overheating.

Every 150 hours.

Check discs for scoring, distortion, damaged plating and evidence of overheating.

Every 300 hours.

Clean foreign material (dirt, etc.) from wheel wells. Inspect supports between main and aft spars in upper wheel well and the lift leg attach bracket at the main spar for deformation, cracks, etc.

As required.

Check for obvious damage.

Routine.

Remove and clean. Inspect for abrasions, cracks and chipped rims, bearings for wear, corrosion, fretting and bluing. Check seals for distortion, deterioration, proper fit and security.

Every 150 hours.

Check shock strut inflation. Perform METHOD 1 of NOSE LANDING GEAR SHOCK ABSORBER SERVICING - INFLATION (Ref. Chapter 32-20-00). Service as necessary.

Routine.

(3) Nose Landing Gear (a) Wheel

(b) Shock Strut

Thoroughly clean and inspect for leaks, damage and security. Clean exposed surface of shock strut piston with clean cloth moistened with hydraulic fluid (39, Table 1, Chapter 91-00-00). Check fluid level. Perform NOSE LANDING GEAR SHOCK ABSORBER SERVICING (Ref. Chapter 32-20-00).

Every 150 hours.

(c) Fork Assembly

Check for cleanliness and obvious damage.

Routine.

(d) Nose Wheel Steering

Check for obvious damage, associated rods and connections for damage.

Every 150 hours.

(e) Actuator Linkage

Check for excessive play, safety and security.

(f) Actuator

Check actuator and support brackets for visible damage and condition. Inspect bracket for loose or missing rivets. Inspect cover and bottom assembly of actuator for cracks at mounting lug.

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05-50-00

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

3. INSPECTION AFTER HARD OR OVERWEIGHT LANDING NOTE: A hard landing is any landing made by an airplane with a vertical decent rate greater than 600 ft/min (feet per minute) when the airplanes gross weight is less than or equal to the Maximum Landing Weight (MLW). Closely related to a hard landing is an overweight landing, which is defined as landing the airplane with a vertical decent rate greater than 360 ft/min (feet per minute) when the airplanes gross weight is greater than MLW but less than Maximum Take-Off Weight (MTOW). An overweight landing is also any landing when the airplane gross weight is greater than MTOW. As it is difficult to accurately determine vertical descent velocity, the following inspections, checks and tests must be performed whenever a hard or overweight landing has been reported or suspected. It is not possible to define specific details of the inspection procedure to be performed after every incident due to the wide variations in weight, speed, nature, and direction of loads that can be encountered. It is therefore recommended that before starting the inspection the pilot is consulted for information regarding the landing conditions. Ascertain, for example: 1. Whether the landing was straight, drifting, wing low, nose or tail heavy. 2. If any noise indicative of structural damage was heard. 3. The weight of the airplane and the fuel. This inspection should be carried out after a hard landing and before the airplane is certified as ready for further flight. The inspections are conducted at two levels. The first level consists of determining if any external damage has occurred and looking for evidence of internal structural failure. The second level is concerned with a more detailed inspection of any damaged areas which were indicated in the findings of the first level inspection. If it is determined by the first level inspection that there is no damage to the airplane, it is not necessary to proceed to the second level inspection. WARNING: Even though wrinkles in the wing or fuselage skin surface may be slight enough to be considered as negligible, a close inspection of the internal supporting structure may reveal serious damage.

A. First Level Item

Inspection Requirement

Inspection Interval

(1) General Appearance

Determine that the airframe components (nacelles, wings, fuselage) are in their normal configuration.

After hard or overweight landing.

(2) Landing Gear

Inspect tires for pressure, excessive wear, splits in the tread, bottoming out or folding over the sidewalls.

After hard or overweight landing.

Check the wheels for flat spots or cracked castings. Check shock struts and attachment lugs for cracks. Check shock absorbers for fluid leakage. After removal from jacks, check main and nose shock absorbers for proper inflation. Inspect hydraulic brake lines for leaks. Inspect downlock, drag link and gear door retract linkage for damage.

05-50-00

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Item

Inspection Requirement

Inspection Interval

Inspect landing gear actuator attachment lugs. Inspect supports between main and aft spars in upper wheel well and the lift leg attach bracket at the main spar for deformation, cracks, etc. Inspect the main and nose landing gear actuators for external leakage. Inspect areas around landing gear attach points. Inspect the main landing gear drag brace support structure as follows: (a) Place the airplane on jacks. (b) Disconnect the upper drag brace from the airplane structure in both the LH and RH wheel wells. (c) Inspect the upper drag brace attach bolts (hollow “lube type” bolts) for cracks. (d) Using a flashlight and mirror, inspect all of the drag brace support structure for possible cracks, particularly at the lower radius of the U-channel where it attaches to the main spar. (e) If cracks are suspected but are not clearly defined, the suspect area should be fluorescent liquid penetrant inspected, using procedures as outlined in AC43.13-1B. (f) If cracks are found, contact Hawker Beechcraft Corporation Technical Support, Hawker Beechcraft Corporation, Wichita, KS. 67206, and report the findings for evaluation. (g) If no cracks are found, install the drag brace. Hawker Beechcraft Corporation recommends that airplanes having experienced severe, or hard landings or other abnormal landing incidents which may have placed undue stress on the landing gears, are to be inspected within the first 150 service hours after such hard landing and at each 600 service hours thereafter. Airplanes that have received repairs in this area, upon Hawker Beechcraft Corporation recommendations, are exempt from this inspection except in the event of a future hard or abnormal landing incident.

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05-50-00

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Item (3) Nacelles

Inspection Requirement

Inspection Interval

Inspect external skin surfaces for distortion, loose or missing rivets.

After hard or overweight landing.

Check cowling attachment fittings for alignment or damage. Inspect engine control cables for smooth operation. Check plumbing and wiring for security and attachment. Inspect engine support mounts for cracks or structural failure. Check tips of propellers for damage. Check propeller spinner and back plate for evidence of interference with cowling. Inspect wheel well structure for damage or cracks. Check area surrounding the landing gear attachment points for distortion or cracks. Inspect engine isolator mounts for mounting bracket pin engagement. If found dislodged, inspect for damage (Ref. Chapter 71 of the Model 1900 Airliner Series Component Maintenance Manual). Repair or replace the isolator mounts as required. (4) Wing Center Section

Inspect external skin surface (upper and lower) for cracks, abnormal wrinkles and loose or missing rivets.

After hard or overweight landing.

Inspect plumbing, wiring and actuators for damage and security of attachment. Check keel, front and rear spar on the lower side of fuselage for damage and alignment. (5) Outboard Wing Panels

Inspect external wing surface skin for cracks, abnormal wrinkles and loose or missing rivets.

After hard or overweight landing.

(6) Fuselage Nose Section

Check external skin surface for cracks, abnormal wrinkles and loose or missing rivets.

After hard or overweight landing.

Check wheel well structure and area surrounding gear attach point for damage. Inspect avionics, radar antenna, wiring and plumbing for security and attachment. (7) Fuselage Center Section Inspect external skin surface for cracks, abnormal wrinkles and loose or missing rivets.

After hard or overweight landing.

Inspect around cabin windows for structural cracks. (8) Fuselage Aft Section

Check external skin surface the entire length for cracks, After hard or abnormal wrinkles and loose or missing rivets. overweight landing. Inspect empennage and control surfaces for freedom of movement.

05-50-00

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

B. Second Level NOTE: Because shock loading may be transmitted along one structural member to another, carefully inspect the surrounding and supporting structure in any damaged area found in the first level inspection. Item

Inspection Requirement

(1) Landing Gear

Place the airplane on jacks and check shock strut for free up and down movement.

Inspection Interval After hard or overweight landing.

Remove the tires and inspect internally for cuts or broken areas. Disassemble and examine wheels for cracks or distortion (Ref. Chapter 32 in the Component Maintenance manual). Visually inspect axle with 10X power Magnifying glass. If suspect, fluorescent liquid penetrant or magnaflux inspect. Replace or inspect wheel bearings and perform lubrication. Remove and replace or magnaflux the landing gear attach bolts, check bolt holes for cracks or elongation. Remove and replace or magnaflux drag link bolts and supports. Cycle the landing gear up and down using the power pack. Use airplane hand pump to extend landing gear. Cycle the gear with the power pack through at least one complete cycle before removing the airplane from jacks. (2) Nacelles

If tips of propeller have been damaged, (Ref. the Pratt After hard or and Whitney PT6A-65B Engine Maintenance Manual overweight landing. for the engine inspection procedure for propeller strike). Inspect areas surrounding the engine mounts. Check the internal structure of the wheel well for cracks or damage. Test plumbing and wiring for proper operation.

(3) Wing Center Section

Remove floorboards and access plates and inspect the front and rear spar and keel structure for evidence of deformation or structural failure. Test plumbing, wiring, flaps, control cables, pulley mounts, and any other system found in this area for proper operation.

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05-50-00

After hard or overweight landing.

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Item (4) Outboard Wing Panels

Inspection Requirement

Inspection Interval

Test plumbing and wiring for proper operation. Inspect plumbing and wiring for security of attachment.

After hard or overweight landing.

Inspect fuel cells and lines for leakage and damage. Inspect internal structure and fuel cells through access panels. (5) Fuselage Nose Section

Remove baggage compartment floorboards and inspect the keel structure and supporting members for damage.

After hard or overweight landing.

Inspect wheel well structure and surrounding areas for signs of structural failure. Test avionics, radar antenna, plumbing and wiring for proper operation. (6) Fuselage Center and Aft Section

Examine stringers, frames and side walls for deformation or structural failure.

After hard or overweight landing.

Test plumbing and wiring for proper operation. Inspect heating and air conditioning ducts for damage. Examine the control cables and pulley mountings and check for clearance from structure at pass-through locations. Ensure a smooth operation. (7) Flight Controls

Perform a flight controls, full travel, sweep inspection of the aileron, elevator and rudder primary systems.

After hard or overweight landing.

Perform a flap operational check. REPAIR OF DAMAGE Due to the variety and degree of structural damage which may be involved, the best repair or replacement procedure must be based on the inspection findings of the individual airplane. If the hard landing inspection indicates that serious structural damage has occurred, contact Hawker Beechcraft Corporation Technical Support, Hawker Beechcraft Corporation, Wichita, KS, 67206, for assistance. LOG BOOK ENTRY Following a hard landing inspection, an entry covering the extent of inspection, the damage and the repair (if applicable) must be noted in the airplane permanent records.

05-50-00

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

4. INSPECTION AFTER ENCOUNTERING TURBULENT AIR NOTE: This inspection should be carried out after the airplane has been subjected to high G loading while flying through extreme or severe turbulent air and before the airplane is returned to service. The inspection is conducted on two levels. The first level consists of determining if any external damage has occurred and looking for evidence of internal structural failure. The second level is concerned with a more detailed inspection of damaged areas which were indicated in the findings of the first level inspection. If it is determined by the first inspection that there is no damage to the airplane, it is not necessary to proceed to the second level inspection. Extreme - Airplane is violently tossed about and is practically impossible to control. May cause structural damage. Severe - Airplane may be momentarily out of control. Occupants are thrown violently against the belts and back into the seat. Unsecured objects are tossed about. WARNING: Even though wrinkles in the wing or fuselage skin surface may be slight enough to be considered as negligible, a close inspection of the internal supporting structure may reveal serious damage.

A. First Level Item

Inspection Requirement

(1) General Appearance

Determine that the airframe components (nacelles, wings, fuselage and empennage) are in their normal configuration.

(2) Wing Center Section

Inspect the external skin surface (upper and lower) for cracks, wrinkles and loose or missing rivets.

Inspection Interval After Encountering Turbulent Air.

Inspect plumbing, wiring and actuators for damage and security of attachment. Check the keel and the front and rear spar on the lower side of the fuselage for damage and alignment. (3) Nacelles

Inspect external skin surfaces for wrinkles and loose or missing rivets. Check cowling attachment fittings for alignment or damage. Inspect engine support mounts for cracks, deformation or structural failure. Inspect engine control cables for smooth operation and check plumbing and wiring for security and attachment. Inspect structure in wheel well for damage or cracks. Inspect engine isolator mounts for mounting bracket pin engagement. If found dislodged, inspect for damage (Ref. Chapter 71 of the Model 1900 Airliner Series Component Maintenance Manual). Repair or replace the isolator mounts as required.

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05-50-00

After Encountering Turbulent Air.

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Item (4) Outboard Wing Panels

Inspection Requirement

Inspection Interval

Inspect the top and bottom wing surface for cracks, wrinkles and loose or missing rivets.

After Encountering Turbulent Air.

Inspect aileron, aileron tab and flaps for wrinkles or cracks. Inspect internal structure and fuel cells through access panel openings. Inspect plumbing and wiring for security of attachment. (5) Fuselage Nose Section

Check external skin surface for cracks, wrinkles and loose or missing rivets.

After Encountering Turbulent Air.

Inspect area forward of windshield for evidence of structural deformation or failure. Inspect avionics, antenna and components for security and attachment. (6) Fuselage Center Section Inspect the entire length of the external skin surface for cracks, stress wrinkles and loose or missing rivets.

After Encountering Turbulent Air.

(7) Fuselage Aft Section

After Encountering Turbulent Air.

Inspect the entire length of the external skin surface for cracks, stress wrinkles and loose or missing rivets. Check the empennage surfaces for damage and freedom of movement. Inspect for skin wrinkles at the juncture of the fuselage and empennage. Check controls for freedom of movement.

05-50-00

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

B. Second Level NOTE: Because G loading may be transmitted along one structural member to another, carefully inspect the surrounding and supporting structure in any damaged area found in the first level inspection. Item

Inspection Requirement

(1) Wing Center Section

Remove floorboards and access panels and inspect the front and rear spar and keel structure for evidence of deformation or structural failure.

Inspection Interval After Encountering Turbulent Air.

Operational test plumbing, wiring, flaps, control cables, pulley mounts and any other system found in this area. (2) Nacelles

Inspect areas surrounding the engine mounts. Inspect internal structure for cracks or damage.

After Encountering Turbulent Air.

Operational test plumbing and wiring. (3) Outboard Wing Panels

If there is evidence of damage to the fuel cells or fuel lines, remove the cells and inspect the fuel cell liners and liner support structure.

After Encountering Turbulent Air.

Operational test the plumbing and wiring, flap actuator, aileron and tab mounting. (4) Fuselage Nose Section

Remove the floorboards and inspect the keel structure and supporting members for damage.

After Encountering Turbulent Air.

Examine any fixed equipment for loose, broken or cracked mountings. Operational test the avionics, radar antenna, plumbing and wiring. (5) Fuselage Center and Aft Section

Examine stringers, frames and side walls for deformation or structural failure.

After Encountering Turbulent Air.

Examine heating and air-conditioning ducts for damage. Operational test plumbing and wiring. Examine the control cables, pulley mountings and cable clearance at areas the cables pass through the structure. Ensure a smooth, normal operation. (6) Empennage

Inspect elevator pushrods, torque tubes and bell cranks for damage. Inspect the attachment of the vertical stabilizer spars to the top of the fuselage for evidence of damage. Inspect skin surfaces for condition and loose or missing rivets.

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After Encountering Turbulent Air.

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Item

Inspection Requirement

Inspection Interval

Check structure for cracks, loose or missing rivets and/ After Encountering or concealed damage. Turbulent Air. Check rudder for freedom of movement and attachment. Check elevator for freedom of movement and attachment. Check trim tab actuators for smoothness of operation and attachment. Check the wiring of the electrical trim tab actuator for connection, security of attachment and condition. Check the electrical trim tab actuator for full travel and security of attachment. REPAIR OF DAMAGE Due to the variety and degree of structural damage which may be involved, the best repair or replacement procedure must be based on the inspection findings of the individual airplane. If the turbulent air inspection indicates that serious structural damage has occurred, contact Hawker Beechcraft Corporation Technical Support, Hawker Beechcraft Corporation, Wichita, KS, 67206, for assistance. LOG BOOK ENTRY Following a turbulent air inspection, an entry covering the extent of inspection, the damage and the repair (if applicable) must be noted in the airplane permanent records.

5. INSPECTION AFTER LIGHTNING STRIKE CAUTION: Following a confirmed lightning strike to a propeller, it must be removed, disassembled, and inspected. Refer to the Special Inspections section of Hartzell Propeller Inc.'s Standard Practices Manual, No. 202A or subsequent revision. This manual may be found in Chapter 61 of the Model 1900 Airliner Series Component Maintenance Manual. Item (1) Propeller

Inspection Requirement

Inspection Interval

Lightning strikes usually enter the metal erosion shield After Lightning Strike. directly. The charge travels through the erosion shield generally exiting at the butt end where it enters the next conductive element in the path. If a lightning strike is present, a darkened area and possible pitting, usually in proximity of the tip and at the most inboard end of the metal erosion shield, will be noticeable. Whenever the propeller has been struck by lightning, the propeller governors must be replaced or overhauled (Ref. Woodward Service Bulletin No. 33574B or subsequent).

(2) Engines

Inspect as instructed in the Pratt and Whitney PT6A-65B Engine Maintenance Manual.

After Lightning Strike.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Item

Inspection Requirement

Inspection Interval

(3) Fuselage, Empennage and Wing Surfaces

Carefully inspect the exterior of the airplane. Evidence of a strike will usually appear as a burned hole or as a series of burned holes in metallic surfaces. Plastic parts may be delaminated and/or deformed due to high internal pressures. Normally two or more points will be found, the entry and the exit points. Antennas are frequently an entry point of lighting and should be carefully inspected for evidence of arcing, sooting or pitting.

After Lightning Strike.

From the point of entry, the strike usually spreads aft in a series of small holes or burn marks. After the points of entry and exit are found, the structure between these points should be carefully inspected. Attention should be given to hinges and hinge pins for possible pitting. Cables, pulleys, bearings, bolts and all bonding jumpers in the area should be inspected for possible damage. Antennas, electrical and electronic equipment should be visually checked for damage and functionally checked for operation. If the strike was near the fuel vent, all plumbing should be carefully inspected for damage. Steel components may exhibit magnetism and require degaussing so as not to affect compass systems.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

6. ENGINE INSPECTION AFTER SUDDEN STOPPAGE CAUTION: After a sudden stoppage the engine’s governor, autofeathering valve or pump must not be returned to service, they must be overhauled or replaced. Item

Inspection Requirement

Inspection Interval

(1) Engine

Inspect as instructed in the Pratt and Whitney PT6A-65B Engine Maintenance Manual.

(2) Propeller Governor

The propeller governor should be overhauled or replaced as instructed in Woodward Maintenance Manual P/N 33048F or subsequent (Ref. Woodward Service Bulletin No. 33574B or subsequent).

(3) Propeller

Overhaul or replace propeller.

After sudden engine stoppage.

7. INSPECTION AFTER HEAVY EQUIPMENT CARGO OPERATION Item (1) Cargo Door Lower Attachment Lug Inspection

Inspection Requirement

Inspection Interval

The cargo door lower attachment lugs should have a visual inspection after each heavy equipment cargo operation. Inspect for cracks, breakage, proper alignment and security of attachment.

After Heavy Equipment Cargo Operation.

05-50-00

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

8. INSPECTION AFTER DEPLOYMENT OF LANDING GEAR ABOVE CRITICAL SPEED CONDITION NOTE: This inspection should be carried out after the landing gear doors have been deployed at an airspeed above the critical deployment speed and before the airplane is returned to service. The inspection will be conducted on two levels. The first level consists of determining if any external damage has occurred and looking for evidence of internal structural failure. The second level is concerned with a more detailed inspection of damaged areas which were indicated in the findings of the first level inspection. If it is determined by the first level of inspection that there is no damage to the landing gear door and surrounding structure, it is not necessary to proceed to the second level inspection. WARNING: Even though wrinkles in the skin surfaces may be considered slight enough to be considered as negligible, a close inspection of the internal supporting structure may reveal serious damage. Determine that the surfaces are in their normal configuration when stowed or deployed.

A. First Level Item

Inspection Requirement

Inspection Interval

(1) General Appearance

Determine that the airframe components (landing gear and flaps) are in their normal configuration.

(2) Landing Gear Door

Inspect the skin panels for wrinkles, cracks and bond separation.

After landing gear doors have been opened above critical speed condition.

Inspect for loose or missing rivets, bolts, and bearings. Inspect hinges, linkages, fittings and support structure for damage, alignment and security attachments. Check required clearances and overcenter requirements. Check gear door free play.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

B. Second Level NOTE: Since loads may be transmitted along one structural member to another, carefully inspect the adjacent members to any damaged element found in the first level inspection. Item (1) Inspection of Damaged Areas

Inspection Requirement

Inspection Interval

Remove the skin panels and conduct inspection by employing nondestructive test methods (acoustic, x-ray, and/or fluorescent liquid penetrant inspection). Inspect fastener holes for cracking. Repair or replace as determined by the extent of the damage.

After landing gear doors have been opened above critical speed condition.

In case of missing or loose fasteners, disassemble and inspect holes and fittings for distortion of holes and cracking. Repair or replace accordingly as determined by the extent of damage. Hinges, linkages, fittings, bearings and support structure which exhibit damage, alignment, and/or security attachment will be disassembled if possible and inspected. Repair or replace accordingly as determined by extent of damage. Inspect gear door free play after all repairs, checks and alignments have been made. Check operation of door from stowed through deployment. Repair or replace accordingly as determined by extent of damage. REPAIR OF DAMAGE Due to the variety and degree of structural damage which may be involved, the best repair or replacement procedure must be based on the inspection findings of the individual airplane. If the preceding inspection indicates that serious structural damage has occurred, contact Hawker Beechcraft Corporation Technical Support, Hawker Beechcraft Corporation, Wichita, KS, 67206, for assistance. LOG BOOK ENTRY Following the inspection, an entry covering the extent of inspection, the damage and the repair (if applicable) must be noted in the airplane permanent records.

05-50-00

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

9. INSPECTION AFTER DEPLOYMENT OF FLAPS ABOVE CRITICAL SPEED CONDITION NOTE: This inspection should be carried out after the flaps have been deployed at an airspeed above the critical deployment speed and before the airplane is returned to service. The inspection will be conducted on two levels. The first level consists of determining if any external damage has occurred and looking for evidence of internal structural failure. The second level is concerned with a more detailed inspection of damaged areas which were indicated in the findings of the first level inspection. If it is determined by the first level of inspection that there is no damage to the flaps and surrounding structure, it is not necessary to proceed to the second level inspection. WARNING: Even though wrinkles in the skin surfaces may be considered slight enough to be considered as negligible, a close inspection of the internal supporting structure may reveal serious damage. Determine that the surfaces are in their normal configuration.

A. First Level Item

Inspection Requirement

(1) General Appearance

Determine that the airframe components (landing gear and flaps) are in their normal configuration.

(2) Flaps

Inspect the skin panels for wrinkles, cracks, and bond separation. Inspect for loose or missing rivets, bolts, and bearings. Inspect tracks, screws, linkages, fittings, flap brackets, actuators, wing brackets and support structure for damage, alignment and security attachment. Check phase alignment between adjacent flaps. Check correlation between flap position and cockpit indicator.

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Inspection Interval After the flaps have been actuated above critical speed condition.

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

B. Second Level NOTE: Since loads maybe transmitted along one structural member to another, carefully inspect the members adjacent to any damaged element found in the first level inspection. Item (1) Inspection of Damaged Areas

Inspection Requirement

Inspection Interval

Remove the skin panels and conduct inspection by employing nondestructive test methods (acoustic, x-ray, and/or fluorescent liquid penetrant inspection). Inspect fastener holes for cracking. Repair or replace as determined by the extent of the damage.

After the flaps have been actuated above critical speed condition.

In case of missing or loose fasteners, disassemble and inspect holes and fittings for distortion of holes and cracking. Repair or replace accordingly as determined by the extent of damage. Hinges, linkages, fittings, bearings and support structure which exhibit damage, alignment, and/or security attachment must be disassembled if possible and inspected. Repair or replace accordingly as determined by extent of damage. Inspect hinges, linkages, fittings and support structure for damage, alignment, and security attachments. Check required clearances and over-center requirements. REPAIR OF DAMAGE Due to the variety and degree of structural damage which may be involved, the best repair or replacement procedure must be based on the inspection findings of the individual airplane. If the preceding inspection indicates that serious structural damage has occurred, contact Hawker Beechcraft Corporation Technical Support, Hawker Beechcraft Corporation, Wichita, KS, 67206, for assistance. LOG BOOK ENTRY Following the inspection, an entry covering the extent of inspection, the damage and the repair (if applicable) must be noted in the airplane permanent records.

05-50-00

Page 17 May 1/12

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

10. INSPECTION IN THE EVENT OF A BENT NOSE STEERING STOP Item

Inspection Requirement

(1) Nose Landing Gear

Inspect the lower portion of the upper gear brace assembly for evidence of hydraulic leaks. If leaks are found remove and repair or replace the nose landing gear (Ref. Chapter 32-20-00).

(2) Steering Stop Support Lugs

If no fluid leakage is found, perform the MECHANICAL STEERING NOSE GEAR STOP REMOVAL procedure (Ref. Chapter 32-50-00) and inspect both stop support lugs for cracks. If no cracks are detected visually, perform the FLUORESCENT LIQUID PENETRANT INSPECTION PROCEDURES or the EDDY CURRENT GENERAL PROCEDURE FOR SURFACE INSPECTIONS procedure in the Model 1900/1900C Airliner Structural Inspection Manual (Ref. Chapter 20-00-00). If cracks are found remove and repair or replace the nose landing gear (Ref. Chapter 32-20-00). If no cracks are found perform the MECHANICAL STEERING NOSE GEAR STOP INSTALLATION procedure (Ref. Chapter 32-50-00).

Page 18 May 1/12

05-50-00

Inspection Interval In the event of a bent steering stop.

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

11. INSPECTION AFTER FLIGHT IN AIRSPACE WITH A LOW CONTAMINATION OF VOLCANIC ASH NOTE: Refer to Hawker Beechcraft Corporation Model 1900 Airliner Series Model Communique No. 94. Hawker Beechcraft Corporation does not recommend airplanes operate in areas of volcanic emissions. However, if operated in low ash concentrations, limit the amount of time spent in the environment and perform a visual inspection to include (but not limited to) the following areas for signs of abrasion or ash contamination during pre/post-flight walk around (reference appropriate Flight Manual): Item

Inspection Requirement

Inspection Interval

(1) Radome

Visual inspect area for signs of abrasion or ash contamination.

(2) Pitot probes and static ports

Visual inspect area for signs of abrasion or ash contamination.

(3) Windshield

Visual inspect area for signs of abrasion or ash contamination.

(4) Leading edges

Visual inspect area for signs of abrasion or ash contamination.

(5) Navigation and landing lights

Visual inspect area for signs of abrasion or ash contamination.

(6) Air intakes

Visual inspect area for signs of abrasion or ash contamination.

(7) Propellers

Visual inspect area for signs of abrasion or ash contamination.

(8) Nacelles

Visual inspect area for signs of abrasion or ash contamination.

(9) Horizontal Stabilizer

Visual inspect area for signs of abrasion or ash contamination.

After Flight in Airspace with a Low Contamination of Volcanic Ash.

(10) All protruding structures/ Visual inspect area for signs of abrasion or ash components (TAT contamination. probes, antennas, etc.) (11) Any other areas of the airplane subject to impact abrasion

Visual inspect area for signs of abrasion or ash contamination.

NOTE: If you have additional questions or need further support, please contact your Regional Field Representative or Hawker Beechcraft Corporation Technical Support at 1-800- 429-5372 or 316-676-3140. Engine manufacturers have issued guidance for operations in areas where volcanic ash may be present. Hawker Beechcraft Corporation recommends contacting applicable engine manufacturers for specific instructions/recommendations.

05-50-00

Page 19 May 1/12

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL If the airplane is located in an area where volcanic ash may settle, hangar the airplane when possible. If the airplane must be parked outside, ensure that all protective covers are installed. It should be noted that volcanic ash is acidic and can cause corrosion damage unless properly removed. If volcanic ash has collected on airplane surfaces or been deposited in structural cavities or low points, Hawker Beechcraft Corporation recommends removal of loose ash using low pressure compressed air or vacuum equipment followed by a thorough clear, fresh water rinse of the structure making sure to remove any pooling or standing water. Personal Protective Equipment (PPE) should be worn, when applicable, when removing volcanic ash. NOTE: If contamination is noted in avionics compartments, it is recommended that owners/operators contact the applicable avionics manufacturer for guidance. Included are references to current EUROCONTROL documentation and engine manufacturers documentation released to date. Additional information/documentation may become available. Hawker Beechcraft Corporation assumes no responsibility for supplying future supplier documentation or revisions.

A. Eurocontrol Reference Links (1) http://www.metoffice.gov.uk/aviation/vaac/vaacuk_vag.html (2) https://www.cfmu.eurocontrol.int/PUBPORTAL/gateway/spec/index.html (3) http://www.metoffice.gov.uk/corporate/pressoffice/2010/volcano/ashconcentration/index.html

B. Engine Manufacturer References (1) Pratt & Whitney Canada, Special Inspection 25-2010.

C. Avionics Manufacturer References (1) Applicable Honeywell Aerospace, Letter(s) dated April 2010. Honeywell Letters: https://portal.honeywell.com/wps/portal/aero/products/apm (2) Applicable Rockwell Collins Service Information.

Page 20 May 1/12

05-50-00

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

CHAPTER 06 - DIMEMSIONS AND AREAS TABLE OF CONTENTS SUBJECT

PAGE

GENERAL INFORMATION 06-00-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1

AIRPLANE DIMENSIONS 06-10-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1

AIRPLANE STATIONS 06-30-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1

AIRPLANE ZONES 06-40-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1

FUSELAGE ACCESS PANELS 06-50-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Floor Access Panels (UA-1 and After) Figure 2, Sheet 1 of 3 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5 Floor Access Panels (UA-1 and After) Figure 2, Sheet 2 of 3 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7 Floor Access Panels (UA-1 and After) Figure 2 Sheet 3 of 3 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9 Floor Access Panels (UB-1 and After; UC-1 and After) Figure 3, Sheet 1 of 3 . . . . . . . . . . . . . . . . . . . 11 Floor Access Panels (UB-1 and After; UC-1 and After) Figure 3, Sheet 2 of 3 . . . . . . . . . . . . . . . . . . . 13 Floor Access Panels (UB-1 and After; UC-1 and After) Figure 3, Sheet 3 of 3 . . . . . . . . . . . . . . . . . . . 15 Stabilizer Access Panels (Figure 4, Sheet 1 of 2) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 17 Stabilizer Access Panels (Figure 4, Sheet 2 of 2) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 19 Wing Access Panels (UA-1 and After; UB-1 and After) (Figure 5, Sheet 1 of 4) . . . . . . . . . . . . . . . . . . 21 Wing Access Panels (UA-1 and After; UB-1 and After) (Figure 5, Sheet 2 of 4) . . . . . . . . . . . . . . . . . . 23 Wing Access Panels (UA-1 and After; UB-1 and After) (Figure 5, Sheet 3 of 4) . . . . . . . . . . . . . . . . . . 25 Wing Access Panels (UA-1 and After; UB-1 and After) (Figure 5, Sheet 4 of 4) . . . . . . . . . . . . . . . . . . 27 Wing Access Panels (UC-1 and After) Figure 6 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 29

06-CONTENTS

Page 1 Nov 1/09

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

List of Effective Pages CH-SE-SU

PAGE

DATE

06-LOEP

1

Nov 1/13

06-CONTENTS

1

Nov 1/09

06-00-00

1

Nov 1/09

06-10-00

1 thru 3

Nov 1/09

06-30-00

1 thru 5

Nov 1/09

06-40-00

1 thru 3

Nov 1/13

06-50-00

1 thru 29

May 1/10

06-LOEP

Page 1 Nov 1/13

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

DIMENSIONS AND AREAS GENERAL INFORMATION DESCRIPTION AND OPERATION

06-00-00 00

1. GENERAL This chapter provides dimensions, stations and zone illustrations to aid maintenance personnel in locating various components on the airplane as referenced elsewhere in this manual. This Chapter is broken down in several sub-system/sections. The following is a list of the information contained in this Chapter. 6-10-00 - Airplane Dimensions 6-30-00 - Station Locations 6-40-00 - Zone Diagrams 6-50-00 - Fuselage Access Panels

06-00-00

Page 1 Nov 1/09

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

DIMENSIONS AND AREAS AIRPLANE DIMENSIONS DESCRIPTION AND OPERATION

06-10-00 00

1. GENERAL For the dimensions of the Model 1900 airplane (Ref. Figure 1). For the dimensions of the Model 1900C airplane (Ref. Figure 2).

06-10-00

Page 1 Nov 1/09

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 1 Airplane Dimensions (UA-1 and After)

Page 2 Nov 1/09

06-10-00

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 2 Airplane Dimensions (UB-1 and After; UC-1 and After)

06-10-00

Page 3 Nov 1/09

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

DIMENSIONS AND AREAS AIRPLANE STATIONS DESCRIPTION AND OPERATION

06-30-00 00

1. GENERAL For the station locations of the Model 1900 airplane (Ref. Figure 1). For the station locations of the Model 1900C airplane (Ref. Figure 2). For the vertical stabilizer station locations of the Model 1900 and 1900C airplane (Ref. Figure 3).

06-30-00

Page 1 Nov 1/09

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 1 (Sheet 1 of 2) Station Locations (UA-1 and After)

Page 2 Nov 1/09

06-30-00

FWD PRESSURE BULKHEAD

FUSELAGE

AFT PRESSURE BULKHEAD

06-30-00

635.31 651.61

293.68

605.98

326.00

588.10

280.50 290.50

544.25 557.50 570.107

468.25 482.75 498.25 509.50 523.50

363.25 378.25 393.25 408.25 423.25 438.25 451.00

WING

640.86

531.00

456.00

430.75

400.75

370.75

327.63 340.93

242.10

310.75 318.25 333.25 348.25

280.75

273.25 288.00 303.25

250.75

220.75

198.25

177.35

150.60

107.00 116.00 134.00

47.50 57.50 70.75 84.00

261.22

213.25 228.25 243.25 258.25

183.25

163.975

143.00

94.00

36.69

0.00 14.20 30.00

166.735

209.016

46.75 37.00 25.00

55.00

69.00 61.50

83.00

140.00 135.11 132.48 124.616

159.73 153.73 145.735

194.84 180.78 173.73

213.90

223.485

236.92

249.610

263.29

276.016

306.29 298.735 291.73

316.915

328.28

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

5.00 20.00 C L 35.00 50.00 66.00 82.00 98.197 110.634

NACELLE HORIZONTAL STABILIZER

FS 664.76 FS 641.89 CSS 69.18 CSS 54.03 CSS 39.29 CSS 24.97 CSS 11.50 WL 125.00 102.56 100.00

UC06B 062127AA.AI

Figure 1 (Sheet 2 of 2) Station Locations (UB-1 and After)

Page 3 Nov 1/09

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 2 Station Locations (UC-1 and After)

Page 4 Nov 1/09

06-30-00

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

A 10.5% CHORD 20% CHORD FRONT SPAR 29.75% CHORD 39.5% CHORD 49.25% CHORD 58.75% CHORD REAR SPAR V.S.S. 91.10 WL 216.10

CSS 69.184

CSS 54.026 REAR SPAR VSS 84.10 RUDDER HINGE BEARING (3 PLACES)

CSS 39.289 RUDDER HINGE LINE CSS 24.973

RUDDER CANTED STATION 46.5

FRONT SPAR CSS 11.500 CSS 00.000

RS 46.5 TRIM TAB HINGE LINE

VSS 11.500 WL 125.00 VSS 0.000

RUDDER CANTED STA 0.000

RUDDER HINGE LINE INT. WITH WL 125.00 67.08 INCHES

RS 0.000

94.217 INCHES FS 551.90 DETAIL

A

FS 618.980

FS 646.12

UC06B 053213AA.AI

Figure 3 Vertical Stabilizer Stations Diagram

06-30-00

Page 5 Nov 1/09

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

DIMENSIONS AND AREAS AIRPLANE ZONES DESCRIPTION AND OPERATION

06-40-00 00

1. GENERAL For the zone diagrams of the Model 1900 and 1900C airplanes (Ref. Figure 1).

06-40-00

Page 1 Nov 1/13

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 1 (Sheet 1 of 2) Zone Diagrams

Page 2 Nov 1/13

06-40-00

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 1 (Sheet 2 of 2) Zone Diagrams

06-40-00

Page 3 Nov 1/13

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

DIMENSIONS AND AREAS FUSELAGE ACCESS PANELS DESCRIPTION AND OPERATION

06-50-00 00

1. GENERAL For the locations of the fuselage access panels on the Model 1900 and 1900C airplanes (Ref. Figure 1).

06-50-00

Page 1 May 1/10

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

18

1 TOP SIDE

5 16

14

9

15

17 7

11

2

INDEX 1.

LEFT SIDE

DESCRIPTION

INDEX 10. 11. 12. 13. 14. 15. 16. 17.

NOSE LANDING GEAR RETRACT CYLINDER OXYGEN SERVICE CONTROL CABLE AIR CONDITIONING SERVICE LH FLIGHT CONTROL CABLES RH FLIGHT CONTROL CABLES LH FLIGHT CONTROL CABLES RH FLIGHT CONTROL CABLES & ELT LH UPPER PITOT AND STATIC

2. 3. 4. 5. 6. 7. 8. 9.

18.

DESCRIPTION RH UPPER PITOT AND STATIC LH LOWER PITOT AND STATIC RH LOWER PITOT AND STATIC NOSE AVIONICS COMPARTMENT RADAR ANTENNA NOSE BAGGAGE COMPARTMENT RUDDER HORN AND PRIMARY STOP BOLTS RUDDER PRIMARY STOP BOLTS, RUDDER PUSH-PULL ROD, RUDDER HORN TAILCONE ACCESS

6 10

13

8 RIGHT SIDE

12

4

BOTTOM SIDE AS VIEWED FROM BELOW

Figure 1 Fuselage Access Panels

Page 2 May 1/10

06-50-00

3

UC27B 042935AB.AI

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

This Page Intentionally Left Blank

06-50-00

Page 3 May 1/10

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

1900

18

1

1

23

2

2

21

19

4

5 20

22

F.S. 145.00

3

UA06B 035150AB.AI

Figure 2 (Sheet 1 of 3) Floor Access Panels (UA-1 and After)

Page 4 May 1/10

06-50-00

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

A. Floor Access Panels (UA-1 and After) Figure 2, Sheet 1 of 3 1.

Nose Gear Steering Potentiometer, Brake Plumbing.

2.

Rudder Bellcrank.

3.

Hydraulic Lines.

4.

Control Cables, Hydraulic Plumbing.

5.

Power Steering Amplifier Control Box.

18.

Nose Landing Gear Aft Steering Link Rod End.

19.

Air-Conditioning Ducts.

20.

Hydraulic Tubing.

21.

Flight Control Cables and Air-Conditioning Ducts.

22.

Elevator Forward Bellcrank Rig Pin Hole and Emergency Landing Gear Extension Pump.

23.

Elevator Trim Tab Cables.

06-50-00

Page 5 May 1/10

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

6

A

A

7 B B

C

C 8 D

D 9

E

E 10 16

F

F

17

11 G

G 12

H

H 13

I

I 14

J

J 15

24 25 1900

UA06B 035149AB.AI

Figure 2 (Sheet 2 of 3) Floor Access Panels (UA-1 and After)

Page 6 May 1/10

06-50-00

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

B. Floor Access Panels (UA-1 and After) Figure 2, Sheet 2 of 3 6.

Parking Brake Valve, Electroluminescent Power Supply Terminal Board, Thermocouple Resistor, Heatsink Diode Assembly, Subpanel and Fuel Control Panel Feeder and Feeder Base, Circuit Board Feeders, Avionics Feeder, Prop Synchrophaser Control Box, Circuit Breaker Panel Feeder Base, Recording Accelerometer, Stall Lift Computer, Electroluminescent Power Supply, No. 1 Avionics Bus Relay, No. 2 Avionics Bus Relay, No. 3 Avionics Relay.

7.

Annunciator P.C. Board Box Assembly, Aural Warning Amplifier, Cabin Altitude Warning Switch, Air-Conditioner Forward Blower Motor, Air-Conditioner Forward Evaporator, High and Low Speed Blower Motor Relays, Deice Boot Valve, Hydraulic Plumbing, P.C. Board Relay Box Assembly, Forward Blower Motor Circuit Breaker, Forward DME Antenna Connector.

8.

Pneumatic Pressure Regulator, Air Cycle Ejector and Ejector Muffler, Bleed Air Fail Warning Pressure Switches, Venturi Suction for Instrument Air, Air-Conditioner Receiver-Dryer, Transponder Antenna Connector.

9.

Aileron Servo for Autopilot, Flap Motor, Flap Motor Relays, LH and RH Generator Control Panel, LH and RH Overvolt Test Switch, Bleed Air Overtemp Module and Connector, Fire Extinguisher Module and Connector, Fire Extinguisher Monitor Module and Connector, Time Delay Module and Connector, Prop Deice Timer, Air-Conditioner command Relay, Spar Cover Light Filter, VHF Antenna Connector.

10.

Air-Conditioner Pressure Switch, Yaw Damper, Aileron Control Cables.

11.

Air-Conditioner Aft Evaporator, Air-Conditioner Aft Blower Motor, High and Low Speed Blower Motor Relay, Beacon Light, Expansion Valve.

12.

Air-Conditioning Ducts.

13.

Air-Conditioning Ducts.

14.

Air-Conditioning Ducts.

15.

Control Cables, Air-Conditioning Ducts, Aft DME Antenna Connector.

16.

LH Side Floorboard Panels.

A.

Control Cables, Air-Conditioning Ducts.

B.

Control Cables, Air-Conditioning Ducts.

C.

Control Cables, Air-Conditioning Ducts.

D.

Control Cables, Air-Conditioning Ducts.

E.

Control Cables, Air-Conditioning Ducts, Aileron Quadrant.

F.

Control Cables, Air-Conditioning Ducts.

G.

Control Cables, Air-Conditioning Ducts.

H.

Control Cables, Air-Conditioning Ducts.

I.

Control Cables and Turnbuckles, Air-Conditioning Ducts.

J.

Control Cables, Air-Conditioning Ducts.

06-50-00

Page 7 May 1/10

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

6

A

A

7 B B

C

C 8 D

D 9

E

E 10 16

F

F

17

11 G

G 12

H

H 13

I

I 14

J

J 15

24 25 1900

UA06B 035149AB.AI

Figure 2 (Sheet 3 of 3) Floor Access Panels (UA-1 and After)

Page 8 May 1/10

06-50-00

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

C. Floor Access Panels (UA-1 and After) Figure 2 Sheet 3 of 3 17.

RH Side Floorboard Panels.

A.

Air-Conditioning Ducts, Plumbing.

B.

Air-Conditioning Ducts, Plumbing.

C.

Air-Conditioning Ducts, Plumbing.

D.

Air-Conditioning Ducts, Plumbing.

E.

Air-Conditioning Ducts, Plumbing.

F.

Air-Conditioning Ducts, Plumbing.

G.

Air-Conditioning Ducts, Plumbing.

H.

Air-Conditioning Ducts, Plumbing.

I.

Air-Conditioning Ducts, Plumbing.

J.

Air-Conditioning Ducts, Plumbing.

24.

Control Cables.

25.

Control Cables.

06-50-00

Page 9 May 1/10

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

1900C

18

1

1

23

2

2

21

19

4

5 20

22

F.S. 145.00

3

UC06B 035152AB.AI

Figure 3 (Sheet 1 of 3) Floor Access Panels (UB-1 and After; UC-1 and After)

Page 10 May 1/10

06-50-00

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

D. Floor Access Panels (UB-1 and After; UC-1 and After) Figure 3, Sheet 1 of 3 1.

Nose Gear Steering Potentiometer, Brake Plumbing.

2.

Rudder Bellcrank.

3.

Hydraulic Lines.

4.

Control Cables, Hydraulic Plumbing.

5.

Power Steering Amplifier Control Box.

18

Nose Landing Gear Aft Steering Link Rod End.

19.

Elevator Cables and Air-Conditioning Ducts.

20.

Hydraulic Tubing.

21.

Flight Control Cables and Air-Conditioning Ducts.

22.

Elevator Forward Bellcrank Rig Pin Hole and Emergency Landing Gear Extension Pump.

23.

Elevator Trim Tab Cables.

06-50-00

Page 11 May 1/10

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

A 6 7

A

B

8 B

C

9 C

D

10 D

E

11 E

F

12 F

G

13 G

H

14 H

17

16

I

15

24 25 1900C

UC06B 035151AB.AI

Figure 3 (Sheet 2 of 3) Floor Access Panels (UB-1 and After; UC-1 and After)

Page 12 May 1/10

06-50-00

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

E. Floor Access Panels (UB-1 and After; UC-1 and After) Figure 3, Sheet 2 of 3 6.

Parking Brake Valve, Electroluminescent Power Supply Terminal Board, Thermocouple Resistors, Heatsink Diode Assembly, Subpanel and Fuel Control Panel Feeder and Feeder Base, Circuit Board Feeders, Avionics Feeder, Prop Synchrophaser Control Box, Circuit Breaker Panel Feeder Base, Recording Accelerometer, Stall Lift Computer, Electroluminescent Power Supply, No. 1 Avionics Bus Relay, No. 2 Avionics Bus Relay, No. 3 Avionics Relay, VHF COMM Antenna Connector.

7.

Annunciator P.C. Board Box Assembly, Aural Warning Amplifier, Cabin Altitude Warning Switch, Pilot and Co-Pilot Audio Amplifiers.

8.

Air-Conditioner Forward Blower Motor, Air-Conditioner Forward Evaporator, High and Low Speed Blower Motor Relays, Deice Boot Timer, P.C. Board Relay Box Assembly, Forward DME Antenna Connector.

9.

Pneumatic Pressure Regulator, Air Cycle Ejector and Ejector Muffler, Bleed Air Fail Warning Pressure Switches, Venturi Suction for Instrument Air, Air-Conditioner Receiver-Dryer, Transponder Antenna Connector, Hydraulic Plumbing.

10.

Aileron Servo for Autopilot, Flap Motor, Flap Motor Relays, LH and RH Generator Control Panel, LH and RH Overvolt Test Switch, Prop Deice Timer, Air-Conditioner Command Relay, Fuel Quantity Relay Panel.

11.

Air-Conditioner Plumbing.

12.

Air-Conditioner Aft Evaporator, Air-Conditioner Aft Blower Motor, High and Low Speed Blower Motor Relay, Beacon Light, Expansion Valve, Transponder Antenna Connector.

13.

VHF Antenna Connector.

14.

Air-Conditioning Ducts.

15.

Air-Conditioning Ducts, Control Cables, Aft DME Antenna Connector.

16.

LH Side Floorboard Panels.

A.

Control Cables, Air-Conditioning Ducts.

B.

Control Cables, Air-Conditioning Ducts.

C.

Control Cables, Air-Conditioning Ducts.

D.

Control Cables, Air-Conditioning Ducts.

E.

Control Cables, Air-Conditioning Ducts, Aileron Quadrant.

F.

Control Cables, Air-Conditioning Ducts.

G.

Control Cables, Air-Conditioning Ducts.

H.

Control Cables and Turnbuckles, Air-Conditioning Ducts.

I.

Control Cables, Air-Conditioning Ducts.

06-50-00

Page 13 May 1/10

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

A 6 7

A

B

8 B

C

9 C

D

10 D

E

11 E

F

12 F

G

13 G

H

14 H

17

16

I

15

24 25 1900C

UC06B 035151AB.AI

Figure 3 (Sheet 3 of 3) Floor Access Panels (UB-1 and After; UC-1 and After)

Page 14 May 1/10

06-50-00

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

F. Floor Access Panels (UB-1 and After; UC-1 and After) Figure 3, Sheet 3 of 3 17.

RH Side Floorboard Panels.

A.

Air-Conditioning Ducts, Plumbing.

B.

Air-Conditioning Ducts, Plumbing.

C.

Air-Conditioning Ducts, Plumbing.

D.

Air-Conditioning Ducts, Plumbing.

E.

Air-Conditioning Ducts, Plumbing.

F.

Air-Conditioning Ducts, Plumbing.

G.

Air-Conditioning Ducts, Plumbing.

H.

Air-Conditioning Ducts, Plumbing.

24.

Control Cables.

25.

Control Cables.

06-50-00

Page 15 May 1/10

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

C 18

19

B

17

A

33 36

15

DETAIL

C

25

23

22 21 16 27A

15 21 22

28A

14 13

29A 26

23

4

20 35 34 30

3

26

9

5

29B 7

31

2

9

6

DETAIL

B

06-50-00

A UC27B 031223AB.AI

Figure 4 (Sheet 1 of 2) Stabilizer Access Panels

Page 16 May 1/10

12

10 11

1

DETAIL

27B

8

24 32

28B

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

G. Stabilizer Access Panels (Figure 4, Sheet 1 of 2) 1.

Dorsal Fin.

2.

Aft Dorsal Fin.

3.

Seal.

4.

Forward Fairing.

5.

Vertical Stabilizer.

6.

Fillet.

7.

Rudder.

8.

Rudder Tab.

9.

Rudder Hinge Bracket.

10.

Rudder Torque Tube.

11.

Rudder Control Horn.

12.

Horizontal Stabilizer.

13.

Elevator.

14.

Elevator Trim Tab.

15.

Aft Tail Fairing.

16.

Rotating Beacon.

17.

Access Plate, FWD Horizontal Stabilizer.

18.

Access Plate, AFT Horizontal Stabilizer.

19.

Aft Navigation Light.

20.

Access Plate, Elevator Bellcrank.

21.

Access Plate, Inboard Elevator.

22.

Access Plate, Outboard Elevator.

23.

Access Plate, Taillet.

24.

Access Plate, Rudder Tab Actuator.

25.

LH Fairing Angle.

26.

Flux Valve (UC-1 and After).

27.

A (B) Inspection Area Access Panel (Kit 114-4060).

28.

A (B) Inspection Area Access Panel (Kit 114-4060).

29.

A (B) Inspection Area Access Panel (Kit 114-4060).

30.

Inspection Area Access Panel (Kit 114-4060).

31.

Inspection Area Access Panel (Kit 114-4060).

06-50-00

Page 17 May 1/10

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

C 18

19

B

17

A

33 36

15

DETAIL

C

25

23

22 21 16 27A

15 21 22

28A

14 13

29A 26

23

4

20 35 34 30

3

26

9

5

29B 7

31

2

9

6

DETAIL

B

06-50-00

A UC27B 031223AB.AI

Figure 4 (Sheet 2 of 2) Stabilizer Access Panels

Page 18 May 1/10

12

10 11

1

DETAIL

27B

8

24 32

28B

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

H. Stabilizer Access Panels (Figure 4, Sheet 2 of 2) 32.

Inspection Area Access Panel (Kit 114-4060).

33.

Access Plate.

34.

Elevator Aft Bellcrank Rig Pin Hole Access Panel.

35.

Elevator Aft Bellcrank Mount Bolt Access Panel.

36.

RH Fairing Angle.

06-50-00

Page 19 May 1/10

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

12 10

13

2

3

4 LEFT SIDE

11

2

1

4

VIEW LOOKING DOWN AT TOP OF WING

8 7

LEFT SIDE

3

9

RIGHT SIDE

9 7

7

5

6

8

7

VIEW LOOKING UP AT BOTTOM OF WING

RIGHT SIDE UC06B 023675AB.AI

Figure 5 (Sheet 1 of 4) Wing Access Panels (UA-1 and After; UB-1 and After)

Page 20 May 1/10

06-50-00

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

I. Wing Access Panels (UA-1 and After; UB-1 and After) (Figure 5, Sheet 1 of 4) 1.

Flap Switch.

2.

Fuel Quantity Probe.

3.

Fuel Quantity Probe, Low Level Sensor.

4.

Fuel Plumbing, Flight Controls.

5.

LH Center Section Fuel Cell and Collector Tank.

6.

RH Center Section Fuel Cell and Collector Tank.

7.

Flap Controls, Aileron Controls.

8.

Flap Actuator.

9.

Fuel Plumbing.

10.

Hydraulic Service Valves (LH Only).

11.

Refrigeration Spray Nozzle (LH Only).

12.

Hydraulic Filler (LH Only).

13.

Battery (RH Only).

06-50-00

Page 21 May 1/10

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

23

21

19

17

18

VIEW LOOKING DOWN AT TOP OF WING

14

LEFT SIDE

22

24

28 26

25 27 LEFT SIDE

20

15

RIGHT SIDE

16

VIEW LOOKING UP AT BOTTOM OF WING

RIGHT SIDE

UC06B 062523AA.AI

Figure 5 (Sheet 2 of 4) Wing Access Panels (UA-1 and After; UB-1 and After)

Page 22 May 1/10

06-50-00

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

J. Wing Access Panels (UA-1 and After; UB-1 and After) (Figure 5, Sheet 2 of 4) 14.

Bleed Air Plumbing (LH Only).

15.

Air Cycle Machine Filter (LH Only).

16.

Condenser Coil (RH Only).

17.

LH Fuel Lines, Deicer Lines.

18.

RH Fuel Lines, Deicer Lines.

19.

LH Fuel Probe.

20.

RH Fuel Probe.

21.

LH Fuel Probe.

22.

RH Fuel Probe.

23.

LH Fuel Line Connection.

24.

RH Fuel Line Connection.

25.

LH Heated Fuel Vent.

26.

RH Heated Fuel Vent.

27.

LH Flap Cables, Aileron Cables.

28.

RH Flap Cables, Aileron Cables.

06-50-00

Page 23 May 1/10

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

59

53

55 57

56

54

58

53

31

32 29

61

LEFT SIDE

50

44

46

39 37

41

RIGHT SIDE

36

35

38

45 40

49

51 42

33 63

LEFT SIDE

30

VIEW LOOKING DOWN AT TOP OF WING

43 48

60

62

52

47

34

VIEW LOOKING UP AT BOTTOM OF WING

RIGHT SIDE

UC06B 062522AA.AI

Figure 5 (Sheet 3 of 4) Wing Access Panels (UA-1 and After; UB-1 and After)

Page 24 May 1/10

06-50-00

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

K. Wing Access Panels (UA-1 and After; UB-1 and After) (Figure 5, Sheet 3 of 4) 29.

LH Aileron Actuator.

30.

RH Aileron Actuator.

31.

LH Remote Compass.

32.

RH Remote Compass.

33.

LH Aft Inboard Wing Fuel Cell.

34.

RH Aft Inboard Wing Fuel Cell.

35.

LH Flap Actuator.

36.

RH Flap Actuator.

37.

LH Aileron Pulley.

38.

RH Aileron Pulley.

39.

LH Aft Center Wing Fuel Cell.

40.

RH Aft Center Wing Fuel Cell.

41.

LH Fuel Plumbing, Electrical Connector.

42.

RH Fuel Plumbing, Electrical Connector.

43.

Aileron Trim Tab Cable (LH Only).

44.

LH Fuel Plumbing.

45.

RH Fuel Plumbing.

46.

LH Aileron Pulley Bracket, Fuel Plumbing.

47.

RH Aileron Pulley Bracket, Fuel Plumbing.

48.

LH Integral (Wet Wing) Fuel Cell.

49.

RH Integral (Wet Wing) Fuel Cell.

50.

LH Fuel Vent Plumbing.

51.

RH Fuel Vent Plumbing.

52.

Air Conditioner Blower Motor.

53.

Fuel Filler Cap.

54.

RH Leading Edge.

55.

LH Leading Edge.

56.

Vapor Cycle System.

57.

Air Cycle System and Power Pack.

58.

RH and Center Bus.

06-50-00

Page 25 May 1/10

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

59

53

55 57

56

54

58

53

31

32 29

61

LEFT SIDE

50

44

46

39 37

41

RIGHT SIDE

36

35

38

45 40

49

51 42

33 63

LEFT SIDE

30

VIEW LOOKING DOWN AT TOP OF WING

43 48

60

62

52

47

34

VIEW LOOKING UP AT BOTTOM OF WING

RIGHT SIDE

UC06B 062522AA.AI

Figure 5 (Sheet 4 of 4) Wing Access Panels (UA-1 and After; UB-1 and After)

Page 26 May 1/10

06-50-00

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

L. Wing Access Panels (UA-1 and After; UB-1 and After) (Figure 5, Sheet 4 of 4) 59.

LH and Center Bus.

60.

RH Inverter.

61.

LH Inverter.

62.

Vapor Cycle System.

63.

Air Cycle System.

06-50-00

Page 27 May 1/10

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

22

10

11 12

21

27

22

18

13

25

21

26

3

3 2

3

3

1

2

4 LEFT SIDE

VIEW LOOKING DOWN AT TOP OF WING

5 17

5

4

19

2

28 5

19 17 19 25

25 19 17 19

5

RIGHT SIDE

2

17 19

5

9 7

20

20

20

21

6 22

LEFT SIDE

29 14

8 18 8 5 15

8

20

16

5

20

21

8

23

29

20

20

7

20

6 22

24

VIEW LOOKING UP AT BOTTOM OF WING

RIGHT SIDE

UC57B 062408AB.AI

Figure 6 Wing Access Panels (UC-1 and After)

Page 28 May 1/10

06-50-00

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

M. Wing Access Panels (UC-1 and After) Figure 6 1.

Flap Switch.

2.

Fuel Probe, Low Level Sensor.

3.

Fuel Filler Cap.

4.

Inverter.

5.

Fuel Probe Access.

6.

Low Level Sensor.

7.

Wiggins Fittings (Strobe & Recog. Wiring).

8.

Fuel Fittings, Fuel Strainer.

9.

Fuel Pump, Drain Valve, Float Switch.

10.

Hydraulic Service Valves (Left Only).

11.

Refrigeration Pack Spray Nozzle (Left Only).

12.

Hydraulic Filler (Left Only).

13.

Battery (Right Only).

14.

Bleed Air Plumbing (Left Only).

15.

Air Cycle Machine Filter (Left Only).

16.

Condenser Air Inlet (Right Only).

17.

Flap Actuators.

18.

Air Cycle System And Powerpack.

19.

Flight Control Rigging.

20.

Integral (Wet Wing) Fuel Cell.

21.

Fuel Filter And Shutoff Valve.

22.

Landing Lights.

23.

Right Leading Edge - Bleed Air Lines.

24.

Left Leading Edge - Bleed Air Lines.

25.

Vapor Cycle System.

26.

Right and Center Bus.

27.

Left and Center Bus.

28.

External Power Access.

29.

Fuel Filter Drain Access.

06-50-00

Page 29 May 1/10

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

CHAPTER 07 - LIFTING AND SHORING TABLE OF CONTENTS SUBJECT

PAGE

GENERAL INFORMATION 07-00-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Lifting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Hoisting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Jacking . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

1 1 1 1

JACKING 07-10-00 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Three-Point Jacking (Preferred Procedure) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Lowering the Airplane After Three-Point Jacking . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Nose Jacking . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Lowering the Airplane After Nose Jacking . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Single-Point Jacking (For Wheel, Tire and Brake Maintenance Only) . . . . . . . . . . . . . . . . . . . . . . . . . Lowering the Airplane After Single-Point Jacking . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 201 202 202 203 203 204

SHORING 07-20-00 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Hoisting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201

07-CONTENTS

Page 1 Nov 1/09

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

List of Effective Pages CH-SE-SU

PAGE

DATE

07-LOEP

1

Nov 1/09

07-CONTENTS

1

Nov 1/09

07-00-00

1 thru 3

Nov 1/09

07-10-00

201 thru 207

Nov 1/09

07-20-00

201 thru 202

Nov 1/09

07-LOEP

Page 1 Nov 1/09

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

LIFTING AND SHORING GENERAL INFORMATION DESCRIPTION AND OPERATION

07-00-00 00

1. LIFTING Bleed air from the engines enters the cabin duct work through the ACM bypass valve and the ejector bypass valve. The ACM bypass valve begins opening first upon receiving a heat command from the cabin temperature controller. When the ACM bypass valve is fully open, the intergral open limit switch in the valve shunts the heat command from the cabin temperature controller to the ejector bypass valve, opening it. As the cabin temperature controller starts to issue cool commands, the ejector bypass valve begins closing. When the ejector bypass valve is fully closed, the cool command is shunted to the ACM bypass valve, closing it.

A. Hoisting WARNING: Prior to attaching the sling to the airplane, inspect the sling for damage or wear. Any defect could diminish the strength of the sling. CAUTION: Before hoisting the airplane, the fuel tanks must be drained and the airplane should be empty. The airplane may be hoisted for maintenance or parts replacement. A minimum overhead clearance of 17 feet and a hoist of at least 14,000 to 16,000 pound capacity are required for hoisting. If it is necessary to hoist the airplane with one or both engines removed, use a sling under the tail. The tail mooring hole can be used to hold the sling in position. The 99-590029-1 hoisting sling, or its equivalent, is required to hoist the airplane (Ref. Chapter 91-00-00, SPECIAL TOOLS). The hoisting sling is constructed of two broad band nylon straps supported by a structural steel crossbar with wire ropes for attachment to an overhead crane or hoist. The sling is designed to support approximately 33,000 pounds to ensure a high safety factor during hoisting.

07-00-00

Page 1 Nov 1/09

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

B. Jacking WARNING: When an airplane has experienced abnormal landing gear procedures of any type, as a safety precaution, place the airplane on jacks prior to performing any inspection or maintenance. Ensure that all three landing gears are down and locked prior to removing the airplane from the jacks. Never jack the airplane in an unsheltered area where winds in excess of 35 knots will be encountered. NOTE: Each of the three jacks used to lift the airplane should have a lifting capacity of at least 10,000 pounds. The Model 1900 Series Airliner is equipped with a three-point jack pad system. Two main gear jack pads are located on the center wing rear spar, just inboard of each nacelle. The nose gear jack pad is located near the aft edge of the nose gear wheel well. All three points are easily identified by the placarding, JACK PAD, adjacent to the jack points. CAUTION: The fuel must be evenly distributed in both wings to ensure stability while the airplane is on jacks. Jacking of an airplane for the purpose of landing gear operation, inspection, servicing or maintenance, should be accomplished within an enclosed building or hangar. In the interest of safety, should it become necessary to jack the airplane in the open, wind velocity in any direction and terrain variations must be compensated for prior to jacking the airplane.

Page 2 Nov 1/09

07-00-00

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 1 Hoisting

UC27B 042921AA.AI

Figure 2 10,000 Pound Capacity Jack

07-00-00

Page 3 Nov 1/09

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

LIFTING AND SHORING JACKING MAINTENANCE PRACTICES

07-10-00 200200

1. PROCEDURES WARNING: When an airplane has experienced abnormal landing gear procedures of any type, as a safety precaution, place the airplane on jacks prior to performing any inspection or maintenance. Ensure that all three landing gears are down and locked prior to removing the airplane from the jacks. CAUTION: Jacking of an airplane for the purpose of landing gear operation, inspection, servicing or maintenance should be accomplished within an enclosed building or hangar. In the interest of safety, should it become necessary to jack the airplane in the open, wind velocity in any direction and terrain variations must be compensated for prior to jacking the airplane.

A. Three-Point Jacking (Preferred Procedure) NOTE: Each of the three jacks used to lift the airplane should have a lifting capacity of at least 10,000 pounds. WARNING: Never jack the airplane in an unsheltered area where winds in excess of 35 knots will be encountered. CAUTION: Position airplane on a flat, hard, oil free surface prior to any jacking operations. Do not attempt any jacking operations outdoors during strong, gusty winds. If the airplane must be lifted outdoors, always face the airplane into the wind. Ensure area is clear of obstacles that could damage the airplane during raising operations. NOTE: If tail stand is installed, remove tail stand during jacking operations. (1) Position a tripod jack (3) at the nose of the airplane directly under each jack pad with two of the jack legs (4) perpendicular to the fuselage (Ref. Figure 201). (2) Position a tripod jack (3) at each wing of the airplane directly under each jack pad with two of the jack legs (4) parallel to the fuselage (Ref. Figure 202). (3) Extend jack extension (1) of each jack to contact securely with its associated jack pad (Ref. Figures 201 and 202). CAUTION: One person should be positioned forward of the airplane to ensure that the airplane remains level during three-point jacking. Keep the jack follower nut within two inches of the jack shoulder at all times during three-point jacking and lowering. (4) Operate the wing and nose jacks together until the clearance between the tires and the ground is at least 2 inches (51 mm). (5) Tighten the follower nut (2) of each jack (3) against the jack shoulder and release pressure from jack pump cylinder (6) (Ref. Figures 201 and 202). (6) Install a tail stand (2) at the aft fuselage mooring point (1) of the airplane and bring it into secure contact with the ground (Ref. Figure 203).

07-10-00

Page 201 Nov 1/09

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

B. Lowering the Airplane After Three-Point Jacking WARNING: Ensure that the landing gear is in the fully down and locked position. Failure to do so may result in injury to personnel and damage to the airplane. CAUTION: Ensure area is clear of obstacles that could damage the airplane during lowering operations. Keep the jack follower nut within two inches of the jack shoulder at all times during three-point jacking and lowering. Never lower just one side of the airplane at a wing jack point. Imbalance could cause the airplane to topple from the jack. (1) Ensure that the landing gear is in the fully down and locked position. (2) Wipe down the extended surface of each landing gear strut using a shop rag dampened with hydraulic fluid. CAUTION: One person should be positioned forward of the airplane to ensure that the airplane remains level during the lowering operation. (3) Ensure that the airplane weight is off the tail stand (2), then remove the tail stand (2) (Ref. Figure 203). CAUTION: Keep the jack follower nut within two inches of the jack shoulder at all times during three-point jacking and lowering. (4) Operate jacks to loosen the jack follower nuts (2), then slowly release pressure from the jack pump cylinder (6) to lower the airplane slowly to the ground (Ref. Figures 201 and 202). CAUTION: Ensure that all jacks are fully retracted and that the airplane has settled prior to jack removal. Damage to the airplane may occur. (5) Remove the jacks (3) from under the airplane.

C. Nose Jacking CAUTION: Position airplane on a flat, hard, oil free surface prior to nose jacking operations. Do not attempt nose jacking operations outdoors if the wind velocity will exceed 30 knots. If the airplane must be lifted outdoors, always face the airplane into the wind. Ensure area is clear of obstacles that could damage the airplane during raising operations. NOTE: If tail stand is installed, remove tail stand during jacking operations. (1) Set the parking brake and place airplane chocks forward and aft of the main landing gear tires. (2) Position a tripod jack (3) at the nose of the airplane directly under the jack pad with two of the jack legs (4) perpendicular to the fuselage (Ref. Figure 201). (3) Extend jack extension (1) to contact securely with the jack pad (Ref. Figure 201). CAUTION: Keep the jack follower nut within two inches of the jack shoulder at all times during nose jacking and lowering.

Page 202 Nov 1/09

07-10-00

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (4) Operate the nose jack until the clearance between the tire and the ground is at least 2 inches (51 mm). (5) Tighten the follower nut (2) of the nose jack (3) against the jack shoulder and release pressure from the jack pump cylinder (6) (Ref. Figure 201). (6) Install a tail stand (2) at the aft fuselage mooring point (1) of the airplane and bring it into secure contact with the ground (Ref. Figure 203).

D. Lowering the Airplane After Nose Jacking CAUTION: Ensure area is clear of obstacles that could damage the airplane during lowering operations. Ensure that the nose landing gear is in the fully down and locked position. Failure to do so may result in injury to personnel and damage to the airplane. Keep the jack follower nut within two inches of the jack shoulder at all times during nose jacking and lowering. (1) Ensure that the nose landing gear is in the fully down and locked position. (2) Wipe down the extended surface of the nose landing gear strut using a shop rag dampened with hydraulic fluid. (3) Ensure that the airplane weight is off the tail stand (2), then remove the tail stand (2) (Ref. Figure 203). (4) Operate nose jack to loosen the jack follower nut (2), then slowly release pressure from the jack pump cylinder (6) to lower the airplane slowly to the ground (Ref. Figure 201). (5) Remove the jack (3) from under the airplane.

E. Single-Point Jacking (For Wheel, Tire and Brake Maintenance Only) WARNING: This procedure is intended for wheel and brake maintenance only. Never perform any other landing gear maintenance procedures using the Single-Point Jacking Procedure. Never jack the airplane in an unsheltered area where winds in excess of 35 knots will be encountered. NOTE: Any jack used to lift the airplane should have a lifting capacity of at least 10,000 pounds. CAUTION: Position airplane on a flat, hard, oil free surface prior to any jacking operations. Do not attempt any jacking operations outdoors during strong, gusty winds. If the airplane must be lifted outdoors, always face the airplane into the wind. Ensure area is clear of obstacles that could damage the airplane during raising operations. NOTE: If tail stand is installed, remove tail stand during jacking operations. (1) Place airplane chocks forward and aft of the landing gear wheels that are not being jacked. NOTE: In ice or snow conditions, ice-grip wheel chocks are preferred. Sandbags may be used if ice-grip chocks are not available, or if the airplane is parked on a steel mat.

07-10-00

Page 203 Nov 1/09

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (2) Position a tripod jack (3) at the wing of the airplane directly under the jack pad with two of the jack legs (4) parallel to the fuselage (Ref. Figure 202). (3) Extend jack extension (1) of the jack to contact securely with its associated jack pad (Ref. Figure 202). CAUTION: Keep the jack follower nut within two inches of the jack shoulder at all times during jacking and lowering procedures. (4) Install a strut limiter (Special Tools, Figure 1, Chapter 91-00-00) to the landing gear strut being worked. (5) Operate the jack as required to raise the affected wheel until the clearance between the tires and the ground is at least 2 inches (51 mm). (6) Tighten the follower nut (2) of the jack (3) against the jack shoulder and release pressure from the jack pump cylinder (Ref. Figures 202).

F. Lowering the Airplane After Single-Point Jacking CAUTION: Ensure area is clear of obstacles that could damage the airplane during lowering operations. (1) Wipe down the extended surface of the landing gear strut using a shop rag dampened with hydraulic fluid. CAUTION: Keep the jack follower nut within two inches of the jack shoulder at all times during jacking and lowering procedures. (2) Operate jack to loosen the jack follower nut (2), then slowly release pressure from the jack pump cylinder to lower the airplane slowly to the ground (Ref. Figure 202). CAUTION: Ensure that the jack is fully retracted and that the airplane has settled prior to jack removal. Damage to the airplane may occur. (3) Remove the jack (3) from under the airplane. (4) Remove the strut limiter from the landing gear strut.

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1. JACK EXTENSION 2. FOLLOWER NUT 3. JACK 4. JACK LEGS 5. JACK HANDLE 6. JACK PUMP

A

1 2 3

4

5

6

DETAIL

A

UC27B 043243AB.AI

Figure 201 Nose Gear Jacking Point

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1. JACK EXTENSION 2. FOLLOWER NUT 3. JACK 4. JACK LEGS 5. FUSELAGE

A A

1 2

5

3

4

VIEW LOOK AFT FROM NOSE OF AIRPLANE (LH SHOWN, RH OPPOSITE) DETAIL

A

Figure 202 Main Gear Jacking Point

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1. MOORING POINT 2. TAIL STAND

A

1

2

DETAIL

A

UE27B 043244AA.AI

Figure 203 Tail Stand Location

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LIFTING AND SHORING SHORING MAINTENANCE PRACTICES

07-20-00 200200

1. PROCEDURES A. Hoisting WARNING: Prior to attaching the sling to the airplane, inspect the sling for damage or wear. Any defect could diminish the strength of the sling. CAUTION: Before hoisting the airplane, the fuel tanks must be drained and the airplane should be empty. The airplane may be hoisted for maintenance or parts replacement. An overhead crane capable of 20 feet of vertical lift is required to lift the aircraft four feet from the ground, allowing for one foot of vertical stretch in the nylon webbing. NOTE: If it is necessary to hoist the airplane with one or both engines removed, use a sling under the tail of the airplane. The airplane may be hoisted for maintenance or parts replacement. A minimum overhead clearance of 17 feet and a hoist of at least 14,000 to 16,000 pound capacity are required for hoisting. If it is necessary to hoist the airplane with one or both engines removed, use a sling under the tail. The tail mooring hole can be used to hold the sling into position. The 99-590029-1 hoisting sling, or its equivalent, is required to hoist the airplane (Ref. Chapter 91-00-00, SPECIAL TOOLS). The hoisting sling is constructed of two broad band nylon straps supported by a structural steel crossbar with wire ropes for attachment to an overhead crane or hoist. The sling is designed to support approximately 33,000 pounds to ensure a high safety factor during hoisting. The straps of the sling encircle the fuselage when the airplane is being hoisted as follows: (1) Place one nylon band at fuselage station 213.25, located just forward of the first cabin window (Ref. Figure 201). Place the other nylon band at fuselage station 373.75, located on the aft side of the sixth cabin window. (2) Attach the sling to the cable or chain on an overhead hoisting crane at a position in line with fuselage station 286.00. (3) Guy ropes between the wing tip and lift point or a man stationed at each wing tip may be required to offset any imbalanced weight in the wings. (4) Remove all loose equipment from the airplane and hoist the airplane smoothly. Take up slack in the hoist and begin the lift gently, watching carefully for any unbalance.

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F.S. 373.75 PLACE AFT STRAP HERE

F.S. 213.25 PLACE FWD STRAP HERE

PICK UP POINT TO CORRESPOND WITH F.S. 286.00 CONNECTING LINK

NYLON STRAP CABLE

NYLON STRAP

BAR

AIRPLANE HOISTING SLING

Figure 201 Hoisting Sling Attach Point

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CHAPTER 08 - LEVELING AND WEIGHING TABLE OF CONTENTS SUBJECT

PAGE

GENERAL INFORMATION 08-00-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Special Tools and Equipment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1

WEIGHING AND BALANCING 08-10-00 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Procedure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Weighing the Airplane . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201

LEVELING 08-20-00 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Procedure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Leveling the Airplane . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201

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LEVELING AND WEIGHING GENERAL INFORMATION DESCRIPTION AND OPERATION

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1. SPECIAL TOOLS AND EQUIPMENT Special tools listed in Table 1 as meeting federal, military or supplier specifications are provided for reference only and are not specifically required by Hawker Beechcraft Corporation. Any product conforming to the specification listed may be used. The products included in these Tables have been tested and approved for aviation usage by Hawker Beechcraft Corporation, by the supplier, or by compliance with the applicable specifications. Generic or locally manufactured products which conform to the requirements of specification may be used even though not included in the Tables. Only the basic number of each specification is listed. No attempt has been made to update the listing to the latest revision. It is the responsibility of the technician or mechanic to determine the current revision of the applicable specification prior to usage of the product listed. This can be done by contacting the supplier of the product to be used. Table 1 Special Tools Tool Name

Part Number

Supplier

Use

1. Plumb Bob Support Assy.

114-590025-1

Hawker Beechcraft Corporation 9709 E. Central, Wichita, KS 67201

Leveling the airplane.

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LEVELING AND WEIGHING WEIGHING AND BALANCING MAINTENANCE PRACTICES

08-10-00 200200

1. PROCEDURE A. Weighing the Airplane Periodic weighing of the airplane may be required to keep the Basic Empty Weight current (Ref. Figures 201, 202 and 203). Frequency of weighing is to be determined by the operator. All changes to the airplane affecting weight and/or balance are the responsibility of the airplane operator. (1) The airplane may be weighed on wheels or jack points. Three jack points are provided. One is on the nose section of the fuselage at station 83.5 and the two main support points are on the wing center section rear spar at station 326.1. Wheel reaction locations should be measured as described in Step (6). (2) Fuel should be drained preparatory to weighing. Fuel is drained from the drain ports with the airplane in static ground attitude. NOTE: (UA-1 and after, UB-1 and after) After tanks are drained, 8.6 pounds of trapped fuel remains in the airplane at fuselage station 301.0. The remainder of the unusable fuel (drainable fuel) is 32.0 pounds at station 297.9 and must be added to a drained system to obtain the aircraft Basic Empty Weight configuration. (UC-1 and after) After tanks are drained, 14.6 pounds of trapped fuel remains in the airplane at fuselage station 278.6. The remainder of the unusable fuel (drainable fuel) is 58.6 pounds at station 304.7 and must be added to a drained system to obtain the aircraft Basic Empty Weight configuration. (3) Engine oil must be at the full level in each tank. Total engine oil aboard when both tanks are full is 57.5 pounds at an arm of 249.3 inches. (4) To determine airplane configuration at time of weighing, installed equipment is checked against the airplane equipment list or superseding forms. All equipment must be in its proper place during weighing. (5) The airplane must be supported on the scales in a level attitude. Leveling reference points are located on the forward entrance door frame. Leveling is accomplished with a plumb bob or optical methods. Jack pad weighing may require the nose gear shock to be secured in the static position to prevent its extension. Leveling for wheel weighing may be accomplished by varying the amounts of air in the shocks and tires. (6) Measurement of the reaction arms for a wheel weighing is made using the nose reference point (FS 95.89). Using a steel measuring tape, measure the distance (with the airplane level on the scales) from the reference (a plumb bob hung from the center of the reference point) to the axle center line of the nose gear and then from the later point to the main wheel axle center line. The main wheel axle center line is best located by stretching a string parallel to the fuselage center line. The locations of the wheel reactions will be approximately at an arm of 315 inches from main wheels and 30 inches for the nose wheel.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (7) The Basic Empty Weight and Moment must include full oil and unusable fuel as well as a full charge for any other fluid system in the airplane (e.g. hydraulic or oxygen). The appropriate weights and moments must be added on the weighing form if any of these items are not in the airplane are subtracted, e.g. usable fuel. (8) Weighing should always be made in an enclosed area which is free from air currents. The scales used should be properly calibrated and certified in accordance with the Bureau of Standards. NOTE: Each new airplane is delivered with a completed sample loading, basic empty weight and center of gravity (C.G.), and equipment list, all pertinent to that specific airplane. It is the owner’s responsibility to ensure that changes in equipment are reflected in a new weight and balance and in an addendum to the equipment list. There are may ways of doing this; it is suggested that a running tally of equipment changes and their effect on basic empty weight and C.G. is a suitable means for meeting both requirements. The current equipment list and basic empty weight and C.G. information must be retained with the airplane when it changes ownership. Hawker Beechcraft Corporation cannot maintain this information; the current status is known only to the owner. If these papers become lost, the FAA will require that the airplane be re-weighed to establish the basic empty weight and C.G. and that an inventory of installed equipment be conducted to create a new equipment list. It is recommended that duplicate copies of the Basic Empty weight and Balance sheet and the Equipment List be made and kept in an alternate location in the event the original is misplaced.

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Figure 201 Basic Weight and Balance Worksheet (UA-1 and After; UB-1 and After)

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Figure 202 Basic Weight and Balance Worksheet (UC-1 and After)

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Figure 203 Dimensional Data

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LEVELING AND WEIGHING LEVELING MAINTENANCE PRACTICES

08-20-00 200200

1. PROCEDURE A. Leveling the Airplane The airplane may be leveled by placing the airplane on jacks and raising or lowering the jacks (Ref. Figure 201). Use a precision level on the aft baggage compartment floorboard and raise or lower the jacks on the high side of the airplane to level the airplane laterally. To level the airplane longitudinally, attach a cord and plumb bob to the upper Phillips head screw just aft of the forward cabin door. Raise or lower the nose-gear jack as required to pass the cord over the center line of a second Phillips head screw directly near the bottom of the door. Suspending the plumb bob in a can of light engine oil will dampen its movement.

Figure 201 Leveling the Airplane

08-20-00

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CHAPTER 09 - TOWING AND TAXIING TABLE OF CONTENTS SUBJECT

PAGE

TOWING 09-10-00 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Towing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201

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TOWING AND TAXIING TOWING MAINTENANCE PRACTICES

09-10-00 200200

1. PROCEDURES A. Towing There are two different tow bars available. The 50-59001 is intended to be used only as a manual tow bar. Even though this tow bar has a hole in the handle which may fit a tractor tow hitch, the bar should not be used with any tow vehicle. 50-590017 may be purchased for use with a tow vehicle (Ref. Chapter 91-00-00, SPECIAL TOOLS). With the tow bar connected to the towing lugs on the upper torque knee fitting of the nose strut, the airplane can be steered with the nose wheel when moving it by hand or with a tug (Ref. Figure 201). Although steering is automatic when the airplane is being towed by the nose strut, someone should ride in the pilot's seat to operate the brakes in the event of an emergency. CAUTION: Do not tow the airplane with rudder locks installed, except on airplanes equipped with power steering, as severe damage to the steering linkage can result. Never exceed the turning limits marked on the nose gear strut during ground handling (Ref. Figure 201). The nose gear steering stop lugs are designed to withstand the loads normally imposed through steering from the cockpit, not to prevent turn limitations from being exceeded during towing. It is possible to overcome the stop during ground handling and damage the steering linkage and nose strut. If the steering stop limitations are exceeded inspect the nose gear steering stop lugs for cracks, bending or distortion (Ref. Chapter 5-50-00, INSPECTION IN THE EVENT OF A BENT NOSE STEERING STOP). When using a tug, observe turn limits marked on the nose gear strut to prevent damage to nose gear. When spotting the airplane, do not push on the propeller or control surfaces. CAUTION: Never tow or taxi with a deflated strut. Even brief towing or taxiing with a deflated strut may cause severe damage.

09-10-00

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Figure 201 Towing the Airplane

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CHAPTER 10 - PARKING, MOORING, STORAGE AND RETURN TO SERVICE TABLE OF CONTENTS SUBJECT

PAGE

PARKING 10-10-00 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Parking . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Control Locks . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 201 201

STORAGE 10-10-01 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Storage . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Engine Preservation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Airplane Preservation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Flight Ready Storage . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Short Term Storage (30 Days or Less) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Intermediate Storage (90 Days or Less) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Long Term Storage (90 days or Longer) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . For Airplanes Scheduled to Remain Inactive for Longer than One Year . . . . . . . . . . . . . . . . . . . . . . . For Airplanes Scheduled to Remain Inactive for Longer than Two Years . . . . . . . . . . . . . . . . . . . . . . In Storage Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 201 201 201 201 202 203 204 206 206 206

MOORING 10-20-00 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Mooring . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201

RETURN TO SERVICE 10-30-00 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Engine Return to Service . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Flight Ready Storage (Less Than 7 Days) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Short Term Storage (8 - 29 Days) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Intermediate Storage (30 - 89 Days) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Long Term Storage (More Than 90 Days) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

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PARKING, MOORING, STORAGE AND RETURN TO SERVICE PARKING MAINTENANCE PRACTICES 1. PROCEDURES A. Parking Set the brakes by depressing the pilot’s brake pedals and pulling out the parking brake control handle. Do not attempt to lock the parking brake by applying force to the parking brake handle. It controls a valve only and cannot apply pressure to the brake system. To release the brakes, depress the brake pedals and push the parking brake control handle in. CAUTION: Do not set the parking brake during low temperatures when the accumulation of moisture may cause the brakes to freeze, or when they are hot from severe use.

B. Control Locks WARNING: The flight control gust locks provided by Hawker Beechcraft Corporation for its products are in compliance with federal regulations to provide an unmistakable warning to the pilot when the lock is engaged. When necessary or desirable to use flight control gust locks, use only the flight control gust lock assembly specified by Hawker Beechcraft Corporation for that particular airplane. When a flight control gust lock assembly is used, the lock must be correctly and fully installed, including the rudder pedal lock and throttle control lock. The control lock (29, Table 7, Chapter 91-00-00) consist of a U-shaped clamp and two pins connected by a chain. The pins lock the primary flight controls and the U-shaped clamp fits around the engine power control levers and serves to warn the pilot not to start the engines with the control locks installed. It is important that the locks be installed or removed together to prevent the possibility of an attempt to taxi or fly the airplane with the power levers released and the pins still installed in the flight controls. Perform the Control Lock Installation Procedure (Ref. Chapter 27-70-00). WARNING: Before staring the engines, remove the control locks. Remove the control locks before towing the airplane. If towed while the rudder lock is installed, serious damage to the steering linkage can result.

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PARKING, MOORING, STORAGE AND RETURN TO SERVICE STORAGE MAINTENANCE PRACTICES 1. GENERAL A. Storage The following procedures are designed to protect the airplane while it is scheduled to be inactive for periods of time as outlined below. These procedures should be considered the minimum necessary to protect the airplane. Each operator should take whatever additional Steps to protect the airplane they consider necessary. These procedures are not intended to be accomplished on airplanes that are not flown because of extensive maintenance activities or for reasons other than scheduled periods of inactivity.

B. Engine Preservation Aircraft in flyable storage should have their engines run at least once per week. Refer to Pratt & Whitney Maintenance Manual PN 3032842, Chapter 72-00, Preservation Procedures, for detailed instructions.

2. AIRPLANE PRESERVATION A. Flight Ready Storage MECH

INSP

(1) Perform the following Steps: (a) Check tires for proper inflation. (b) Install prop restraint and engine exhaust covers. (c) Install inlet covers. (d) Install air conditioner condenser plug into right inboard wing leading edge or seal opening with barrier material. (e) Install pitot head covers. (f) Statically ground the airplane. (g) Position nose of airplane into prevailing wind, if possible. (h) Chock both main landing gear wheels. (i) Install approved gust locks (29, Table 7, Chapter 91-00-00).

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B. Short Term Storage (30 Days or Less) Airplanes scheduled to be inactive for a period of less than 30 days, in addition to the Flight Ready Storage preservation procedures, should perform the following Steps: MECH (1) Airframe (a) Inspect wings, wheel well and engine compartment areas for fuel leaks. Repair all leaks. (b) Ensure main and auxiliary fuel tanks are approximately 3/4 full. 1 Prevent microbial growth in the fuel tanks by adding Biobor JF or equivalent in sufficient quantity to maintain a concentration of 270 PPM (63, Table 1, Chapter 91-00-00). 2 Check fuel vents are clear. Clean if blocked. Cover fuel vents with screen mesh. 3 Fill fuel bladder cells to capacity to minimize fuel vapor and protect the cells inner liner. CAUTION It is imperative that the storage areas of nickel-cadmium batteries be separated (in different buildings if possible) from lead acid batteries so that no cross contamination, even from fumes, is possible. (c) Remove battery from airplane and place in suitable storage area. 1 Clean battery compartment, quick disconnect plug, cables, and vent hoses with a solution of 5 ounces of boric acid dissolved in one gallon of water, then rinse with clean water and allow to dry. Seal the battery vent tubes and cover quick disconnect plug with barrier material. Secure with tape. 2 Seal battery compartment with barrier material and secure with tape. (d) Fill brake reservoir and hydraulic landing gear reservoir to operational levels. Inspect system for leaks. Repair leaks prior to placing aircraft into storage. (e) Apply Agemaster preservative or equivalent to deicer boots and propeller deicer boots (177, Table 1, Chapter 91-00-00). (f) Install landing gear down locks. (g) Cover air outlet port (top of aft nacelle) with barrier material. Secure with tape.

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C. Intermediate Storage (90 Days or Less) Airplanes scheduled to be inactive for a period of less than 90 days, in addition to the Short Term Storage preservation procedures should perform the following Steps: MECH

INSP

(1) Cabin (a) Drain and clean lavatory, if installed. Reinstall empty lavatory unit. Install placard on or near flush button that reads “LAV NOT SERVICED. DO NOT FLUSH”. (2) Propellers (a) Wrap propeller spinner and blade bases with barrier material. Secure with tape. CAUTION Do not clean the propeller with Methyl Ethyl Ketone or Methyl Propyl Ketone (b) Remove dirt, oil, and bug accumulation from propellers with water and a soft brush. Clean propeller with water, denatured alcohol or naphtha. (3) Avionics (a) Clean and cover any equipment sensitive to dust or moisture and comply with any additional requirements recommended by the manufacturer. (b) Remove all electrical cockpit displays such as EFIS and/or EHIS displays, LCD displays for TCAS or altimeters and store in suitable storage area preferably in a controlled environment NOTE The following item may be skipped if airplane is to be stored in a hangar. (c) If airplane is equipped with weather radar, install fresh desiccant in the wave guide of the weather radar located in the avionics bay. (9, Table 1, Chapter 91-00-00). (4) Airframe (a) Install cabin and cockpit seat covers. NOTE Compliance with the following items is not required if airplane is to be stored in a hangar. (b) Cover cabin windows with barrier material. Secure with tape. (c) Seal cockpit storm windows with barrier material. Secure with tape. (d) Seal emergency hatches and cargo door seams with barrier material. Secure with tape.

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INSP

(e) Install barrier material around radome to fuselage joint. Secure with tape. (f) Ensure all fuselage and wing drain holes are clear. Clean if blocked. Cover all drain holes with screen mesh. Secure with tape. (g) Cover avionics bay door seals with barrier material. Secure with tape. (h) Cover generator inlet scoop with barrier material. Secure with tape. (i) Cover oil cooler inlet with barrier material. Secure with tape. (j) Moor airplane (Ref. 10-20-00).

D. Long Term Storage (90 days or Longer) Airplanes scheduled to be inactive for a period exceeding 90 days, in addition to the Intermediate Storage preservation procedures should perform the following Steps: MECH (1) Preparation (a) Perform the following Steps if not previously accomplished: 1 Cover cabin windows with barrier material. Secure with tape. 2 Seal cockpit storm windows with barrier material. Secure with tape. 3 Seal emergency hatches and cargo door seams with barrier material. Seal with tape. 4 Install barrier material around radome to fuselage joint. Secure with tape. 5 Ensure all fuselage and wing drain holes are clear. Clean if blocked. Cover all drain holes with screen mesh. Secure with tape. 6 Cover avionics bay doors seals with barrier material. Secure with tape. 7 Cover generator inlet scoop with barrier material. Secure with tape. 8 Cover oil cooler inlet with barrier material. Secure with tape. (b) Remove windshield wipers, wrap with barrier material and secure with tape. Store the wipers as appropriate (Ref. Chapter 30-40-00) (c) Cover windshields with barrier material. Secure with tape. (d) Clean all exposed antennas and connections. (e) Cover stall warning unit with barrier material. Ensure barrier material does not apply pressure to the vane. Secure with tape.

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INSP

(f) Cover landing light and taxi light lenses with barrier material. Secure with tape. (g) Cover ram air intake with barrier material. Secure with tape. (2) Flight Controls (a) Lubricate all hinge pins, bearings, bellcranks, chains, control rods and quadrants. Lightly coat with corrosion preventive compound (11, Table 1, Chapter 91-00-00). (b) Coat flap tracks and rollers with corrosion preventive compound (11, Table 1, Chapter 91-00-00). Retract flaps. (3) Fuel Cells (UA-1 and After, UB-1 and After) (a) Drain fuel bladder cells, clean inner cell and spray or rub a thin coating of light engine oil on inner liners of all fuel cells (10, Table 1, Chapter 91-00-00). (4) Landing Gear NOTE: It is recommended that unserviceable tires be used on airplanes stored for more than 90 days. (a) Clean brakes and apply coating of primer (12, Table 1, Chapter 91-00-00). Wheel removal required. (b) Apply corrosion preventive compound (11, Table 1, Chapter 91-00-00) to inner wheel brake keyways. (c) Touch up all spots where paint has been chipped from the wheels. (Primer may be used for this purpose). (d) Touch up wheel bolts and nuts with corrosion preventive compound (11, Table 1, Chapter 91-00-00). (e) Coat the exposed surfaces of the shock strut pistons and nose gear shimmy damper piston with preservative hydraulic fluid (13, Table 1, Chapter 91-00-00). (f) Wrap lower portion of the landing gear in barrier material. Secure with tape. (g) Fabricate hardwood collars for the strut pistons to prevent bottoming of the struts when deflating (Ref. Chapter 91-00-00, Special Tools, Figure 1 (Sheet 6 of 9)). (h) Install collars over barrier material. Secure with tape. CAUTION Never tow or taxi with a deflated strut. Even brief towing or taxiing with a deflated strut may cause severe damage. (i) Deflate struts until they rest on the wooden collars.

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E. For Airplanes Scheduled to Remain Inactive for Longer than One Year MECH

INSP

(1) Perform the following Steps: (a) Remove engines for storage in containers. (b) Store propellers in protected environment.

F. For Airplanes Scheduled to Remain Inactive for Longer than Two Years MECH

INSP

MECH

INSP

(1) Perform the following Step: (a) Perform internal corrosion inspection of propellers (disassembly required).

G. In Storage Inspection

(1) For airplanes scheduled to remain inactive, perform the following Steps weekly: (a) Perform a walk around inspection of the airplane. 1 Check all covers and barriers for condition and security of attachment. 2 Check for possible external damage and FOD. Check static wicks and flight control surface trailing edges for evidence of lightening strikes. 3 Check for evidence of fuel leaks. 4 Check all access areas such as flight control gaps and wheel wells for possible accumulation of debris such as bird or insect nests. 5 Check static ground for condition and attachment. 6 Check landing gear for evidence of leaks. Do not remove barrier material for this check. (b) Check drain sump of the fuel tanks for water and evidence of microbiological contamination. (c) Check tires for proper inflation.

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INSP

(2) Perform the following Step every 30 days: (a) Apply Agemaster preservative or equivalent to propeller and leading edge deice boots (177, Table 1, Chapter 91-00-00). (3) Perform the following Step every 90 days: CAUTION The airplane may be towed to accomplish this Step. Do not tow the airplane any appreciable distance with the struts deflated. Do not turn the nose landing gear while moving the airplane to accomplish this Step. The airplane may be towed only as far as necessary to rotate the tires 90 degrees. (a) To prevent flat spots from occurring, rotate wheels a minimum of 90 degrees.

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PARKING AND MOORING MOORING MAINTENANCE PRACTICES

10-20-00 200200

1. PROCEDURES A. Mooring Three mooring eyes are provided, one on each outer main wing spar and one on the rear fuselage ventral fin. To moor the airplane, chock the wheels fore and aft, install the control lock and tie the airplane down at all three mooring eye points and the landing gear torque knees on the nose and main gear. Avoid over-tightening the rear line and pulling the nose of the airplane up so far that wind will create lift on the wings. If extreme weather is anticipated, it is advisable to nose the airplane into the wind. When mooring the airplane, install the engine inlet and exhaust covers, pitot mast covers, propeller restraints and the optional nose and main landing gear shock strut limiters (Ref. Figure 201).

Figure 201 Mooring the Airplane

10-20-00

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10-30-00 200200

PARKING, MOORING, STORAGE AND RETURN TO SERVICE RETURN TO SERVICE MAINTENANCE PRACTICES 1. PROCEDURES A. Engine Return to Service Refer to Pratt & Whitney Maintenance Manual P/N 3041195, Chapter 72-00, Depreservation Procedures, for detailed instructions.

B. Flight Ready Storage (Less Than 7 Days) Mechanic

Inspector

Mechanic

Inspector

(1) No special requirements. Perform preflight inspection in accordance with the Pilots Operating Manual/FAA Approved Flight Manual. (2) Review airplane maintenance records to ensure airplane is in compliance with all applicable Inspection Requirements, Service Bulletins, Airworthiness Directives and Regulations.

C. Short Term Storage (8 - 29 Days)

(1) Remove all tape, barrier material, covers and plugs installed on airplane. (2) Ensure all fuselage and wing drain holes have screen mesh material removed. (3) Install serviceable airplane battery. (4) Remove landing gear downlocks. (5) Adjust fuel levels to meet Pilots Operating Manual/FAA Approved Flight Manual requirements. (6) Review airplane maintenance records to ensure airplane is in compliance with all applicable Inspection Requirements, Service Bulletins, Airworthiness Directives and Regulations.

10-30-00

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D. Intermediate Storage (30 - 89 Days) In addition to Short Term Storage return to service procedures, the following Steps should be accomplished: Mechanic

Inspector

(1) Remove all covers from avionics equipment, if installed. (2) Install cockpit electronic displays, if previously removed. (3) Remove all cockpit and cabin seat covers, if installed. (4) Remove all desiccant from wave guide of weather radar, if equipped. (5) Check landing gear hydraulic accumulator precharge. Service as required (Ref. Chapter 32-30-00). (6) Service lavatory, if installed. Remove “LAV NOT SERVICED. DO NOT FLUSH” placard. (7) Review airplane maintenance records to ensure airplane is in compliance with all applicable Inspection Requirements, Service Bulletins, Airworthiness Directives and Regulations.

E. Long Term Storage (More Than 90 Days) In addition to the Intermediate Storage depreservation procedures the following Steps should be accomplished: Mechanic

Inspector

(1) Install windshield wipers (Ref. Chapter 30-40-00). CAUTION Never tow or taxi with a deflated strut. Even brief towing or taxiing with a deflated strut may cause severe damage. (2) Inflate and service landing gear struts. Remove wooden collars from the landing gear struts (Ref. MAIN LANDING GEAR SHOCK ABSORBER SERVICING, Chapter 32-10-00 and NOSE GEAR SHOCK ABSORBER SERVICING, Chapter 32-20-00). (3) Clean corrosion preventative compound from the flap tracks. (4) Perform complete airplane lubrication (Ref. Chapter 12-20-00). (5) Review airplane maintenance records to ensure airplane is in compliance with all applicable Inspection Requirements, Service Bulletins, Airworthiness Directives and Regulations.

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CHAPTER 11 - PLACARDS AND MARKINGS TABLE OF CONTENTS SUBJECT

PAGE

EXTERIOR PLACARDS AND MARKINGS (UA-1 AND AFTER; UB-1 AND AFTER) 11-20-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Model Designation Placard (UA-1 and After; UB-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Exterior Placards and Markings (UA-1 and After; UB-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

1 1 1 1

EXTERIOR PLACARDS AND MARKINGS (UC-1 AND AFTER) 11-21-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Model Designation Placard (UC-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Exterior Placards and Markings (UC-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

1 1 1 1

FLIGHT CONTROL RIG PIN FUSELAGE AND EMPENNAGE PLACARDS 11-30-00 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Fuselage Rig Pin Placards . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Empennage Rig Pin Placard . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

11-CONTENTS

201 201 201 201

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List of Effective Pages CH-SE-SU

PAGE

DATE

11-LOEP

1

May 1/10

11-CONTENTS

1

Nov 1/09

11-20-00

1 thru 14

Nov 1/09

11-21-00

1 thru 13

May 1/10

11-30-00

201 thru 206

Nov 1/09

C2

11-LOEP

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PLACARDS AND MARKINGS EXTERIOR PLACARDS AND MARKINGS (UA-1 AND AFTER; UB-1 AND AFTER) DESCRIPTION AND OPERATION

11-20-00 00

1. GENERAL A. Model Designation Placard (UA-1 and After; UB-1 and After) The model designation placard is located on the left side of the airplane on the forward cabin entrance doorframe (Ref. Figure 1, Placard 22). The door must be open to observe the placard. The placard identifies the airplane by its model number and serial number. Should a question arise concerning the care of the airplane, it is important to include the airplane serial number in any correspondence to Hawker Beechcraft Corporation.

B. Exterior Placards and Markings (UA-1 and After; UB-1 and After) Required interior placards and markings are listed in Section II Limitations of the Pilot’s Operating Handbook and FAA Approved Airplane Flight Manual. Exterior placards and markings, with their locations indicated, are shown in this chapter (Ref. Figure 1). NOTE: Any time an airplane is repainted or touched up, inspect all placards and markings to ensure that they are securely attached, are legible, and are not covered with paint.

11-20-00

Page 1 Nov 1/09

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 1 (Sheet 1 of 13) Exterior Placards (UA-1 and After; UB-1 and After)

Page 2 Nov 1/09

11-20-00

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Figure 1 (Sheet 2 of 13) Exterior Placards (UA-1 and After; UB-1 and After)

11-20-00

Page 3 Nov 1/09

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Figure 1 (Sheet 3 of 13) Exterior Placards (UA-1 and After; UB-1 and After)

Page 4 Nov 1/09

11-20-00

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Figure 1 (Sheet 4 of 13) Exterior Placards (UA-1 and After; UB-1 and After)

11-20-00

Page 5 Nov 1/09

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Figure 1 (Sheet 5 of 13) Exterior Placards (UA-1 and After; UB-1 and After)

Page 6 Nov 1/09

11-20-00

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Figure 1 (Sheet 6 of 13) Exterior Placards (UA-1 and After; UB-1 and After)

11-20-00

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Figure 1 (Sheet 7 of 13) Exterior Placards (UA-1 and After; UB-1 and After)

Page 8 Nov 1/09

11-20-00

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Figure 1 (Sheet 8 of 13) Exterior Placards (UA-1 and After; UB-1 and After)

11-20-00

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Figure 1 (Sheet 9 of 13) Exterior Placards (UA-1 and After; UB-1 and After)

Page 10 Nov 1/09

11-20-00

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

OIL AIR STRUT PART NO. 114-820021BEECH AIRCRAFT CORPORATION WICHITA, KANSAS USA

INSTRUCTIONS FILL WITH MIL-H-5606 HYDRAULIC FLUID TO CHECK FLUID AND FILL REMOVE VALVE CAP, DEPRESS VALVE CORE AND ALLOW STRUT TO FULLY COMPRESS. REMOVE VALVE CORE, CONNECT ONE END OF A 1/4 INCH HOSE TO VALVE STEM AND SUBMERGE OTHER END IN HYDRAULIC FLUID. SLOWLY EXTEND STRUT TO FILL, THEN COMPRESS STRUT TO EXPELL EXCESS FLUID. RECYCLE AS NECESSARY TO EXPELL ALL AIR. WITH STRUT COMPRESSED, REPLACE VALVE CORE. WITH AIRCRAFT EMPTY EXCEPT FOR FULL FUEL AND OIL KEEP STRUT INFLATED TO 5.25 TO 5.75 INCHES OF PISTON SHOWING.

WARNING RELEASE AIR IN STRUT BEFORE DISASSEMBLING

PLACARD 29 UC11B 972956AA.AI

Figure 1 (Sheet 10 of 13) Exterior Placards (UA-1 and After; UB-1 and After)

11-20-00

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Figure 1 (Sheet 11 of 13) Exterior Placards (UA-1 and After; UB-1 and After)

Page 12 Nov 1/09

11-20-00

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 1 (Sheet 12 of 13) Exterior Placards (UA-1 and After; UB-1 and After)

11-20-00

Page 13 Nov 1/09

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Figure 1 (Sheet 13 of 13) Exterior Placards (UA-1 and After; UB-1 and After)

Page 14 Nov 1/09

11-20-00

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PLACARDS AND MARKINGS EXTERIOR PLACARDS AND MARKINGS (UC-1 AND AFTER) DESCRIPTION AND OPERATION

11-21-00 00

1. GENERAL A. Model Designation Placard (UC-1 and After) The model designation placard is located on the left side of the airplane on the forward cabin entrance doorframe. The door must be open to observe the placard. The placard identifies the airplane by its model number and serial number. Should a question arise concerning the care of the airplane, it is important to include the airplane serial number in any correspondence to Hawker Beechcraft Corporation.

B. Exterior Placards and Markings (UC-1 and After) Required interior placards and markings are listed in Section II Limitations of the Pilot's Operating Handbook and FAA Approved Airplane Flight Manual. Exterior placards and markings, with their locations indicated, are shown in this Chapter (Ref. Figure 1). Figure 1 illustrates exterior markings around exits. The external marking of each exit should include a 2-inch wide band of contrasting color that is readily distinguishable from the surrounding fuselage surface outlining the exit. Refer to the appropriate Airworthiness regulations for more details. NOTE: Any time an airplane is repainted or touched-up, inspect all placards and markings to ensure that they are securely attached, legible, and not covered with paint.

11-21-00

Page 1 May 1/10

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 1 (Sheet 1 of 12) Exterior Placards (UC-1 and After)

Page 2 May 1/10

11-21-00

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 1 (Sheet 2 of 12) Exterior Placards (UC-1 and After)

11-21-00

Page 3 May 1/10

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 1 (Sheet 3 of 12) Exterior Placards (UC-1 and After)

Page 4 May 1/10

11-21-00

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 1 (Sheet 4 of 12) Exterior Placards (UC-1 and After)

11-21-00

Page 5 May 1/10

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 1 (Sheet 5 of 12) Exterior Placards (UC-1 and After)

Page 6 May 1/10

11-21-00

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 1 (Sheet 6 of 12) Exterior Placards (UC-1 and After)

11-21-00

Page 7 May 1/10

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Figure 1 (Sheet 7 of 12) Exterior Placards (UC-1 and After)

Page 8 May 1/10

11-21-00

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Figure 1 (Sheet 8 of 12) Exterior Placards (UC-1 and After)

11-21-00

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Figure 1 (Sheet 9 of 12) Exterior Placards (UC-1 and After)

Page 10 May 1/10

11-21-00

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Figure 1 (Sheet 10 of 12) Exterior Placards (UC-1 and After)

11-21-00

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Figure 1 (Sheet 11 of 12) Exterior Placards (UC-1 and After)

Page 12 May 1/10

11-21-00

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Figure 1 (Sheet 12 of 12) Exterior Placards (UC-1 and After)

11-21-00

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11-30-00 200200

PLACARDS AND MARKINGS FLIGHT CONTROL RIG PIN FUSELAGE AND EMPENNAGE PLACARDS MAINTENANCE PRACTICES 1. GENERAL WARNING: When installing new placards, rig pin must be properly installed in the hole to verify that the correct hole is identified. The flight control rig pin fuselage and empennage placards are located near the applicable bellcrank in the fuselage and empennage. These placards identify the location to install the applicable rig pin for the flight control system being worked. These placards were installed per Field Service Kit 114-4062-0001. Figures 201 through 205 illustrate the approximate location of these placards.

A. Fuselage Rig Pin Placards The fuselage rig pin placards are located in four separate locations of the fuselage. Two placards are in the flight compartment area, one placard is located in the main fuselage area and one placard is located in the aft fuselage area. The two placards located in the flight compartment area are for the rudder forward bellcrank, located on the left side of the flight compartment forward of the pilot’s seat (rudder gust lock hole), and the elevator forward bellcrank located on the left side of the flight compartment forward and right of the pilot’s seat under panel 22 (Ref. Chapter 6-50-00, FLOOR ACCESS PANELS). The placard for the aileron system is located in the main fuselage area just aft of the main spar under panel 16E (Ref. Chapter 6-50-00, FLOOR ACCESS PANELS). The placard in the aft fuselage area is located on the right side of the aft rudder torque tube. Access to this location is behind panel 8 (Ref. Chapter 6-50-00, FUSELAGE ACCESS PANELS).

B. Empennage Rig Pin Placard The empennage rig pin placards are for the aft elevator bellcrank located at the top aft section of the vertical stabilizer under panel 34 (Ref. Chapter 6-50-00, STABILIZER ACCESS PANELS). One placard is installed on the left side and one on the right side of the vertical stabilizer.

11-30-00

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1. RIG PIN HOLE 2. AILERON QUADRANT SUPPORT BRACKET 3. AILERON QUADRANT 4. PLACARD

A

4

RIG PIN 3

1 DETAIL

B

2

B

LE FT

AILERON QUADRANT

DETAIL

A UC11B 052024AA.AI

Figure 201 Aileron Rig Pin Placard (Main Fuselage Area)

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11-30-00

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A

1. FORWARD RUDDER BELLCRANK 2. RIG PIN HOLE 3. PLACARD (2 PIECES) 3

RIG DETAIL

PIN

B

1

B

2

DETAIL

A

UE11B 052025AA.AI

Figure 202 Forward Rudder Bellcrank Rig Pin Placard (Left Side Flight Compartment)

11-30-00

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1. FORWARD ELEVATOR BELLCRANK RIG PIN HOLE 2. FORWARD ELEVATOR BELLCRANK (UNDER STRUCTURE) 3. PLACARD

A

3

RIG PIN DETAIL

B

2

1

B

VIEW LOOKING DOWN LEFT SIDE OF PEDESTAL DETAIL

A

Figure 203 Elevator Forward Bellcrank Rig Pin Placard (Left Side Flight Compartment)

Page 204 Nov 1/09

11-30-00

UC11B 052026AA.AI

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

1. AFT ELEVATOR BELLCRANK 2. AFT ELEVATOR RIG PIN HOLE 3. PLACARD

A 3

RIG PIN 1 DETAIL

B

B

FWD

2

VERTICAL STABILIZER (REF)

LEFT SIDE SHOWN; RIGHT SIDE OPPOSITE DETAIL

A

UC11B 052027AA.AI

Figure 204 Elevator Aft Bellcrank Rig Pin Placard (Empennage Area)

11-30-00

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A

1. AFT RUDDER TORQUE TUBE 2. RIG PIN HOLE 3. PLACARD 3

RIG PIN DETAIL

B NOTE: EARLIER VERSIONS OF THE TORQUE TUBE SECTOR MAY NOT HAVE LIGHTENING HOLES INSTALLED.

1 2

B

VIEW

A UC11B 052028AA.AI

Figure 205 Aft Rudder Torque Tube Rig Pin Placard (Aft Fuselage Area)

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CHAPTER 12 - SERVICING TABLE OF CONTENTS SUBJECT

PAGE

GENERAL INFORMATION 12-00-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1

REPLENISHING 12-10-00 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Fuel System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Water and Foreign Material Contamination . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Fuel Contamination Control . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Fuel-Handling Safety Information . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Fuel Grades and Types . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Fuel Tank Filling (UA-1 and After; UB-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Fuel Tank Filling (UC-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Draining the Fuel System (UA-1 and After; UB-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Draining the Fuel System (UC-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Draining The Fuel System (Alternate Method) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Fuel Additives . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Anti-Icing Additive, MIL-I-27686 or MIL-I-85470A . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Biocidal Agent . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Oil System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tires . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Hydraulic System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Brake System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Antiskid Brake System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Air Cycle Machine . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Oxygen System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 201 201 202 203 204 205 205 205 206 209 209 209 211 212 212 212 213 213 213

SCHEDULED SERVICING 12-20-00 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Engine External Washing Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Salt Water Contamination . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Fire Extinguisher Agent Contamination . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Cleaning Airplane Exteriors . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Cleaning During Curing Period (One Month) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Cleaning After Curing Period . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Environmental Fallout (Acid Rain) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Waxing Airplane Finishes . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Surface Deicer Boots . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Landing Gear . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Placard Replacement . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Plastic Windows . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Windshields . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Cleaning Airplane Interiors . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Upholstery . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Interior Cabin Trim . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Lubrication . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

12-CONTENTS

201 201 201 201 201 201 202 202 202 202 202 202 202 203 203 204 204 204 205

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

CHAPTER 12 - SERVICING TABLE OF CONTENTS (CONTINUED) SUBJECT PAGE Sealed Bearings . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 205 Wheel Bearing Lubrication . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 205 Spline Drives . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 205 Gaskets and Packings . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 205 Control Cables and Cable Pressure Seals . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 205 Lubrication of Threads . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 206 Lubrication Schedule . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 206 Nacelle Engine Controls and Inertial Anti-ice Lubrication Figure 202 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 212 Flight Compartment Engine Controls and Propeller Lubrication Figure 203 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 214 Nose Landing Gear Lubrication Figure 204 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 216 Main Landing Gear Lubrication Figure 205 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 218 Flight Compartment and Elevator Controls Lubrication Figure 206 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 220 Rudder Control System Lubrication Figure 207 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 222 Flap and Aileron Control System Lubrication Figure 208 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 224 Nose Avionics Door Lubrication Figure 209 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 226 Cabin Door Lubrication Figure 210 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 228 Cargo Door Lubrication Figure 211 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 230

UNSCHEDULED SERVICING 12-30-00 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Information . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . De-icing and Anti-icing of Aiplanes on the Ground . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . De-icing and Anti-icing Fluids . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Remove Frost, Snow or Ice . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Guidelines to Holdover Times (HOT) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General Application . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Remove Frost . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Remove Sleet and Freezing Rain . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Remove Snow . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . De-ice the Windshield . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . De-ice the Landing Gear . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Anti-icing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Fluid Spills . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Remove Salt or Chemical Agents . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Residue fron De-ice/Anti-ice . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Fire Extinguisher Agent Contamination . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Biocidal Agent Treatment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

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List of Effective Pages CH-SE-SU

PAGE

DATE

12-LOEP

1

Nov 1/13

12-CONTENTS

1 and 2

Feb 1/10

12-00-00

1

Nov 1/09

12-10-00

201 thru 214

May 1/10

12-20-00

201 thru 232

May 1/11

12-30-00

201 thru 207

Feb 1/10

12-LOEP

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

SERVICING GENERAL INFORMATION DESCRIPTION AND OPERATION

12-00-00 00

1. GENERAL WARNING: WHEN AN AIRPLANE HAS EXPERIENCED ABNORMAL LANDING GEAR PROCEDURES OF ANY TYPE, AS A SAFETY PRECAUTION, PLACE THE AIRPLANE ON JACKS PRIOR TO PERFORMING ANY INSPECTION OR MAINTENANCE. ENSURE THAT ALL THREE LANDING GEARS ARE DOWN AND LOCKED PRIOR TO REMOVING THE AIRPLANE FROM THE JACKS. CAUTION: Jacking of an airplane for the purpose of landing gear operation, inspection, servicing or maintenance, should be accomplished within an enclosed building or hangar. In the interest of safety, should it become necessary to jack the airplane in the open, wind velocity in any direction and terrain variations must be compensated for prior to jacking the airplane. Servicing information contained in this chapter is limited to the types of servicing that are of a general nature concerned with servicing of the overall airplane. Servicing procedures for specific components of the airplane are in the chapter applicable to the component. Section 12-10-00 covers information pertinent to the replenishing of fuel, oil, hydraulic fluid, tire pressures, etc. Section 12-20-00 contains information concerning lubrication of components and cleaning of the airplane exterior and interior parts. These servicing procedures are normally performed according to time schedules. Section 12-30-00 contains information pertaining to servicing of an unscheduled nature such as the removal of ice and snow. Servicing time limits for parts or components that must be serviced according to a specific time schedule will be found in Chapter 5, TIME LIMITS - MAINTENANCE CHECKS. Lubrication necessary for the performance of maintenance procedures, such as packing of gearboxes or lubrication of spline drives, will be covered in the chapter applicable to the system or component being maintained.

12-00-00

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

SERVICING REPLENISHING MAINTENANCE PRACTICES

12-10-00 200200

1. FUEL SYSTEM A. Water and Foreign Material Contamination All hydrocarbon fuels contain some dissolved and some suspended water. The quantity of water contained in the fuel depends on temperature and the type of fuel. Kerosene tends to absorb and suspend more water than aviation gasoline. Along with the water, it will suspend rust, lint and other foreign materials longer. Given sufficient time, these suspended contaminants will settle to the bottom of the tank. However, the settling time for kerosene is five times that of aviation gasoline. Due to this fact, jet fuels require good fuel handling practices to assure that the aircraft is serviced with clean fuel. If recommended ground procedures are carefully followed, solid contaminants will settle and free water can be reduced to 30 parts per million (ppm), a value that is currently accepted by the major airlines. Since most suspended matter can be removed from the fuel by sufficient settling time and proper filtration, it is not a major problem. Dissolved water has been found to be the major fuel contamination problem. Its effects are multiplied in airplanes operating primarily in humid regions and warm climates. Dissolved water cannot be filtered from the fuel by a micronic type filter, but can be released by lowering the fuel temperature, such as will occur in flight. For example, a kerosene fuel may contain 65 ppm (8 ounces per 1000 gallons) of dissolved water at 80°F. When the fuel temperature is lowered to 15°F, only about 25 ppm will remain in solution. The difference of 40 ppm will have been released as super-cooled water droplets which need only a piece of solid contaminant or an impact shock to convert them to ice crystals. Tests indicate that these water droplets will not settle during flight and are pumped freely through the system. If they become ice crystals in the tank, they will not settle since the specific gravity of ice is approximately equal to that of kerosene. Forty (40) ppm of suspended water seems like a very small quantity, but when added to suspended water in the fuel at the time of delivery, it is sufficient to ice a filter. While the critical fuel temperature range is from 0° to -20°F, which produces severe system icing, water droplets can freeze at any temperature below 32°F. Water in jet fuel also creates an environment favorable to the growth of a microbiological sludge in the settlement areas of the fuel cells. This sludge, plus other contaminants in the fuel, can cause corrosion of metal parts in the fuel system as well as clogging the fuel filters. Even though the Model 1900 Airliner utilizes the latest corrosion resistant materials and techniques, the possibility of filter clogging and corrosive attacks on various fuel system components may occur if contaminated fuels are introduced. Since fuel temperature and settling time affect total water content and foreign matter suspension, contamination can be minimized by keeping equipment clean, using adequate filtration equipment and careful water drainage procedures, storing the fuel in the coolest areas possible, and allowing adequate settling time. Underground storage is recommended for fuels. Filtering the fuel each time it is transferred will minimize the quantity of suspended contaminants carried by the fuel.

B. Fuel Contamination Control The primary means of fuel contamination control by the owner/operator is good housekeeping. This applies not only to fuel supply, but to keeping the airplane system clean. The following is a list of Steps that may be taken to recognize and prevent contamination problems. (1) Know your supplier. It is impractical to assume that fuel free from contaminants will always be available, but it is feasible to exercise precaution and be watchful for signs of fuel contamination.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (2) Assure, as much as possible, that the fuel obtained has been properly stored, that it is filtered as it is pumped to the truck, and again as it is pumped from the truck to the airplane. (3) Perform filter inspections to determine if sludge is present. (4) Maintain good housekeeping by periodically flushing the fuel tankage system. The frequency of flushing will be determined by the climate and the presence of sludge. (5) Use only clean fuel servicing equipment. (6) After refueling, allow a three-hour settling period whenever possible, then drain a small amount of fuel from each drain.

C. Fuel-Handling Safety Information Fuel handling and maintenance on or near the fuel system when flammable fuel vapors are present must be performed as cautiously as possible to prevent a fire or an explosion. The following safety information must be complied with when handling fuel or during maintenance of the airplane fuel system. WARNING: The airplane and all equipment used in fueling or defueling must be properly grounded to each other and to the ramp. This includes defueling equipment, work stands, purging equipment, and any powered or pneumatic devices. Equip work stands with a personnel static discharge plate of copper or zinc, affixed in such a position that personnel can contact the plate before coming in contact with the airplane. High static electrical charges are created by the contact and separation of unlike substances, or by any sort of motion of persons or material, and are a constant source of danger when generated in the presence of fuels or flammable vapors. Grounding jacks are located near the fuel filler cap on each wing and on each side of the fuselage nose section. Do not drain fuel tanks near the end of the working day and allow them to stand empty overnight. It could make conducive conditions for producing explosive vapors. If the system is not completely empty, residual fuel drains down the sides of the tank and forms puddles. During the night, fuel from the puddles evaporates into the air in the tank and a critical fuel-air ratio develops. An explosion could be set off by a spark. Avoid such a lapse of time between draining and purging of the fuel tanks. Ensure that the area is well ventilated before draining fuel. No smoking within 50 feet of the airplane or any place where flammable fuel vapors are present. Place the battery and generator switches in the OFF position. Disconnect all electrical power from the airplane before fueling. If the fuel contacts the eyes, rinse with cool, fresh water and seek medical attention immediately. Avoid allowing fuel to contact the skin. When contact cannot be avoided, wash with mild soap and water. Visually inspect all connections and hoses for leaks. If any leaks are indicated or develop, discontinue servicing.

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12-10-00

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

CAUTION: Remove spilled fuel immediately to prevent the fuel-contaminated surface from causing deicer boot and/or tire deterioration. Do not fill the auxiliary fuel tanks until the main tanks are full. Avoid damage to the deicer boots with refueling equipment or fuel spillage. The deicer boots are made of a soft, flexible material which may be damaged if refueling hoses, ladders or platforms are dragged across or rested against the deicer boot surfaces. Do not rest the fuel nozzle in the filler neck of the tanks.

D. Fuel Grades and Types Table 201 gives fuel refiner's brand name, along with the corresponding designations established by the American Petroleum Institute (API) and the American Society of Testing Material (ASTM). The brand names are listed for ready reference and are not specifically recommended by Hawker Beechcraft Corporation. Any product conforming to the recommended specification may be used. Jet A, Jet A-1, Jet B, JP-4, JP-5, and JP-8 fuels may be mixed in any ratio. Aviation gasoline, grades 80 Red (formerly 80/87), 91/98, 100LL Blue, 100 Green (formerly 100/130), and 115/145 Purple are emergency fuels and may be mixed in any ratio with the normal fuels when necessary. However, use of the lowest octane rating available is suggested due to its lower lead content. NOTE: In some countries, 100LL Blue is colored Green and designated 100L. CAUTION: The use of aviation gasoline shall be limited to 150 hours operation during each Time Between Overhaul (TBO) period. The use of gasoline as a jet fuel should be minimized wherever possible due to adverse effects to the hot section parts and the corrosion of turbine vanes. Table 201 Fuel Brands and Type Designations COMPANY

PRODUCT NAME

DESIGNATION

AMERICAN OIL COMPANY

American Jet Fuel Type A American Jet Fuel Type A-1

Jet A Jet A-1

ATLANTIC REFINING COMPANY

Arcojet-A Arcojet-A-1 Arcojet-B

Jet-A Jet-A-1 Jet-B

BP TRADING COMPANY

BP A.T.K BP A.T.G.

Jet A-1 Jet B

CALIFORNIA TEXAS COMPANY

Caltex Jet A-1 Caltex Jet B

Jet A-1 Jet B

CITIES SERVICE COMPANY

Turbine Type A

Jet A

CONTINENTAL OIL COMPANY

Conoco Jet-40 Conoco Jet-50 Conoco Jet-60 Conoco Jet JP-4

Jet A Jet A Jet A-1 Jet B

GULF OIL COMPANY

Gulf Jet A Gulf Jet A-1 Gulf Jet B

Jet A Jet A-1 Jet B

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Table 201 Fuel Brands and Type Designations (Continued) COMPANY

PRODUCT NAME

DESIGNATION

EXXON OIL COMPANY

Exxon Turbo Fuel A Exxon Turbo Fuel 1-A Exxon Turbo Fuel 4

Jet A Jet A-1 Jet B

MOBIL OIL COMPANY

Mobil Jet A Mobil Jet A-1 Mobil Jet B

Jet A Jet A-1 Jet B

PHILLIPS PETROLEUM COMPANY

Philjet A-50 Philjet JP-4

Jet A Jet B

PURE OIL COMPANY

Purejet Turbine Fuel Type A Purejet Turbine Fuel Type A-1

Jet A Jet A-1

RICHFIELD PETROLEUM COMPANY

Richfield Turbine Fuel A Richfield Turbine Fuel A-1

Jet-A Jet-A-1

SHELL OIL COMPANY

Aeroshell Turbine Fuel 640 Aeroshell Turbine Fuel 650 Aeroshell Turbine Fuel JP-4

Jet A Jet A-1 Jet B

SINCLAIR OIL COMPANY

Sinclair Superjet Fuel Sinclair Superjet Fuel

Jet A Jet A-1

STANDARD OIL OF CALIFORNIA

Chevron TF-1 Chevron JP-4

Jet A-1 Jet B

STANDARD OIL OF KENTUCKY

Standard JF A Standard JF A-1 Standard JF B

Jet A Jet A-1 Jet B

STANDARD OIL OF OHIO

Jet A Kerosene Jet A-1 Kerosene

Jet A Jet A-1

TEXACO

Texaco Avjet K-40 Texaco Avjet K-58 Texaco Avjet JP-4

Jet A Jet A-1 Jet B

UNION OIL COMPANY

76 Turbine Fuel Union JP-4

Jet A-1 Jet B

NOTE: Jet A - Aviation Kerosene Type fuel with -40°F (-40°C) Freeze Point. Jet A-1 - Aviation Kerosene Type Fuel with -58°F (-50°C) Freeze Point. Jet B - A low grade kerosene type fuel with a freeze point of -60°F (-51°C), similar to MIL-T-5624 grade JP-4, which has a freeze point of -76°F (-60°C).

E. Fuel Tank Filling (UA-1 and After; UB-1 and After) When filling the airplane fuel tanks, always observe the following: (1) Make sure the airplane is statically grounded to the servicing unit and to the ramp.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL CAUTION: Use a protective pad on the deicer boot. When inserting or removing the fuel nozzle, use extreme care to prevent the fuel hose from rubbing against the deicer boot. Also, do not allow fuel to contact the deicer boot. (2) The fuel filler caps are located on the outboard leading edge of each wing near the wing tip, providing single point refueling for each wing. (3) Allow a three-hour settle period whenever possible, then drain a small amount of fuel from each drain point.

F. Fuel Tank Filling (UC-1 and After) (1) Ground the airplane to the servicing unit and to the ramp. (2) The filler caps for the auxiliary tanks are located inboard of the nacelles and aft of the main spars. The main tank filler caps are located on the leading edges near the wing tips. Fill the main tanks first, then the auxiliary tanks. CAUTION: Do not take off with fuel in auxiliary tanks only, even for flights of short duration. Use a nonabrasive protective pad on the deicer boot to prevent contact with the fuel hose. Do not allow any fuel to come into contact with the deicer boot. (3) If possible, allow the fuel to settle for three hours and then drain a small amount of fuel from each drain point.

G. Draining the Fuel System (UA-1 and After; UB-1 and After) NOTE: Before beginning the defueling operation, always statically ground the airplane structure. Remove the fuel filler cap from the wing being defueled during draining to prevent damage or collapse of the fuel cells. (1) Provide a large suitable container or tank for the fuel to drain into. (2) Unscrew the drain valve from the bottom of the sump tank on the wing being defueled. The valve is located on the LH or RH inboard wing near the fuselage. (3) Screw in an AN815-12D union into the drain valve opening and connect a drain line. The fuel will gravity drain as the check valve is unseated by the union. (4) When fuel stops flowing from the fuel drain valve, open the sump drain valves to drain the residual fuel. Open the sump drain valves with a fuel sump drain wrench (Figure 1 (Sheet 2 of 10), Chapter 91-00-00) or with a screw driver by pushing up. Rotate the valve one half turn to lock it open. (5) After fuel has drained to desired level ensure fuel drain valves and sump drain valves are properly closed. Install plug on fuel drain valve (removed in Step (2)) and safety wire. (6) After refueling the affected tank ensure the fuel drain valves and sump drain valves (opened in Steps (3) and (4)) do not leak.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

H. Draining the Fuel System (UC-1 and After) NOTE: Before beginning the defueling operation, always statically ground the airplane structure. (1) Position an approved container large enough to hold the fuel that is to be drained under the drain valve, and open the filler cap on the tank being drained. NOTE: The drain valves for the main tanks are located immediately outboard of the lower nacelle structure. The drain valves for the auxiliary tanks are located under the auxiliary pump access covers. Each tank may be drained through its respective drain valve or the auxiliary tank fuel may be pumped into the main tank and all fuel drained through the main tank fuel drain. (2) Remove the plug from the drain valve and screw an AN815-12 union and drain line into the valve. Screwing the fitting into the valve will open the fuel check; thus, the drain line should be in the catch container before the fitting is screwed into the valve. (3) When fuel stops flowing from the fuel drain valve, open the sump drain valves to drain the residual fuel. Open the sump drain valves with a fuel sump drain wrench (Figure 1 (Sheet 2 of 10), Chapter 91-00-00) or with a screw driver by pushing up. Rotate the valve one half turn to lock it open. (4) After fuel has drained to desired level ensure fuel drain valves and sump drain valves are properly closed. Install plug on fuel drain valve (removed in Step (2)) and safety wire. (5) After refueling the affected tank ensure the fuel drain valves and sump drain valves (opened in Steps (2) and (3)) do not leak.

I. Draining The Fuel System (Alternate Method) WARNING: The fuel handling safety information under the heading FUEL HANDLING SAFETY INFORMATION in this section must be complied with. CAUTION: Do not fly the airplane with fuel in the auxiliary tanks only, even for flights of short duration. (1) Position an approved defueling container large enough to accommodate the amount of fuel to be removed from the airplane and statically ground the fuel container to the ramp. (2) Before beginning the defueling operation, statically ground the airplane structure to the defueling container and to the ramp. (3) Remove the lower accessory panel from the applicable nacelle. (4) Place the BATT switch to the ON position and pull the applicable FIRE PULL handle to close the firewall shutoff valve. (5) Place the BATT switch to the OFF position. (6) Disconnect firewall fuel hose (2) at engine driven fuel pump (1) and connect a drain line to the firewall fuel hose. Cap the engine driven fuel pump fitting to prevent contamination (Ref. Figure 201). (7) Apply ground power to the airplane (Ref. Chapter 24-40-00). (8) Push in the applicable FIRE PULL handle to open the firewall shutoff valve.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL NOTE: This procedure will also drain the auxiliary fuel tank by placing the AUX PUMP switch in the AUTO position. If desired, fuel may be left in the auxiliary fuel tank by placing the AUX PUMP switch to the OFF position. WARNING: Do not leave the airplane unattended while defueling. (9) Move the appropriate (LEFT or RIGHT) STANDBY PUMP toggle switch, located on the fuel control panel on the LH sidewall, to the ON position. WARNING: Do not allow the standby boost pump to run without fuel flowing for more than three minutes. (10) To prevent overheating of the standby boost pump, visually confirm fuel is flowing into the fuel container. (11) When the fuel flow begins to slow, place the applicable (LEFT or RIGHT) STANDBY PUMP toggle switch, located on the fuel control panel on the LH sidewall, to the OFF position. NOTE: The opposite wing may be defueled without relocating the equipment by using the fuel cross transfer system, although the rate of flow will be much slower. It would be much faster to move the equipment to the opposite wing and repeat the entire procedure. (12) If no further defueling is required proceed to Step (13). If the opposite wing is to be defueled using the cross transfer system, perform the following Steps: (a) Position the TRANSFER FLOW toggle switch, located on the fuel control panel, toward the wing with the defueling equipment connected. WARNING: Do not allow the standby boost pump to run without fuel flowing for more than three minutes. (b) To prevent overheating of the standby boost pump, visually confirm fuel is flowing into the fuel container. (c) When the fuel flow begins to slow, place the TRANSFER FLOW toggle switch to the OFF position. (13) Pull the applicable FIRE PULL handle to close the firewall shutoff valve. (14) Disconnect the drain line from the firewall fuel hose (2) (Ref. Figure 201). (15) Remove the cap from the engine driven fuel pump (1) and connect the firewall fuel hose (2). (16) Install the lower accessory panel on the applicable nacelle. (17) Push in the applicable FIRE PULL handle to open the firewall shutoff valve. (18) Remove ground power from the airplane (Ref. Chapter 24-40-00). (19) Use the sump drain valves to remove remaining fuel.

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1. ENGINE DRIVEN FUEL PUMP 2. FIREWALL FUEL HOSE

A A

1

2

VIEW

A

Figure 201 Alternate Defueling Connection

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UC12B 074153AA.AI

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

2. FUEL ADDITIVES All JP-4 jet fuel produced in the U.S.A. is required to contain an anti-icing additive conforming to MIL-I-27686 or MIL-I-85470A (62, Table 1, Chapter 91-00-00) in concentrations of 0.10 to 0.15 volume percent. However, all other fuels, including aircraft kerosene, contain no additives. Condensation of water in the fuel tanks increases the possibility of microbiological contamination of the fuel, which can damage skins, coatings and sealants in the fuel tanks. The biocidal agent (63, Table 1, Chapter 91-00-00) contains anti-fungicidal and biocidal agents. Use of this additive in the fuel tanks will reduce the possibility of contamination of the fuel and clogging of fuel filters and lines. These anti-icing additives and biocidal agents may be used separately or together in the fuel system with no detrimental effect on fuel system components.

A. Anti-Icing Additive, MIL-I-27686 or MIL-I-85470A MIL-I-27686 or MIL-I-85470A Jet Fuel Anti-Icing Additive (62, Table 1, Chapter 91-00-00) is primarily an anti-icing agent. It should be noted that anti-icing additive does not alter the freeze point of the fuel; however, when dissolved water separates from the fuel during a drop in temperature, the additive quickly separates from the fuel to preferentially dissolve in the water, thereby depressing the freezing point of the water to prevent the formation of ice in the fuel. The additive must be precisely blended into the fuel by a metering device that permits injection of the agent into a flowing stream of fuel to ensure even dispersal. Fuel distributors may tank or batch blend, or it may be preferred to blend at the airplane when fueling. If the tanker truck is not equipped with a HI FLO Prist blender (Model PHF-204), it may be necessary to carry the anti-icing and the blending device in the airplane. When blending the anti-icing agent with fuel, the concentration of additive should not be less than 20 fluid ounces per 156 gallons of fuel or more than 20 fluid ounces per 104 gallons of fuel. When adding previously blended fuel, the additive concentration should not be less than 0.10 percent by volume or more than 0.15 percent by volume. This additive should be used on a continuous basis. What biocidal/antifungal properties do MIL-I-27686 OR MIL-I85470 fuel additives have? Fuel additives conforming to MIL-I-27686 OR MIL-I-85470 do not have any specific referenced biocidal requirements. Prior to the mid 90’s MIL-I-27686, or ethylene glycol monomethyl (EGMME) was used. Some manufacturers (namely Prist) were able to market the anti-ice compound as microbiostat (not microbiocide, - stat means it controls or retards growth, - cide means it kills microbes) because they had their product certified as a pesticide. In the mid 90’s, the industry transferred from the use of MIL-I-27686 to MIL-I-85470 diethylene glycol monomethyl ether (DGMME). It is widely believed that DEGMME does have a retarding effect on microbial growth, but is not officially claimed by additive manufacturers. If any fungal or microbial growth is found in the fuel system, a biocidal agent should be used.

B. Biocidal Agent CAUTION: Drain water prior to refueling with biocidal agent. Excessive water concentrations in contact with excessive biocidal agent concentrations can result in formation of solid crystalline products in a fuel system. Biocidal Agent (63, Table 1, Chapter 91-00-00) is not an anti-icing agent and is intended to be used specifically as a fuel biocide. The compound is an extremely efficient biocidal agent and is soluble in fuel as well as water. Biocidal agent disperses throughout the entire fuel system to even the most remote areas soon after introduction into the system. The compound is used as a periodic preventive treatment in concentrations of 135 ppm when the airplane has been operated in an environment conducive to fungal or microbial contamination, or when such contamination is evidenced by dirty sump drains, clogged filters, odor, or visual evidence in the tanks, etc.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL CAUTION: Kathon FP 1.5 is approved for the initial shock of fuel tanks. Do not use Kathon FP 1.5 for preventative treatment. NOTE: Kathon FP 1.5 suggested use is once every six months as a shock treatment. Operating the airplane with a biocidal dosage at less than the recommended concentration may create and environment conducive to fungal and microbial contamination. Biobor JF in concentrations not to exceed 270 ppm, or Kathon FP 1.5 in a concentration of 100 ppm may be used as a single dose shock treatment to clean out and sterilize a very contaminated system. Any system that is contaminated should be treated at the concentration level for the appropriate biocide used. When sterility is achieved, the Biobor JF at the 135 ppm level may be used. Parked airplanes require only one treatment until fuel is burned off or replaced. The biocidal agent is not volatile and may remain in the tank until the fuel is used. The preferred method of introducing biocidal agents into the fuel is by injection through a metering device. If no metering device is available, blending may be accomplished by batch blending or by over-the-wing blending while filling the tanks. When half of the required quantity of fuel has been added, gradually introduce the compound directly into the stream of fuel while adding the other half of the fuel. Complete mixing is necessary, depending upon the severity of the contamination. Biobor JF or Kathon FP 1.5 must be used at a high enough concentration to kill, not just control the infestation. The mixture must be able to contact the entire surface of the fuel tank interior and remain in contact a minimum of 36 to 72 hours for Biobor JF or 12 to 72 hours for Kathon FP 1.5 to affect the kill. The longer the time the biocide remains the better the biocide will work. During the soak time the airplane should not be moved or the engines run. Refer to Table 202 or Table 203 for fuel ratios for specified concentrations as an aid in blending. Tank surfaces, gages, filters and linings should be inspected or replaced as necessary, depending upon the severity of the contamination. Be sure to account for residual fuel in the tanks so that proper dosage is maintained. Table 202 BIOBOR JF STERILIZATION AND MAINTENANCE TREATMENT LEVELS TURBINE FUEL

BIOBOR JF @ 270 PPM

BIBOR JF @ 135 PPM

LBS.

GALS.

LBS.

GALS.

FL. OZS.

LBS.

GALS.

FL. OZS.

670

100

0.18

0.02

2.63

0.09

0.01

1.32

1,340

200

0.36

0.04

5.26

0.18

0.02

2.63

2,010

300

0.54

0.06

7.89

0.27

0.03

3.95

2,680

400

0.72

0.08

10.53

0.36

0.04

5.26

3,350

500

0.90

0.10

13.16

0.45

0.05

6.58

6,700

1,000

1.81

0.21

26.46

0.90

0.10

13.16

13,400

2,000

3.62

0.41

52.92

1.81

0.21

26.46

16,750

2,500

4.52

0.52

66.08

2.26

0.26

33.04

33,500

5,000

9.01

1.03

132.16

4.52

0.52

66.08

67,000

10,000

18.09

2.07

264.47

9.05

1.03

132.31

NOTE: To estimate the fluid ounces of Biobor JF required to give a concentration of 270 ppm, multiply pounds of fuel by 0.004. For 135 ppm multiply pounds of fluid by 0.002.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Table 203 Kathon FP 1.5 Small Tank - Treatment Ratio @ 100 ppm System Volume

Dose Volume

Gallons

Liters

Cubic Meters

Ounces

Milliliters

Liters

50

189.25

0.19

0.75

22.18

0.022

100

378.50

0.38

1.5

44.36

0.044

150

567.75

0.57

2

59.14

0.059

200

757.00

0.76

2.75

81.32

0.081

250

946.25

0.95

3.5

103.50

0.103

300

1135.50

1.14

4

118.28

0.118

320

1211.20

1.21

4.25

125.67

0.126

350

1324.75

1.32

4.5

133.07

0.133

400

1514.00

1.51

5.25

155.24

0.155

450

1703.25

1.70

6

177.42

0.177

500

1892.50

1.89

6.5

192.21

0.192

1000

3785.00

3.79

13

384.41

0.384

1500

5677.50

5.68

19.5

576.62

0.577

2000

7570.00

7.57

25.75

761.44

0.761

2500

9462.50

9.46

32

946.25

0.946

3000

11,355.00

11.36

38.5

1138.46

1.138

4000

15,140.00

15.14

51.25

1515.48

1.515

500

18,925.00

18.93

64

1892.50

1.893

10,000

37,850.00

37.85

128

3785.00

3.785

NOTE: Density of Jet Fuel: 1 gallon weighs 6.714 pounds.

3. OIL SYSTEM Servicing the engine oil system involves maintaining the engine oil at the proper level. The engine oil tank is provided with a filler neck and a quantity dipstick and cap located at the 11 o'clock position on the accessory gear case. The dipstick is marked in U.S. quarts and indicates the amount required to fill the tank. Access to the oil dipstick cap is gained by opening the small access door in the upper aft cowling. CAUTION: Do not mix different brands of oil when adding oil between oil changes. Different brands of oil may be incompatible due to the difference in their chemical structure. The total oil tank capacity is 10 quarts. An additional 4.4 quarts of oil is required to fill the lines and cooler, giving a total system capacity of 14.4 quarts; however, because of the residual oil trapped in the system, no more than 13 quarts should be added during an oil change.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Check the oil level as follows: any time the oil has been changed or the engine has remained stationary for more than 12 hours: a) Run the engine at idle for 2 minutes. b) Shut down the engine and check the oil level.

4. TIRES The Model 1900 Series Airliner uses FOUR 22 x 6.75 - 10, 8-ply, 160 mph, tubeless type tires on the main landing gears and ONE 19.5 x 6.75 - 8, 10-ply, rib tread, 160-mph, tube type or tubeless tire on the nose landing gear. CAUTION: Tires that have picked up a fuel or oil film should be washed down as soon as possible with a detergent solution to prevent contamination of the rubber. Table 204 Tire pressure Main Landing Gear Tires (Loaded)

Main Landing Gear Tires (On Jacks)

Nose Landing Gear Tires

95 psi

91 psi

60 psi

Tire pressure should be checked on a regular basis. The intervals may be determined by individual experience utilizing operating and servicing requirements. Such a servicing program will help prevent tire damage from excessive pressure loss due to a slow leak or an extensive drop in temperature. Maintaining the proper tire inflation pressure shown in Table 204 will help to avoid damage from landing shock and contact with sharp stones and ruts, and will minimize tread wear. When inflating tires, inspect them for cuts, cracks, breaks and tread wear. The pressure of a serviceable tire that is fully inflated should not drop more than 4% over a 24-hour period. For the most accurate reading, tire pressure should be checked when the tires are cool; consequently, wait at least 2 hours (3 hours in hot weather) after a flight before checking tire pressure.

5. HYDRAULIC SYSTEM Servicing the hydraulic landing gear system consists of maintaining the correct fluid level. A fill can, located just inboard of the LH nacelle and forward of the main spar, contains a cap and dipstick assembly marked FILL WARM - COLD. Prior to removing the fill can lid, the knob on the manual bleed valve must be depressed to relieve air pressure in the power pack reservoir. Add hydraulic fluid (39, Table 1, Chapter 91-00-00) as required to fill the system. Approximately 2 3/4 gallons of hydraulic fluid is required to fill a completely empty system. NOTE: When filling the hydraulic system, the air being displaced by the hydraulic fluid will need to be relieved out the manual bleed valve. Occasionally depress the button on the manual bleed valve to relieve this air pressure and to allow the system to fill faster. This action will also allow the dipstick reading to more accurately indicate the amount of fluid in the system.

6. BRAKE SYSTEM Brake system servicing is limited to maintaining the hydraulic fluid level in the reservoir mounted in the upper LH corner of the aft bulkhead of the nose avionics compartment. When no fluid is visible in the reservoir sight glass, add a sufficient quantity of hydraulic fluid (39, Table 1, Chapter 91-00-00) to raise the fluid level to the

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL lower edge of the filler neck. Approximately 1 gallon of fluid is required to fill a completely empty system if power steering is installed. If power steering is not installed, approximately 1/2 gallon of fluid is required to fill a completely empty system.

7. ANTISKID BRAKE SYSTEM Brake system servicing is limited to maintaining the hydraulic fluid level in the reservoir mounted in the upper RH corner of the aft bulkhead of the nose avionics compartment. To check the fluid level, turn the master switch ON, turn the anti-skid switch OFF, ensure that the parking brake is off, remove the reservoir cap and depress the brake pedals 15 to 20 times to deplete the accumulator. If no fluid is visible in the reservoir sight glass, add a sufficient quantity of hydraulic fluid (39, Table 1, Chapter 91-00-00) to raise the fluid level to the lower edge of the filler neck.

8. AIR CYCLE MACHINE NOTE: Do not mix different types of oil when filling. Check the quantity of oil contained in the see-thru oil sump and add oil as required to maintain a full sump. The manufacturer's recommendation for lubricating oil to be used in the air cycle machine is Exxon 2389; however, any oil conforming to MIL-L-7808G (72, Table 1, Chapter 91-00-00) may be used. Approximately 1/2 pint of fluid is required to fill a completely empty air cycle machine.

9. OXYGEN SYSTEM WARNING: When filling the oxygen system, use only MIL-O-27210 Aviator's Breathing Oxygen (28, Table 1, Chapter 91-00-00). Do not use oxygen intended for medical purposes, or industrial purposes such as welding. Such oxygen may contain excessive moisture that could freeze in the valves and lines of the oxygen system. Access to the pressure gage and filler valve of the oxygen system may be gained through an access door located on the LH side of the nose section below the nose baggage compartment. To recharge the oxygen system, remove the protective cap from the filler valve and attach the hose from an oxygen recharging cart to the filler valve. Make sure that the airplane oxygen system and the servicing equipment are properly grounded before servicing the system. WARNING: Avoid making sparks and keep all burning cigarettes or fire away from the vicinity of the airplane. Make sure that the oxygen shutoff valve control (placarded OXYGEN PULL ON) located on the subpanel to the left of the copilot's seat is in the OFF position. Inspect the filler connection for cleanliness before attaching it to the filler valve. Make sure that your hands, tools, and clothing are clean, particularly of grease or oil, for these contaminants will ignite upon contact with pure oxygen under pressure. As a further precaution against fire, open and close all oxygen valves slowly. To prevent overheating, fill the oxygen system slowly by adjusting the recharging rate with the pressure regulating valve on the cart. At a temperature of 70°F., the dual 76.5-cubic foot cylinders should be filled to 1,850 psig. This pressure may be increased an additional 3.5 psig for each degree of increase in temperature; conversely, for each degree of drop in temperature, reduce the pressure for the cylinder(s) by 3.5 psig (Ref. Table 205). When the oxygen system is properly charged, disconnect the filler hose from the filler valve and replace the protective cap on the filler valve. If at any time, in the process of servicing and purging the system or replacing the oxygen cylinder, it becomes necessary to disconnect a fitting, the threads of the fitting should be wrapped with thread sealer (27, Table 1, Chapter 91-00-00) prior to being connected back into the system. NOTE: Refer to Advisory Circular 43.13-1A for the additional servicing precautions recommended by the FAA on the oxygen systems. For additional information (Ref. Chapter 35-00-00).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Table 205 Oxygen Charging Rate Temp (F)

PSI

Temp (F)

PSI

Temp (F)

PSI

Temp (F)

PSI

Temp (F)

PSI

0

1605

25

1692.5

50

1780

75

1867.5

100

1955

1

1608.5

26

1696

51

1783.5

76

1871

101

1958.5

2

1612

27

1699.5

52

1787

77

1874.5

102

1962

3

1615.5

28

1703

53

1790.5

78

1878

103

1965.5

4

1619

29

1706.5

54

1794

79

1881.5

104

1969

5

1622.5

30

1710

55

1797.5

80

1855

105

1972.5

6

1626

31

1713.5

56

1801

81

1888.5

106

1976

7

1629.5

32

1717

57

1804.5

82

1892

107

1979.5

8

1633

33

1720.5

58

1808

83

1895.5

108

1983

9

1636.5

34

1724

59

1811.5

84

1899

109

1986.5

10

1640

35

1727.5

60

1815

85

1902.5

110

1990

11

1643.5

36

1731

61

1818.5

86

1906

111

1993.5

12

1647

37

1734.5

62

1822

87

1909.5

112

1997

13

1650.5

38

1738

63

1825.5

88

1913

113

2000.5

14

1654

39

1741.5

64

1829

89

1916.5

114

2004

15

1657.5

40

1745

65

1832.5

90

1920

115

2007.5

16

1661

41

1748.5

66

1836

91

1923.5

116

2011

17

1664.5

42

1752

67

1839.5

92

1927

117

2014.5

18

1668

43

1755.5

68

1843

93

1930.5

118

1018

19

1671.5

44

1759

69

1846.5

94

1934

119

2021.5

20

1675

45

1762.5

70

1850

95

1937.5

120

2025

21

1678.5

46

1766

71

1853.5

96

1941

121

2028.5

22

1682

47

1769.5

72

1857

97

1944.5

122

2032

23

1685.5

48

1773

73

1860.5

98

1948

123

2035.5

24

1689

49

1776.5

74

1864

99

1951.5

124

2039

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

SERVICING SCHEDULED SERVICING MAINTENANCE PRACTICES

12-20-00 200200

1. GENERAL This chapter is supplemented by Chapter 20-10-01 of the 1900 Airliner Series Corrosion Control Manual.

2. ENGINE EXTERNAL WASHING PROCEDURES CAUTION: Never wash an engine while it is running or hot. After the engine has been shut down, allow it to cool for at least one hour prior to washing. Electrical components and plugs on the engine and in the engine compartment should be covered/protected during the engine wash. Use dry shop air to dry all components after washing. Do not allow water into the engine air inlet or the exhaust. Do not direct high pressure water or solvent directly into mechanical parts having air vent holes, such as the fuel control units.

A. Salt Water Contamination If the exterior surface of the engine is contaminated with salt, it should be washed thoroughly with water prior to the next flight of the airplane. Demineralized water is not required for this purpose. At no time should an engine be left in a contaminated (salted) condition for any extended period of time, such as overnight.

B. Fire Extinguisher Agent Contamination In the event of engine contamination by fire extinguishing agents, refer to the latest revision of PT6A-65 Engine Maintenance Manual P/N 3032842 or subsequent. CAUTION: If the engine inhales a fire extinguishing compound, it must be cleaned, removed, and disassembled for a thorough internal cleaning. If the engine fire extinguisher is discharged, the engine baffling will prevent entry of the extinguishing compound into the engine; thus only an external engine washing is required. Most incidences of fire extinguisher compound ingestion are the result of ground personnel using an external extinguisher during engine operation.

3. CLEANING AIRPLANE EXTERIORS CAUTION: Contamination or washout of grease in wheel bearings will damage bearings and may result in loss of the wheel. Prior to washing, attach the pitot cover securely and plug or mask off all other openings. Be particularly careful to mask off all static air ports before washing or waxing. Use special care to avoid washing away grease from any lubricated area. Prior to cleaning, cover such areas as wheels, brakes, etc. Always be sure all maskings and coverings are removed before returning the airplane to service. Lubricate after cleaning as necessary. The urethane finish undergoes a curing process for a period of time after application. During the first month after paint application, some special care is required. Airplane owners should observe the following recommendations in order to preserve the durability and appearance of the airplane paint.

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A. Cleaning During Curing Period (One Month) (1) Avoid prolonged flights in heavy rain or sleet. Avoid any operating conditions which might cause abrasion or premature finish deterioration. (2) Clean the airplane with mild detergents and water only. Use a clean soft rag, keeping it free from dirt and grime. Rinse with clear water thoroughly. (3) Use no waxes, polishes, rubbing compounds, or abrasive cleaners of any type. The use of such items can permanently damage the surface finish. (4) Stubborn oil or soot deposits on cowlings, wheel wells, etc. may be removed gently with automotive tar removers.

B. Cleaning After Curing Period (1) Wash the airplane regularly. Use mild detergents and water only. Rinsing thoroughly with clear water prevents detergent residue buildup that can dull the paint appearance. (2) Normally, waxing is not necessary; however if waxing is desired, select a high quality automotive or airplane waxing product. Never use rubbing compounds or abrasive cleaners of any type.

C. Environmental Fallout (Acid Rain) In certain areas of the country where chemicals may be present in the atmosphere, it is best to avoid outside storage when damp conditions exist. Acids which remain in standing water can stain the paint topcoat and cause permanent damage to the finish. Flush off residual moisture with clean tap water and dry the surface. At this time, waxing the surface can provide protection from acid rain damage.

D. Waxing Airplane Finishes A good coat of wax will protect the airplane finish from the sun's rays and protect the surface against oxidation. Use a high quality automotive or airplane wax. Do not use a wax containing silicone because silicone materials are difficult to remove.

E. Surface Deicer Boots The surfaces of the deicer boots should be checked for indications of engine oil after servicing and at the end of each flight. Any oil spots that are found should be removed with a nondetergent soap and water solution. Care should be taken when cleaning to avoid scrubbing the boots because the conductive coating (A56B) must not be removed from the boot surface. The boots are made of soft, flexible stock that can be damaged if gasoline hoses are dragged over the surface of the boots or if ladders or platforms are rested against them.

F. Landing Gear The landing gear (nose and main) should be washed with low pressure water and mild detergent as soon as is practical following operation on salty or muddy runways. Using low pressure air, blow off all water before flight or storage of the airplane.

G. Placard Replacement Ensure all placards are in place and legible whenever the airplane has been repainted or touched up after repairs. Replace any placards that have been defaced after such repainting or repairs. Page 202 May 1/11

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H. Plastic Windows The plastic windows should be kept clean and waxed at all times. To prevent scratches and crazing, wash the windows carefully with plenty of soap and running water. CAUTION: When washing the windshield, do not use water from a bucket or pail. Sand, dirt particles or other debris may collect in the standing water and cause scratches in the plastic. Use the palm of the hand to feel and dislodge dirt and mud. A soft cloth, chamois or sponge may be used only for the purpose of carrying water to the surface of the window. After washing, rinse the window thoroughly with running water and dry it with a clean, moist chamois. Do not rub the plastic window with a dry cloth, because this will cause an electrostatic charge which attracts dust. Remove oil and grease with a cloth moistened with kerosene (49, Table 1, Chapter 91-00-00), solvent (54, Table 1, Chapter 91-00-00) or hexane (51, Table 1, Chapter 91-00-00), then rinse the window with clear water. CAUTION: Never use gasoline, benzene, alcohol, acetone, carbon tetrachloride, fire extinguisher or anti-ice fluid, lacquer thinner, or glass cleaner with a base of these materials, for such materials will soften the plastic and may cause crazing. Aliphatic naptha and similar solvents are highly flammable and extreme care must be exercised when using these chemicals. If it is desirable to use a commercial cleaner to clean the plastic windows, use only cleaners that are approved by Hawker Beechcraft Corporation. There are several cleaners available commercially that state that they are approved for use on acrylic surfaces. However, it has been discovered that some of these cleaners cause acrylic plastic to craze. Therefore, only the following product is approved as a cleaner for acrylic plastic windows: plexiglas polish and cleaner (48, Table 1, Chapter 91-00-00). Follow the directions on the container. After washing plastic windows with soap and water, apply a good grade of commercial wax. The wax will fill in minor scratches and help prevent further scratches. Apply a thin, even coat of wax and bring it to a high polish by rubbing lightly with a clean, dry, soft flannel cloth. Never use a power buffer, as the heat generated by the buffing pad may soften the plastic. If the windows were cleaned with one of the commercial cleaners mentioned previously, it will not be necessary to apply wax. Each of these cleaners contains wax, as well as cleaning agents.

I. Windshields Glass windshields with antistatic coating should be cleaned as follows: (1) Wash excessive dirt and other substances from the glass with clean water. (2) Clean the windshield with mild soap and water or with a 50/50 solution of solvent (30, Table 1, Chapter 91-00-00) and water. Wipe the glass surface in a straight rubbing motion with a soft cloth or sponge. Never use any abrasive materials or any strong acids or bases to clean the glass. (3) Rinse the glass thoroughly and dry, but do not apply wax. NOTE: It is essential that the windshield wipers be thoroughly cleaned. Operating the wipers when they are dirty is a common source of scratches on the windshield. Do not attempt to polish out such nicks or scratches in the glass surface.

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4. CLEANING AIRPLANE INTERIORS The seats, rug, upholstery panels and headlining should be vacuum-cleaned frequently to remove as much surface dust as possible. Experience has shown that commercial, foam-type cleaners or shampoos can be used to condition the surfaces of rugs, carpets and upholstered materials. The upholstery should be vacuum-cleaned, and the stains should be removed. A solution of the cleaner can be prepared by mixing a small amount in a bucket of water and beating the solution until a heavy foam forms. Apply the foam uniformly with a brush over the surface to be cleaned, then remove the suds with a vacuum cleaner or by wiping the surface with a brush or cloth. Because there is very little moisture in this foam, wetting of the fabric or retention of moisture in the warp does not occur. Unlacquered metal fittings and furnishings within the airplanes can be cleaned using most commercial metal polishes. Use a soft, clean rag for application; then polish to a brilliant gloss with a dry cloth. Protect the finish with a good grade of wax.

A. Upholstery The most effective method of cleaning upholstery is directly dependent on the type of upholstery involved. For instance, a fabric type of upholstery that has been flame-proofed should never be treated by the application of cleaners with a water base. The reason for this is that the flame retardant on the fabric is water soluble and will be diluted to a point where the fire-resistant quality is rendered useless. Also, the natural capillary action of the water in the fabric will cause the salts of the flame-retardant chemicals to rise to the surface resulting in unsightly faded spots. NOTE: Clean wool and wool-blended upholstery fabrics by dry-cleaning ONLY. After 5 dry cleanings, have the upholstery treated with a fire retardant at a service company or spray the fabric with a fire retardant (105, Table 1, Chapter 91-00-00) or equivalent. Clean the fabric upholstery manually as follows: (1) Remove the upholstery from the airplane. (2) Use a stiff-bristled brush and brush the upholstery along the weave. (A nylon-bristled fingernail brush can be used). (3) Vacuum the entire surface to remove any salt residue or dirt stains. (4) Apply dry cleaning solvent (2, Table 1, Chapter 91-00-00) sparingly on a lint-free cloth and clean stains as required. (5) Allow the upholstery to completely dry. (6) Treat the upholstery with a fire retardant (105, Table 1, Chapter 91-00-00). (7) Install the upholstery. Clean leather upholstery with a nonabrasive, chemically neutral, nonreactive, emulsion-type cleaner such as saddle soap. Apply it over the dirty surface using a sponge or soft cloth (use a gentle, wiping motion; do not scrub). Do not allow the cleaner to dry on the material surface. Wipe the cleaner off before it drys. Treat leather after cleaning with wax or a leather conditioner.

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B. Interior Cabin Trim CAUTION: To prevent damage to the plastic interior trim, never use MPK, naptha, mufti, stoddard solvent, gasoline, lacquer thinner, or other types of paint cleaners as cleaning agents. Using soap and water, wash the plastic interior trim. Scrubbing with a brush and detergent soap will usually provide adequate cleansing; however, alcohol may be used to remove contaminants that are soluble in alcohol.

5. LUBRICATION A. Sealed Bearings Sealed bearings are prepacked with grease and do not require periodic lubrication. Sealed bearings must be replaced when normal airplane inspection procedures indicate that the bearing will no longer operate satisfactorily. The lubrication of sealed bearings must not be attempted unless facilities are available for removing and replacing seals. When sealed bearings are cleaned and lubricated, the work must be done in strict compliance with applicable bearing maintenance directives.

B. Wheel Bearing Lubrication CAUTION: Improper axle nut installation, mixing of lubricants, contamination or washout of grease in wheel bearings will damage bearings and may result in loss of the wheel. DO NOT MIX lubricants of different types or manufacturers. If the lubricant is changed or unknown, make certain that all the affected components are thoroughly cleaned before lubrication. Wheel bearing grease lubrication intervals vary significantly depending upon the operation of the individual airplane. Grease change intervals at tire change is acceptable if the service history has been satisfactory. If the service history is unknown or if the history dictates shorter intervals, Operators should change grease at 200 hour intervals or at tire change (which ever comes first) unless a longer interval can be shown to be acceptable. Exercise care when washing area. Contamination or washout of grease in the wheel bearing will damage the bearing and may result in loss of the wheel. When performing wheel bearing lubrication, inspect bearings, hub caps, and seals for condition. Replace any questionable parts. Perform the MAIN WHEEL INSTALLATION procedure (Ref. Chapter 32-40-00) or the NOSE WHEEL INSTALLATION procedure (Ref. Chapter 32-40-00). Ensure the axle nuts are properly torqued and safetied.

C. Spline Drives The engines have wet spline lubrication and lubrication of the splines is not required when installing an accessory. The air conditioning compressor end of the quill shaft requires lubrication. For lubrication of the quill shaft splines (Ref. Chapter 21-52-02).

D. Gaskets and Packings When lubricating gaskets and packings, use the type of fluid in the system for the gaskets and packings.

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E. Control Cables and Cable Pressure Seals Clean the pressure seals and the control cables for the length of travel through the pressure seals with cleaning solvent (2, Table 1, Chapter 91-00-00). Fill the seal with lubricant and lubricate the cleaned area of the cable and one inch beyond with grease (23, Table 1, Chapter 91-00-00). On all the remaining length of cable, apply corrosion-preventive compound (11, Table 1, Chapter 91-00-00) with a brush. Wipe off excess with a clean cloth.

F. Lubrication of Threads Lubricate all plumbing fittings with the proper lubricant (Ref. Figure 201 and Table 201). When applying lubricants, observe the following rules: (1) Clean the threads before applying the lubricant. (2) Use only thin coats of the selected thread lubricant. (3) Apply lubricant to the male threads only. (4) Do not lubricate the first two threads. (5) Never allow lubricant to enter fittings or flare areas. (6) On flared tube fittings, apply a small amount of lubricant on the back face of the sleeve shoulder. This is to prevent the sleeve from turning with the nut and galling the flare.

G. Lubrication Schedule The lubrication illustrations are organized so that related items requiring lubrication are grouped together. Each lubrication point listed in Table 202 is identified as indexed on the accompanying illustration. Lubrication time intervals are incremented to occur only at times coincident with the detailed inspection intervals (Ref. Chapter 05-20-00). Table 201 Thread Lubricants for Fluid-Line Fittings Type of Line

Brass, Steel and Aluminum Fittings

Functional Fluid

1. Fuel and Fuel Pressure Line

Fuel

Braycote 236 (VV-P-236)

2. Lubricating Oil and Oil Pressure Line

Lubricating Oil

Braycote 236 (VV-P-236) or MIL-G-6032 Lubricating Grease (Gasoline and Oil Resistant)

3. Automatic Pilot

Air (ambient)

a) Straight Threads

None

b) Tapered Threads

3M Tape #48 or #547 PTFE Tape (A-A-58092, Type III, Size 1 or 2)

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Table 201 Thread Lubricants for Fluid-Line Fittings (Continued) Type of Line 4. Pressurization Control

Brass, Steel and Aluminum Fittings

Functional Fluid Breathable Air (ambient)

a) Straight Threads

None

b) Tapered Threads

3M Tape #48 or #547 PTFE Tape (A-A-58092, Type III, Size 1 or 2)

5. Pitot System

Air (ambient)

a) Straight Threads

None

b) Tapered Threads

Loctite PST 59231 Pipe Sealant

6. Fire Extinguisher System

Trifluoro Bromo Methane

Loctite PST 59231 Pipe Sealant

7. Air Conditioning System

Freon (Possible Refrigeration Oil)

Oil of System

8. Oxygen System

Oxygen

a) Straight Threads

Krytox 240AC Grease (MIL-G-27617 Type III)

b) Tapered Threads

3M Tape #48 or #547 PTFE Tape (A-A-58092, Type III, Size 1 or 2)

9. Bleed Air System a) CRES (straight and tapered threads) b) AL (straight and tapered threads) 10. Hydraulic System

Air (Max. 750°F)

Dow Corning 77 (M-77)

Air

Loctite PST 59231 Pipe Sealant

Hydraulic Fluid

a) Straight Threads

Fluid of System

b) Tapered Threads

Fluid of System or Loctite 545

11. Vacuum

3M Tape #48 or #547 PTFE Tape (A-A-58092, Type III, Size 1 or 2) or Dow Corning 111 Valve & Lubricant Sealant

12. Deicer

Loctite PST 59231 Pipe Sealant

13. Gyro (Edo Air)

Air

3M Tape #48 or #547 PTFE Tape (A-A-58092, Type III, Size 1 or 2)

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Figure 201 Lubrication of Threads Table 202 Lubrication Schedule Lubrication Reference ENGINE CONTROLS (NACELLE)

Fig. 202

Cam Plate and Pins

Index 1

INERTIAL ANTI-ICE SYSTEM

Fig. 202

Hinge Point Bushings

Index 2

AIR-CONDITIONER COMPRESSOR

Fig. 202

Quill Shaft

Index 3

ENGINE CONTROLS (FLIGHT COMPARTMENT)

Fig. 203

Linkage Bushings and Pins

Index 1

PROPELLER

Fig. 203

Propeller Hub Grease Fittings

Index 2

Low Pitch Stop Rods

Index 3

NOSE LANDING GEAR

Fig. 204

Door Hinges and Retract Linkage

Index 2

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100 200 300 400 600 1200 Hours Hours Hours Hours Hours Hours

X

X

X

X

X X

X

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Table 202 Lubrication Schedule (Continued) Lubrication Reference

100 200 300 400 600 1200 Hours Hours Hours Hours Hours Hours

Grease Fittings

Index 3

Upper and Lower Nose Gear Strut Bearing

Index 4

Steering Bellcrank Grease Fitting

Index 5

X

Steering Disconnect Cam (Mechanical Steering Only)

Index 6

X

MAIN LANDING GEAR

Fig. 205

Grease Fittings

Index 2

Door Hinges and Linkage Bearings

Index 3

Door Retract Cam

Index 4

CONTROL COLUMN

Fig. 206

Chain

Index 1

RUDDER PEDALS

Fig. 206

Pedal

Index 2

ELEVATOR CONTROL SYSTEM

Fig. 206

Elevator Trim Tab Hinge

Index 3

Elevator Trim Tab Actuator Grease Fittings

Index 4

Elevator Trim Cable Pressure Seals

Index 5

X

Elevator Cable Pressure Seals

Index 6

X

Elevator Trim Tab Chain

Index 7

X

Elevator and Trim Tab Cables

Index 8

X

LANDING GEAR POWER PACK Fluid Level AIR CYCLE MACHINE

X X

X X X

X

X

X X

Chapter 12-10-00 Index 1

X

Chapter 12-10-00

Fluid Level

Index 1

X

RUDDER CONTROL SYSTEM

Fig. 207

Rudder Trim Tab Actuator Grease Fittings

Index 1

Rudder Trim Tab Hinge

Index 2

Rudder Tab Cable Pressure Seal

Index 3

X

Rudder Cable Pressure Seal

Index 4

X

Rudder and Trim Tab Cables

Index 5

X

X X

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Table 202 Lubrication Schedule (Continued) Lubrication Reference

100 200 300 400 600 1200 Hours Hours Hours Hours Hours Hours

FLAP CONTROL SYSTEM

Fig. 208

Flap Motor Gearbox

Index 1

Flap Tracks

Index 2

Flap Limit Switch Link

Index 3

X

Flap Asymmetry Switch Hub Assembly

Index 4

X

Flap Asymmetry switch Hub Assembly (Modified by Kit 129-5046)

Index 4A

X

AILERON CONTROL SYSTEM

Fig. 208

Aileron Bellcrank

Index 5

Bellcrank Rod Ends

Index 6

Trim Tab Actuator

Index 7

Trim Tab Cable Pressure Seals

Index 8

X

Aileron Cable Pressure Seals

Index 9

X

Aileron and Trim Tab Cables

Index 10

X

Aileron Trim Tab Hinge

Index 11

NOSE BAGGAGE DOOR

Fig. 209

Gas Spring End Fittings

Index 1

Door Hinge

Index 2

Latch Pin and Plate

Index 3

X

Latching Mechanism

Index 4

X

CABIN DOOR

Fig. 210

Cam Lip

Index 1

X

Door Damper

Index 2

X

Door Hinge

Index 3

X

Cam Surface of Pressure Lock

Index 4

Bearing Block Contact Surface or Grease Fittings

Index 5

X

Door Handle

Index 6

X

EMERGENCY EXIT DOORS

Fig. 210

Door Track

Index 7

X

Latching Mechanism

Index 8

X

CARGO DOOR

Fig. 211

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X X

X X X

X

X X

X

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Table 202 Lubrication Schedule (Continued) Lubrication Reference

100 200 300 400 600 1200 Hours Hours Hours Hours Hours Hours

Gas Spring End Fittings

Index 1

X

Door Cam Locks

Index 2

Pushrod Pin and Bushing

Index 3

X

Latch Pin and Latch Plate

Index 4

X

Door Hinge

Index 5

Door Handle

Index 6

X

X X

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H. Nacelle Engine Controls and Inertial Anti-ice Lubrication Figure 202 INDEX NO.

LOCATION

LUBRICANT

INTERVAL HOURS

ENGINE CONTROLS (NACELLE) 1.

Cam Plate and Pins

Grease (68, Table 1, Chapter 91-00-00).

200

Clean and lubricate with grease (23, Table 1, Chapter 91-00-00). Do not lubricate with oil.

400

Lubricant (69, Table 1, Chapter 91-00-00).

600

INERTIAL ANTI-ICE SYSTEM 2.

Hinge Point Bushings AIR CONDITIONER COMPRESSOR SYSTEM

3.

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Air Conditioner Compressor Quill Shaft

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Figure 202 Nacelle Engine Controls and Inertial Anti-Ice Lubrication

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I. Flight Compartment Engine Controls and Propeller Lubrication Figure 203 INDEX NO.

LOCATION

LUBRICANT

INTERVAL HOURS

ENGINE CONTROLS (FLIGHT COMPARTMENT) 1.

Linkage Bushing and Pins (Ref. Figure 203, Warning)

Grease (83, Table 1, Chapter 91-00-00).

600

PROPELLER CAUTION Do not use Aeroshell 17 on the hub grease fittings. 2.

Propeller Hub (Ref. Figure 203, Note)

Aeroshell 6 (preferred, and approved for temperatures down to -40°F) (Ref. Hartzell Owner’s Manual and Log Book No. 139).

400

3.

Low Pitch Stop Rods (Reversing Propeller)

Marvel Mystery Oil (84, Table 1, Chapter 91-00-00).

200

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Figure 203 Flight Compartment Engine Controls and Propeller Lubrication

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J. Nose Landing Gear Lubrication Figure 204 INDEX NO.

LOCATION

LUBRICANT

INTERVAL HOURS

NOSE LANDING GEAR 1.

Wheel Bearings

Clean, inspect bearings and bearing races for pitting, cracks, discoloration, rust, or indications of other wear or damage, and pack with grease (87, Table 1, Chapter 91-00-00).

At Tire Change or (Ref. NOTE: below)

DO NOT mix greases of different types or manufacturers. 2.

Door Hinges and Retract Linkage

Lubricating Oil (41, Table 1, Chapter 91-00-00).

200

3.

Grease Fittings

Lubricating Grease (61, Table 1, Chapter 91-00-00).

400

4.

Upper and Lower Nose Gear Strut Bearing

Use only specified grease (79, Table 1, Chapter 91-00-00).

200

5.

Steering Bellcrank Grease Fitting

Lubricating Grease (61, Table 1, Chapter 91-00-00).

400

6.

Steering Disconnect Cam (Mechanical Steering Only)

Lubricating Grease (68 or 107, Table 1, Chapter 91-00-00) lubricate sparingly.

400

NOTE: Refer to WHEEL BEARING LUBRICATION in this section for more information. After washing airplane, lubricate all lubrication points.

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Figure 204 Nose Landing Gear Lubrication

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K. Main Landing Gear Lubrication Figure 205 INDEX NO.

LOCATION

LUBRICANT

INTERVAL HOURS

MAIN LANDING GEAR 1.

Wheel Bearings

Clean, inspect bearings and bearing races for pitting, cracks, discoloration, rust, or indications of other wear or damage, and pack with grease (87, Table 1, Chapter 91-00-00).

At Tire Change or (Ref. NOTE: below)

DO NOT mix greases of different types or manufacturers. 2.

Grease Fittings

Lubricating Grease (61, Table 1, Chapter 91-00-00).

400

3.

Door Hinges and Linkage Bearings

Lubricating Oil (41, Table 1, Chapter 91-00-00).

200

4.

Door Retract Cam

Lubricating Grease (68 or 107, Table 1, Chapter 91-00-00) lubricate sparingly.

400

NOTE: Refer to WHEEL BEARING LUBRICATION in this section for more information. After washing airplane, lubricate all lubrication points.

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Figure 205 Main Landing Gear Lubrication

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L. Flight Compartment and Elevator Controls Lubrication Figure 206 INDEX NO.

LOCATION

LUBRICANT

INTERVAL HOURS

CONTROL COLUMN 1.

Chain

Clean with a cloth dampened in cleaning solvent (2, Table 1, Chapter 91-00-00) lubricate with SAE 30W mineral oil and wipe off excess.

1200

Lubricating Oil (41, Table 1, Chapter 91-00-00).

200

RUDDER PEDALS 2.

Pedal ELEVATOR CONTROL SYSTEM

3.

Elevator Trim Tab Hinge

Apply Lubricant (106, Table 1, Chapter 91-00-00) with a brush or squirt type can.

200

4.

Elevator Trim Tab Actuator Grease Fittings

Grease (1 or 17, Table 2, Chapter 27-00-00).

400

CAUTION Do not mix greases of different types or manufacturers. Mixing greases reduces lubricant effectiveness. The actuator was originally manufactured with grease (1, Table 2, 27-00-00) and this grease may have been cleaned out and replaced with grease (17, Table 2, 27-00-00).

5.

Elevator Trim Cable Pressure Seals

Grease (23, Table 1, Chapter 91-00-00) Clean and lubricate per CONTROL CABLES AND CABLE PRESSURE SEALS in this chapter.

1200

6.

Elevator Cable Pressure Seals Grease (23, Table 1, Chapter 91-00-00) Clean and lubricate per CONTROL CABLES AND CABLE PRESSURE SEALS in this chapter.

1200

7.

Elevator Trim Tab Chain

Grease (23, Table 1, Chapter 91-00-00).

1200

8.

Elevator and Trim Tab Cables

Apply Corrosion Preventive Compound (11, Table 1, Chapter 91-00-00) Clean and lubricate per CONTROL CABLES AND CABLE PRESSURE SEALS in this section.

1200

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Figure 206 Flight Compartment and Elevator Controls Lubrication

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M. Rudder Control System Lubrication Figure 207 INDEX NO.

LOCATION

LUBRICANT

INTERVAL HOURS

RUDDER CONTROL SYSTEM 1.

Rudder Trim Tab Actuator

Grease (1 or 17, Table 2, Chapter 27-00-00).

400

CAUTION Do not mix greases of different types or manufacturers. Mixing greases reduces lubricant effectiveness. The actuator was originally manufactured with grease (1, Table 2, 27-00-00) and this grease may have been cleaned out and replaced with grease (17, Table 2, 27-00-00). 2.

Rudder Trim Hinge

Apply lubricant (106, Table 1, Chapter 91-00-00) with a brush or squirt can.

200

3.

Rudder Tab Cable Pressure Seal

Grease (23, Table 1, Chapter 91-00-00) Clean and lubricate per CONTROL CABLES AND CABLE PRESSURE SEALS in this chapter.

1200

4.

Rudder Cable Pressure Seal

Grease (23, Table 1, Chapter 91-00-00) Clean and lubricate per CONTROL CABLES AND CABLE PRESSURE SEALS in this chapter.

1200

5.

Rudder and Trim Tab Cables

Apply corrosion preventive compound (11, Table 1, Chapter 91-00-00) with a brush; wipe off excess. clean and lubricate per control cables and cable pressure seals in this chapter.

1200

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Figure 207 Rudder Control System Lubrication

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N. Flap and Aileron Control System Lubrication Figure 208 CAUTION: Do not mix red MIL-G-81322 Grease (61, Table 1, Chapter 91-00-00) with dark tan MIL-G-10924 or SA826-3242 Grease (85, Table 1, Chapter 91-00-00). INDEX NO.

LOCATION

LUBRICANT

INTERVAL HOURS

FLAP CONTROL SYSTEM 1.

Flap Motor Gearbox

Pack gearbox with grease (85 or 61, Table 1, Chapter 91-00-00) to 0.3 inch from cover.

1200

NOTE When using dry film lubricant (82, Table 1, Chapter 91-00-00) ensure proper cure time, follow the manufacturer’s instructions. 2.

Flap Tracks

Lubricating Grease (68, 82 or 107, Table 1, Chapter 91-00-00).

400

3.

Flap Limit Switch Link

Lubricating Oil (41, Table 1, Chapter 91-00-00) apply to holes.

600

4.

Flap Asymmetry Switch Hub Assembly

Grease (23 or 61, Table 1, Chapter 91-00-00) Refer to Chapter 27-50-06 for lube procedures.

600

4A.

Flap Asymmetry Switch Hub Assembly (Modified By Kit 129-5046)

Grease (23 or 61, Table 1, Chapter 91-00-00) one pump from grease gun.

600

AILERON CONTROL SYSTEM 5.

Aileron Bellcrank (UA and UB Serials)

Lubricating Oil (41, Table 1, Chapter 91-00-00).

1200

6.

Bellcrank Rod Ends (UA and UB Serials)

Grease (23, Table 1, Chapter 91-00-00).

100

7.

Trim Tab Actuator

Grease (1 or 17, Table 2, Chapter 27-00-00).

400

CAUTION Do not mix greases of different types or manufacturers. Mixing greases reduces lubricant effectiveness. The actuator was originally manufactured with grease (1, Table 2, 27-00-00) and this grease may have been cleaned out and replaced with grease (17, Table 2, 27-00-00). 8.

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Trim Tab Cable Pressure Seals

12-20-00

Grease (23, Table 1, Chapter 91-00-00) Clean and lubricate per CONTROL CABLES AND CABLE PRESSURE SEALS in this chapter.

1200

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL INDEX NO.

LOCATION

LUBRICANT

INTERVAL HOURS

9.

Aileron Cable Pressure Seals

Grease (23, Table 1, Chapter 91-00-00) Clean and lubricate per CONTROL CABLES AND CABLE PRESSURE SEALS in this chapter.

1200

10.

Aileron and Trim Tab Cables

Apply corrosion preventive compound (11, Table 1, Chapter 91-00-00) clean and lubricate per CONTROL CABLES and CABLE PRESSURE SEALS in this chapter.

1200

11.

Aileron Trim Tab Hinge

Apply lubricant (106, Table 1, Chapter 91-00-00) with a brush or squirt can.

200

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1 10 9

10 8 10

10

4A 3

3 4 3

2

3

11 7 7

11

UC SERIALS 5

6

6 UA AND UB SERIALS

5

Figure 208 Flap and Aileron Control System Lubrication

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UC12B 050558AB.AI

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

O. Nose Avionics Door Lubrication Figure 209 CAUTION: Disassemble Index No. 1 joint only when the door is fully open. INDEX NO.

LOCATION

LUBRICANT

INTERVAL HOURS

NOSE AVIONICS DOOR 1.

Gas Spring End Fittings

Disassemble joint and lubricate sparingly with grease (23, Table 1, Chapter 91-00-00).

1200

2.

Door Hinge

Apply lubricant (106, Table 1, Chapter 91-00-00) with a brush or squirt-type can. Wipe off excess.

600

3.

Latch Pin and Plate

Wipe clean and lubricate the pin with lubricant (45, Table 1, Chapter 91-00-00).

200

4.

Latching Mechanism

Lubricating Oil (41, Table 1, Chapter 91-00-00).

200

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Figure 209 Nose Avionics Door Lubrication

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P. Cabin Door Lubrication Figure 210 INDEX NO.

LOCATION

LUBRICANT

INTERVAL HOURS

CABIN DOOR 1.

Cam Lip

Wipe clean and lubricate the cam lip using lubricant (45, Table 1, Chapter 91-00-00).

200

2.

Door Damper

Hydraulic Fluid (39, Table 1, Chapter 91-00-00).

200

3.

Door Hinge

Apply lubricant (106, Table 1, Chapter 91-00-00) with a brush or squirt-type can.

200

4.

Cam Surface of Pressure Lock Grease (23, Table 1, Chapter 91-00-00).

400

5.

Bearing Block Contact Surface or Grease Fittings

Grease (23, Table 1, Chapter 91-00-00).

600

6.

Door Handle

Lubricate door handle grease fitting with grease (83, Table 1, Chapter 91-00-00).

600

EMERGENCY EXIT DOORS 7.

Door Track

Lubricate track sparingly with grease (80, Table 1, Chapter 91-00-00) (6 Places per Door).

200

8.

Latching Mechanism

Dry Film Lubricant (82, Table 1, Chapter 91-00-00).

200

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 210 Cabin Door Lubrication

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12-20-00

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Q. Cargo Door Lubrication Figure 211 CAUTION: Disassemble Index No. 1 joint only when the cargo door is fully open and the stabilizer is in place. INDEX NO.

LOCATION

LUBRICANT

INTERVAL HOURS

CARGO DOOR 1.

Gas Spring End Fittings

Disassemble joint and lubricate sparingly with grease (23, Table 1, Chapter 91-00-00).

400

2.

Door Cam Lock

Wipe clean and lubricate the lip of the cam lock with lubricant (45, Table 1, Chapter 91-00-00). Do not apply to the face of the cam.

200

3.

Pushrod Pin and Bushing

Lubricate sparingly with lubricating oil (41, Table 1, Chapter 91-00-00).

600

4.

Latch Pin and Latch Plate

Wipe clean and lubricate the pin and the hole in the latch plate with lubricant (45, Table 1, Chapter 91-00-00).

600

5.

Door Hinge

Apply lubricant (106, Table 1, Chapter 91-00-00) with a brush or squirt can.

200

6.

Door Handle

Lubricate door handle grease fitting with grease (83, Table 1, Chapter 91-00-00).

1200

12-20-00

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Figure 211 Cargo Door Lubrication

Page 232 May 1/11

12-20-00

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

SERVICING UNSCHEDULED SERVICING MAINTENANCE PRACTICES

12-30-00 200200

1. INFORMATION A. De-icing and Anti-icing of Airplanes on the Ground (1) De-icing is the removal of ice, frost and snow from the airplane’s exterior after it has formed. Anti-icing is a means of keeping the surface clear of subsequent accumulations of ice, snow and frost. (2) Snow and ice on an airplane will seriously affect its performance. Even formation of a smooth covering of ice on the wing will change the contour of the wing, producing an increase in drag and a reduction in effective lift coefficient. Frost or frozen snow may present an even greater hazard since the surface texture is rough and will seriously disrupt the smooth flow of air across the wing.

B. De-icing and Anti-icing Fluids (1) Hawker Beechcraft Corporation has evaluated and approved the following de-ice/anti-ice fluids for use on this model: (a) SAE Type I Anti-icing Fluids (Unthickened-Type Fluids) Type I fluids (115, Table 1, Chapter 91-00-00) mainly provide protection against refreezing when there is no precipitation. (b) SAE Type II Anti-icing Fluids (Thickened-Type Fluids) Type II fluids (115, Table 1, Chapter 91-00-00) provide protection against refreezing when precipitation occurs. (c) SAE Type III Anti-icing Fluids (Thickened-Type Fluids) Type III fluids (115, Table 1, Chapter 91-00-00) provide protection against refreezing when precipitation occurs (d) SAE Type IV Anti-icing Fluids (Thickened-Type Fluids) Type IV fluids (115, Table 1, Chapter 91-00-00) provide protection against refreezing when precipitation occurs.

12-30-00

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All approved Type I, Type II, Type III and Type IV fluids may be used for either de-icing or anti-icing at any dilution as allowed by the fluid manufacturer's recommendations. Hawker Beechcraft Corporation cannot accept responsibility for damage to the airplane finish, windows, rubber seals, etc. resulting from the use of de-icing fluids not conforming to the specified specifications. These fluids were chosen according to the following specifics: 1 Noncorrosive. 2 Do not deteriorate rubber, painted surfaces, or plastics. 3 Have a high flash point. 4 Nontoxic. 5 Good self-wetting and antifoaming characteristics.

2. REMOVE FROST, SNOW OR ICE WARNING: Remove all snow, ice and frost before flight. Type I, Type II, Type III and Type IV; Glycols are listed in the ‘harmful’ category. You must wear protective goggles, protective gloves and clothing when you handle this material. Keep the material away from skin and eyes. Inhalation of glycol mists, aerosols, or high concentration of heated vapors poses a hazard to humans. Apply deicing fluid only in well-ventilated areas. Avoid inhaling vapors or mists. If adequate ventilation, designed to keep mists or vapors below harmful levels, is not available, maintenance personnel must wear approved respiratory protective devices. De-icing fluid Type I, has a limited period of effectiveness (referred to as ‘Holdover Times (HOT)’). The use of Type I fluid should only be considered where Type II, Type III or Type IV fluids are not available. CAUTION: Make sure that the correct fluid application equipment and correct procedures are used by qualified personnel so that the fluid will perform to the specifications. Fluid failure is complex and dependent on: •

Percent mix

•

Type and rate of precipitation

•

Product Holdover Time (HOT), and other variables

Each product is unique and reference must be made to the manufacturers recommendations and the FAA guidelines which are published each year. Though these fluids pass the DTD crazing test for transparent panels, do not apply hot spray directly onto window panels or seals as damage may occur. The fluids identified in this procedure are for ground de-icing/anti-icing only and are not intended for and do not provide protection during flight.

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12-30-00

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Residue from thickened de-icing/anti-icing fluids (Type ll, Type III or Type lV) can remain in aerodynamically quiet areas and accumulate over time. This residue can re-hydrate and expand into a gel-like material that could freeze during flight and cause restrictions in the flight control systems. The accumulation of residual fluid is more prevalent when using a one-step de-icing/ anti-icing process which is commonly used in Europe. For operators using the one-step application process, visually inspect flight control systems for the presence of dry or re-hydrated fluid at least twice during the winter operations season. For operators using the two-step application process, visually inspect flight control systems for the presence of dry or re-hydrated fluid once at the end of the winter operations season, as a recommended minimum procedure. Each operator must determine their own frequency of inspection based on the operational tempo of their airplane or fleet. For additional information concerning this guidance please contact the HBC field service representative or the Technical Support telephone line.

A. Guidelines to Holdover Times (HOT) (1) Visit FAA web site (www.faa.gov) and in the search field type [Holdover Time Tables] and press go/search. A listing will be called up that will include (but is not limited to) documents that read “FAA - Approved de-icing Program Updates” for the current year. When the Web Site cannot be accessed or questions arise, contact FAA Flight Standards, Washington DC at 202.267.8166 or Hawker Beechcraft Corporation 1.800.429.5372 or 316.676.3140. (2) Before take-off the pilot in command must be satisfied the airplane is clear of frost, snow or ice, within the limitations stated in the applicable Aircraft Flight Manual. (3) If the holdover time is exceeded and visual/tactile investigation of flight surfaces is not possible, then the airplane should be returned for further treatment with de-icing/anti-icing fluid prior to takeoff. (4) Freezing point of SAE/ISO Type I fluid mixture used must be at a minimum of 10°C (18°F) below OAT.

B. General Application WARNING: Make sure the airplane is grounded before any work is started. While these procedures are done, the surfaces will be very slippery due to snow, ice, or de-icing/anti-icing fluid; use a servicing stand whenever possible. If a stand is not available, use mainplane mats, safety belts, and take other similar precautions. Always stand upwind of the airplane. De-icing/anti-icing fluids are toxic. Avoid contact with skin and eyes, use goggles and protective clothing. Functional check of flight controls - If an airplane has been extremely iced or snow covered, a flight control check should be done. This check should be repeated after de-icing.

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CAUTION: Wing skin temperatures may differ and in some instances be lower than the OAT. A stronger mix (more glycol) can be used under these conditions. The de-icing/anti-icing fluid used must be to a concentration and at a temperature in accordance with the fluid manufacturer's instructions, provided the following limitations are not exceeded: •

For heated fluids, a temperature of 60° C (140° F) at the nozzle is desirable. The upper temperature limit shall not exceed fluid and airplane manufacturer’s recommendations.

•

Before you apply fluid, ground the airplane and make sure the covers and blanks are installed to the equipment that follows: - Static vents - Pitot heads - Stall detector vanes

To prevent possible ingestion of de-icing fluid, do not operate the engines or APU (if applicable) while de-icing is in progress. After snow is cleared from the surfaces, make sure the areas listed below are free from snow or ice: - All air intakes - Control surface hinges - Gaps between wing trailing edge shroud and flaps, airbrakes and ailerons - Gaps between horizontal stabilizer and elevators - Gaps between rudder and vertical stabilizer - Wheel brakes - Landing gear bays - All doors and external drains - Cold air unit ram air exhaust outlet and duct The directions given in the fluid manufacturer’s instructions with reference to the possibility of mixing UCAR ULTRA with the residues of other fluids must be strictly adhered to. Minor contamination with Type I fluids can significantly degrade the anti-icing performance.

C. Remove Frost (1) Spray with de-icing/anti-icing fluid to fluid manufacturer’s instructions. (2) Under severe frost forming conditions, after defrosting, give a further light application of the concentrated fluid to make sure the maximum holdover period.

D. Remove Sleet and Freezing Rain CAUTION: After spraying, examine surface thoroughly, as ice formed by freezing rain can be difficult to see under the de-icing fluid and may require touch for confirmation. •

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Spray with undiluted de-icing/anti-icing fluid.

12-30-00

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

E. Remove Snow (1) Dry snow (a) Remove by brushing, or by the application of a cold air blast, taking care not to trap the snow in the control surface gaps and hinges. (b) If the ambient temperature is above freezing, the snow can be removed by the application of a hot air blast, and the cleared surface sprayed with diluted de-icing/anti-icing fluid. (2) Wet snow (a) Remove with rubber squeegees, taking care not to trap the snow in the control surface gaps and hinges. (b) If ice has formed under the snow, clear by spraying with de-icing/anti-icing fluid. (3) Frozen snow and ice films (a) Clear off any loose snow, then apply a heavy spray of de-icing/anti-icing fluid to the manufacturer’s instructions. •

Brush the snow as the fluid is being applied; this will assist in breaking up the deposit and help to retain the fluid on the deposit. When all frozen deposits have been removed, give a final light spray.

(b) If the ambient temperature is above 0°C (32°F), ducted hot air blasts may be used to disperse the ice. •

Do not use hot air blasts near windows, and take extreme care to prevent damage by overheating to painted surfaces, rubber, glass, acrylic or glass fibre, hydraulic pipelines, grease or oiled surfaces.

•

Brush off or mop up water resulting from melted ice, as soon as possible.

F. De-ice the Windshield •

Lightly spray windshield with fluid, windshield washing (DTD.900AA/4939A). If smearing occurs, wipe clear using a warm damp rag.

G. De-ice the Landing Gear CAUTION: De-icing/anti-icing fluid must not come into contact with landing gear electrical plugs and harnesses. Do not let de-icing/anti-icing fluid contact brake units. (1) Brush-off loose accumulations of snow with soft brush. (2) Remove stubborn deposits with a rag soaked in de-icing/anti-icing fluid.

12-30-00

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H. Anti-icing •

If meteorological reports predict the onset of freezing rain or heavy frost deposits, an anti-icing spray of cold (or hot if cold not available) concentrated ground de-icing/anti-icing fluid Type II, Type III or Type IV, will give limited holdover protection times, guidance for which is contained in Web Site mentioned in paragraph GUIDELINES TO HOLDOVER TIMES (HOT). If the precipitation is extremely heavy, deposits will form on top of the coating, but these will have little, if any, adhesion to the surfaces and can be easily removed with a light spray of hot diluted fluid.

I. Fluid Spills Glycol-based deicing fluids are biodegradable in water. Only gross contamination of slow moving or restricted bodies of water would be likely to cause any serious environmental impact. Typical field-use concentrations of deicing fluids, particularly when diluted by snow, ice or water, causes little or no injury to most broad leaf plants, grasses perennial ground cover, and woody plants. Minor leaks or spills of deicing fluid in storage areas must be soaked up with an absorbent material, such as sawdust, vermiculite, an all-purpose commercial oil absorbent, or sand. Carefully shovel the absorbent/deicing fluid into an appropriate container for disposal. Spilled, leaked, or contaminated deicing fluid must be disposed of in strict compliance with all applicable federal, state, and local regulations and ordinances.

J. Remove Salt or Chemical Agents CAUTION: Do not use high pressure water; damage may result to electrical equipment and lubricated components. (1) Where contamination has occurred on the structure, due to the airplane landing on airfields where the snow or ice has been dispersed with salt or chemical melting agents: (a) Wash down the affected areas with clean water as soon as possible. A wetting agent, such as detergent cleaner (Teepol TS610) or Comprex A, may be added in small quantities. (b) When the time or conditions prevent removal of the contamination at outstations, this must be noted in the technical log, so that appropriate action may be taken to remove it at main base.

K. Residue from De-ice/Anti-ice (1) Inspection (a) Gain access to areas, potentially affected by residue, where flight controls and other systems components are located. (b) Visually inspect for the presence of dry or rehydrated residue anywhere in these areas. The residue may be difficult to see, especially if dry. Dry residue tends to be in the form of a thin film that may be partially covered with grease or dirt. Rehydrated residue often appears as a gel-like substance which is thicker and therefore more visible.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

(2) Cleaning (a) After identifying any residue it should be removed by using warm water with rags and/or soft bristle brushes to clean the gel-like material away. CAUTION: Make sure the water or compressed air does not cause any residue to enter areas that are not accessible. Do not allow runoff from the cleaning process to enter other areas of the airplane. Avoid spraying cleaning fluids onto bearings, fittings, control cables and electrical connectors. Do not spray controls with water if the ambient temperature is below freezing unless the airplane is located in a heated hangar. This cleaning process can, potentially, remove grease from control system bearings and fittings, and remove corrosion inhibitors from control cables. (b) The use of a low-pressure stream of water or compressed air to rinse away any residue may prove helpful. Using Type I de-icing fluid, or a mixture of water and Type I fluid, is also a good cleaning agent for removal of residue. (3) Relubrication (if required) •

If residue has been found and removed by cleaning, it is recommended that all bearings, fittings, and control cables in the affected area be relubricated (Ref. Chapter 12-20-00, 201, SCHEDULED SERVICING).

3. FIRE EXTINGUISHER AGENT CONTAMINATION In the event of engine contamination by fire extinguisher agents, (Ref. the latest revision of PT6A-65 ENGINE MAINTENANCE MANUAL P/N 3032842 or subsequent revision and Chapter 72-00-00, ENGINE, TURBOPROP - INSPECTION.

4. BIOCIDAL AGENT TREATMENT If evidence of microbiological contamination is detected, sterilize the fuel system using a biocidal agent (63, Table 1, Chapter 91-00-00) in shock dose treatment quantities (Ref. BIOCIDAL AGENT, Chapter 12-10-00). Repeat this shock treatment every 90 days until no contamination is found. If no evidence of contamination is found, the fuel system should be sterilized annually.

12-30-00

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CHAPTER 20 - STANDARD PRACTICES - AIRFRAME TABLE OF CONTENTS SUBJECT

PAGE

GENERAL INFORMATION 20-00-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1

ELECTRICAL BONDING 20-00-01 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Bonding Surface of Aluminum . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Preparation of Aluminum Surfaces to be Bonded or Grounded . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Preparation of Surface . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Magnesium Alloy Surface Cleaning . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Steel Surface Cleaning . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Steel Brush Cleaning of Aluminum . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Typical Blind Ground Stud Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . CRES Steel or Titanium Bonding If Temp Is Below 330°F . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . CRES Steel or Titanium Bonding If Temp Exceeds 300°F . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Aluminum or Magnesium Alloy Bonding If Temp Is Under 300°F . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Typical Grounding Stud . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Electrical Bonding Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 201 201 201 202 202 203 204 205 207 209 211 213

CONTROL CABLES AND PULLEYS 20-00-02 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Control Cable System Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Control Cable Storage . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Control Cable Pulley Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 201 202 202

WIRING 20-00-03 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Control Column Clearance Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Wiring Forward Of The Instrument Panel . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201

ELECTROSTATIC DISCHARGE SENSITIVITY 20-00-04 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Classification of Electrostatic Discharge Sensitivity . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 ESDS Equipment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Removal/Installation of ESDS Equipment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Handling of ESDS Components and Equipment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Controlling Static Charge Buildup . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202 Permanent Static Control Workstation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202

20-CONTENTS

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CHAPTER 20 - STANDARD PRACTICES - AIRFRAME TABLE OF CONTENTS (CONTINUED) SUBJECT PAGE Portable Static-Control Workstation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 206 Humidity and Dust Effects on ESDS Components and Equipment . . . . . . . . . . . . . . . . . . . . . . . . . . . 208 Packaging of ESDS Components and Equipment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 208 Marking of ESDS Components and Equipment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 210 Storage and Transit of ESDS Components and Equipment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 210

TUBING, HOSE AND FITTINGS 20-00-05 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Tubing, Hose and Fittings . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Hose Assembly Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202 Tube Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202 Fluid Line Fitting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 203 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 203 Nonpositioning Type Fitting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 204 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 204 Pipe Thread Fitting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 204 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 204 Tube Damage . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 205 Limits . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 205 Tube Manufacturing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 205 Conical Seal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 212 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 212

TORQUE WRENCHES 20-01-00 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201

LEADING EDGE EROSION PROTECTION 20-04-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Leading Edge Erosion Protection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201

BEARINGS 20-05-00 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Hydraulic Press Bearing Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Mechanical Press Bearing Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Bearing Housing Inspection. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Page 2 Nov 1/12

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CHAPTER 20 - STANDARD PRACTICES - AIRFRAME TABLE OF CONTENTS SUBJECT PAGE Bearing Installation Using Retaining Compound . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .202 Bearing Installation by Staking . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .203

LOCKING DEVICES 20-07-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Self-Locking Nuts . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Lockwire and Cotter Pin Requirements . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Taper Pins . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Slotted, Steel Locknuts (Prevailing Torque Type) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Standard And Stepped Studs . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Hose, Tubing and Threaded Couplings . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Lockwire . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Retaining Rings (Spirolox, etc.) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202 Taper Pins (AN386 Only) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202

AIRPLANE FINISH CARE 20-08-00 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Cleaning Airplane Finishes . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . During the Curing Period (One Month) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . After the Curing Period . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Environmental Fallout (Acid Rain) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Placard Replacement . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Exterior Finishes (Aluminum Surfaces) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Urethane Paints . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Urethane Paint Repair Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Paint Stripping and Cleaning Urethane Paint . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Urethane Primer . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Urethane Topcoats . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Urethane Touch-Up Repair . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Special Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Paint Free Areas . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Magnesium Surfaces . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Paint Removal From Magnesium Surfaces . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Painting Magnesium Surfaces . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 201 201 201 202 202 202 202 202 202 203 203 204 204 208 210 210 211

CORROSION 20-09-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201

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CHAPTER 20 - STANDARD PRACTICES - AIRFRAME TABLE OF CONTENTS (CONTINUED) SUBJECT

PAGE

AIRFRAME PENETRATION INSPECTION 20-10-00 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201

CONTROL OF LIFE-LIMITED PARTS 20-15-00 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Control Of Life-limited Parts . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Determination Of Serviceability When Part Life Is Unknown . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201

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List of Effective Pages CH-SE-SU

PAGE

DATE

20-LOEP

1

Nov 1/13

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1 thru 4

Nov 1/12

20-00-00

1

Nov 1/09

20-00-01

201 thru 215

May 1/10

20-00-02

201 thru 203

Nov 1/13

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1 201 and 202

Nov 1/09 Nov 1/09

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1 thru 3 201 thru 213

Nov 1/09 Nov 1/09

20-00-05

1 201 thru 214

Nov 1/09 Nov 1/13

20-01-00

201 thru 203

Nov 1/09

20-04-00

1 201 thru 203

Nov 1/09 Nov 1/09

20-05-00

201 thru 204

May 1/12

20-07-00

1 and 2 201 thru 206

Nov 1/09 Nov 1/09

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201 thru 211

Nov 1/09

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Nov 1/09 Nov 1/09

20-10-00

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Nov 1/09

20-15-00

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Nov 1/09

20-LOEP

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

STANDARD PRACTICES - AIRFRAME GENERAL INFORMATION DESCRIPTION AND OPERATION

20-00-00 00

1. GENERAL WARNING: Any maintenance requiring the disconnection and reconnection of flight control cables, plumbing, electrical connectors or wiring requires identification of each side of the component being disconnected to facilitate correct reassembly. At or prior to disassembly, components should be color coded, tagged or properly identified in a way that it will be obvious how to correctly reconnect the components. After reconnection of any component, remove all identification tags. Check all associated systems for correct function prior to returning the airplane to service.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

STANDARD PRACTICES - AIRFRAME ELECTRICAL BONDING MAINTENANCE PRACTICES

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200200

1. PROCEDURES A. Bonding Surface of Aluminum A clean, smooth surface is required for bonding. The surface preparation for an electrical bond requires the complete removal of any anodic film, grease, oil, paint, lacquer, metal finishes or other high-resistance properties from an area slightly larger than the contact area. This will ensure negligible radio frequency (RF) impedance between adjacent metal parts.

B. Preparation of Aluminum Surfaces to be Bonded or Grounded (1) The mating surfaces must be smooth and contoured so that the entire mating surface areas are in actual contact. (2) Remove all protective films with fine sand paper or suitable solvent. (3) Clean all surfaces with solvent (14, Table 1, Chapter 91-00-00). (4) Shake the container of chemical conversion coating (88, Table 1, Chapter 91-00-00) vigorously and apply the chemical conversion coating (brush-on strength) to the cleaned area with a clean Scotch-brite, sponge or equivalent applicator. Keep the area wet with solution 3 to 5 minutes or until a yellow color develops. (5) Using clean water, dampen a clean cloth and gently wipe the treated area. Wipe with care as the formed coating is very soft while wet. (6) Allow the treated area to thoroughly air dry (1 hour maximum) prior to installation of mating parts. If this time limit is exceeded, retouch with chemical conversion coating (88, Table 1, Chapter 91-00-00).

C. Preparation of Surface The following information has been found satisfactory in preparation of metals for electrical mating surfaces. Grease, oil and other non conductive films should be removed with solvent (14, Table 1, Chapter 91-00-00). Non soluble films should be removed by sanding and polishing with very fine garnet, silicon carbide, or other aluminum oxide paper. Use caution so as not to remove any excessive metal. The area should be brushed clean and wiped with solvent (14, Table 1, Chapter 91-00-00). NOTE: No emery or iron oxide papers/cloth may be used to clean the surface. A small area on an aluminum surface may be cleaned by using a stainless steel wire brush with a pilot. Wipe off cleaned area with solvent (14, Table 1, Chapter 91-00-00) and a clean dry cloth. (Ref. Figure 201).

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D. Magnesium Alloy Surface Cleaning (1) Remove grease and oil from surface with solvent (14, Table 1, Chapter 91-00-00). (2) Remove any paint or lacquer, from the surface with lacquer thinner. (3) Brush area liberally with a chrome pickle solution for one minute, then rinse immediately by brushing with clean water to remove all chemicals. (4) Dry thoroughly.

E. Steel Surface Cleaning When the surface is corrosion-resistant (CRES) or plated steel, clean the bonding surfaces as follows: (1) Remove any grease and oil from the surface with solvent (14, Table 1, Chapter 91-00-00). (2) Remove any paint or lacquer from the surface with lacquer thinner and dry thoroughly. NOTE: Do not remove the zinc or cadmium plate from the steel surfaces.

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F. Steel Brush Cleaning of Aluminum (1) A stainless steel wire brush with a pilot shall be used to thoroughly clean the aluminum surface (Ref. Figure 201). (2) Wipe off the cleaned area with a dry cloth and clean with solvent (14, Table 1, Chapter 91-00-00). (3) Chemically treat the cleaned surface as instructed under the heading PREPARATION OF SURFACE in this chapter.

Figure 201 Steel Brush Cleaning of Aluminum

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G. Typical Blind Ground Stud Installation (1) The side walls of the mounting hole must be clean and free of all chemical films, grease and paint. (Cleaning of the upper and lower surfaces are not required (Ref. Figure 202). (2) Bond as instructed under the heading of PREPARATION OF SURFACE in this section.

Figure 202 Typical Blind Ground Stud Installation

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H. CRES Steel or Titanium Bonding If Temp Is Below 330°F NOTE: Lock washers shall be used on all bolted bonding/ground connections. Their function is to ensure a tight connection with plain or self-locking nuts under conditions where thermal expansion of the fastener occurs (Ref. Figure 203). Bolt Size: Use only a No. 6 or No. 8 screw where edge distance will not permit use of a No. 10 screw. A 3/16-inch diameter minimum should be used when possible. Current Return: A 100-amp current return requires a 1/4-inch diameter minimum size fastener. A 200-amp current return requires a 5/16-inch diameter minimum size fastener. BONDING INSTRUCTIONS (1) MS35338 lock washer may be used. (2) MS21042L self-locking nut, MS21047L or MS21069L self-locking nut plate may be used. (3) Location of the nut plate or the head of the bolt is optional. (4) Clean and seal as instructed under the heading PREPARATION OF SURFACE in this section. (5) Corrosion protect as instructed under the heading PREPARATION OF SURFACE in this section.

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JUMPER ASSEMBLY

SCREW OR BOLT CAD PLATED STEEL

LOCK-WASHER WASHER, STEEL LIGHT SERIES PLATED

PLATED STEEL, CRES STEEL OR TITANIUM

SELF-LOCKING NUT OR SELF-LOCKING PLATE NUT

SEAL AFTER INSTL. 1-1/2 DIA. CLEANED AREA

CLEAN TO BASE METAL AREA 1-1/2 DIA. OF TERMINAL

UC27B 043604AA.AI

Figure 203 CRES Steel or Titanium Bonding if Temp is Below 330°F

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I. CRES Steel or Titanium Bonding If Temp Exceeds 300°F NOTE: Lock washers shall be used on all bolted bonding/ground connections. Their function is to ensure a tight connection with plain or self-locking nuts under conditions where thermal expansion of the fastener occurs (Ref. Figure 204). Bolt Size: Use only a No. 6 or No. 8 screw where edge distance will not permit use of a No. 10 screw. A 3/16-inch diameter minimum should be used when possible. Current Return: A 100-amp current return requires a 1/4-inch diameter minimum size fastener. A 200-amp current return requires a 5/16-inch diameter minimum size fastener. BONDING INSTRUCTIONS (1) MS35338 lock washer may be used. (2) MS21042L self-locking nut, MS21047L or MS21069L self-locking nut plate may be used. (3) No sealing required where the maximum temperature exceeds 600°F. (4) Location of the nut plate or the head of the bolt is optional. (5) Clean and seal as instructed under the heading PREPARATION OF SURFACE in this section. (6) Corrosion protect as instructed under the heading PREPARATION OF SURFACE in this section.

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JUMPER ASSEMBLY

TITANIUM OR CRES STEEL SCREW OR BOLT

LOCK-WASHER WASHER, STEEL LIGHT SERIES PLATED

PLATED STEEL, CRES STEEL OR TITANIUM

SELF-LOCKING NUT OR SELF-LOCKING PLATE NUT

SEAL AFTER INSTL. 1-1/2 DIA. CLEANED AREA

CLEAN TO BASE METAL AREA 1-1/2 DIA. OF TERMINAL

UC27B 043605AA.AI

Figure 204 CRES Steel or Titanium Bonding if Temp Exceeds 330°F

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J. Aluminum or Magnesium Alloy Bonding If Temp Is Under 300°F CAUTION: Do not use magnesium for an electrical current return. NOTE: Lock washers shall be used on all bolted bonding/ground connections. Their function is to ensure a tight connection with plain or self-locking nuts under conditions where thermal expansion of the fastener occurs (Ref. Figure 205). Bolt Size: Use only a No. 6 or No. 8 screw where edge distance will not permit use of a No. 10 screw. A 3/16-inch diameter minimum should be used when possible. Current Return: A 100-amp current return requires a 1/4-inch diameter minimum size fastener. A 200-amp current return requires a 5/16-inch diameter minimum size fastener. BONDING INSTRUCTIONS (1) MS35338 lock washer may be used. (2) MS21042L self-locking nut, MS21047L or MS21069L self-locking nut plate may be used. (3) Location of the nut plate or the head of the bolt is optional. (4) Clean mating surface and seal as instructed under the heading PREPARATION OF SURFACE in this section. (5) Corrosion protect as instructed under the heading PREPARATION OF SURFACE in this section.

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Figure 205 Aluminum or Magnesium Alloy Bonding if Temp is Under 300°F

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K. Typical Grounding Stud NOTE: Lock washers shall be used on all bolted bonding/ground connections. Their function is to ensure a tight connection with plain or self-locking nuts under conditions where thermal expansion of the fastener occurs (Ref. Figure 206). Bolt Size: Use only a No. 6 or No. 8 screw where edge distance will not permit use of a No. 10 screw. A 3/16-inch diameter minimum should be used when possible. Current Return: A 100-amp current return requires a 1/4-inch diameter minimum size fastener. A 200-amp current return requires a 5/16-inch diameter minimum size fastener. BONDING INSTRUCTIONS (1) Bond as instructed under the heading PREPARATION OF SURFACE in this section.

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Figure 206 Typical Grounding Stud

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L. Electrical Bonding Check When electrical components are changed or modified, the integrity of the bonding procedure can be verified by utilizing the central grounding point. The central grounding point is located on the bottom of the fuselage in the center just aft of the main spar and is the primary ground point for the bonding check. The bonding check may be accomplished by using the secondary grounding points (adjacent aircraft structure). The central grounding point is covered by a large washer attached to the airplane with a screw. In order to gain access to the grounding point the screw and washer must be removed. After the bonding check is complete, replace the central grounding point washer and screw. Proper bonding procedures will result in minimal resistance readings between the replaced electrical component and the central grounding point (Ref. Figure 207). BONDING INSTRUCTIONS (1) Bonded connections should be located in a protected area and, if at all possible, near an inspection door or in an accessible location that provides ease of replacement and inspection. (2) Components should be bonded directly to the structure, not through other bonded parts (i.e. plumbing, etc.). (3) Components should be bonded to the structure with the shortest acceptable bond cable. (4) When movable parts are involved, bonding jumpers must be installed that will not hinder the component's movement. (5) Bonding connections should not be compression fastened through nonmetallic materials. (6) All bonding surfaces must be cleaned prior to bond installation. (7) Only self-locking nuts should be used for bonding connections. (8) RF current returns should not be made through magnesium. (9) Solder joints cannot be used alone for bonding parts that are subject to movement or vibration. (10) The structural integrity of the airframe must not be adversely affected by any bonding procedure. (11) Nonmetallic inserts or dry film lube nutplates cannot be used for bonding purposes. (12) AC and DC ground returns must be connected separately. (13) Shielded wires should be bonded to the structure unless specifically noted to the contrary. (14) Unless prohibited by space, multiple jumpers or dual system grounds (LH and RH systems) should not be connected to the same ground point. (15) There are six classes of bonding listed in MIL-B-5087. For their purposes and resistant requirements (Ref. Table 201).

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Table 201 RESISTANCE REQUIREMENTS Class

Purpose

Resistance (ohm)

A

Antenna installations: To assure functional performance of antennas by minimizing RF impedance.

C

Current path return: To use part of airplane structure as a current return path (minus line).

0.01

H

Shock hazard: To ground the case of the equipment in order to prevent electric shocks upon casual contact with electric equipment.

0.1

L

Lightning protection: To safely release lightning current caused when the airplane is struck by lightning, where current flows through the airplane structure causing heat (the path of electric resistance is at a high level), resulting in burn loss in the fuselage and equipment.

0.01

R

RF potentials: Provide low impedance to prevent radio noise from those parts generating electromagnetic energy and those affected by radio frequency.

0.00

S

Static charge: To prevent electric shocks due to the accumulation of static charge and RF noise due to discharge.

1.0

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0.0025

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A

DETAIL

A UC27B 042944AA.AI

Figure 207 Central Ground Point

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STANDARD PRACTICES - AIRFRAME CONTROL CABLES AND PULLEYS MAINTENANCE PRACTICES

20-00-02 200200

1. PROCEDURES A. Control Cable System Inspection WARNING: When inspecting control cables, always wear gloves to avoid injury from frayed or broken wires. When a control cable is removed from the airplane, the cable should be dipped in corrosion preventive (11, Table 1, Chapter 91-00-00). Excess corrosion preventive may be removed by wiping with a clean cloth. NOTE: Anywhere a cable is not visible, the flight controls shall be manipulated so that 100% of all cables can be inspected. Inspect the control cable system as follows: (1) Inspect the control cables for incorrect routing, fraying and twisting. Look for interference with adjacent airplane structure, equipment, wiring, plumbing and other control cables. (2) Monitor control cable movement for freedom, looseness and full travel. (3) Visually inspect all swaged fittings for distortion, cracks or broken wires at the fitting. (4) Turnbuckles should have the proper thread exposure and be correctly safety wired. (5) Locate any control cable broken or corroded wires as follows: (a) Inspect the control cables near fairlead pulleys by passing a cloth along the length of the cable. If a snag is found, closely examine the cable to determine the extent of the damage. (b) Any suspect cable should be removed and placed in a loop position and checked for additional broken wires.

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Figure 201 Typical Control Cable With Broken Wires NOTE: Individual broken wires are acceptable in primary and secondary control cables at random locations when there are no more than three broken wires on any given three foot cable segment. (c) Inspect the control cables with broken wires for evidence of corrosion. If necessary, remove the control cable, form it into a loop, and check the center strand for corrosion. Replace any control cable that shows evidence of corrosion. NOTE: The interior of all turnbuckles should be coated or filled with grease (23, Table 1, Chapter 91-00-00) for corrosion protection.

B. Control Cable Storage Control cables should be stored straight or in a coil. When stored in coil form, the coil inside diameter should not be less than 150 times the control cable diameter, or bent in a radius of not less than 75 times the control cable diameter. Coils should not be flattened, twisted or folded during storage. Storage requirements should apply until the control cable is installed in its normal position in the airplane. If only a part of the control cable is installed in an assembly, control cable storage requirements apply to the uninstalled portion of the control cable.

C. Control Cable Pulley Inspection Inspect all control cable pulleys as follows: NOTE: Control cable pulleys are installed along the control cables where a change of direction is needed. (1) Inspect all control cable pulleys for roughness, sharp edges and the presence of foreign material embedded in the grooves (Ref. Figure 202). (2) Inspect all control cable pulley bearings for smooth rotation, freedom from flat spots and foreign material. (3) Inspect all control cable pulleys for proper alignment. Page 202 Nov 1/13

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (4) Inspect the control cable pulley brackets and guards for damage, misalignment and looseness. (5) Control cable pulleys which turn for a short distance must be rotated periodically to provide a new bearing surface for the control cable.

Figure 202 Control Cable Pulley Wear Patterns

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STANDARD PRACTICES - AIRFRAME WIRING DESCRIPTION AND OPERATION

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1. GENERAL Information for inspection of the electrical wiring and criteria for repair or replacement of the electrical wiring is contained in AC 43.13-1B or subsequent. Any questions should be addressed to Hawker Beechcraft Technical Support at 1-800-429-5372 or 316-676-3140.

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200200

STANDARD PRACTICES - AIRFRAME WIRING MAINTENANCE PRACTICES 1. CONTROL COLUMN CLEARANCE INSPECTION A. Wiring Forward Of The Instrument Panel (1) Gain access to the area forward of the instrument panel on either the pilot’s or copilot’s side. (2) Inspect the area forward of the instrument panel to assure that no electrical/avionics wiring, wire bundles, air ducting, plumbing (both hard and soft lines), etc., are not touching or can come in contact with any part of the flight control mechanisms. (3) One mechanic should be positioned to inspect the area forward of the instrument panel while another mechanic moves the control column. (4) Slowly move the pilot (3) or copilot (1) control wheel aft and forward to move the elevator system through the full range of travel while this inspection is being performed (Ref. Figure 201). (a) Observe the chain and cable (5) of the control column and check for clearance. (b) Observe the top of the control column vertical tube (2) and check for clearance. (c) Observe the bob weight (4) and check for clearance. (5) Continue to move the control wheel and aft and forward and turn the control wheel to move the aileron system through the full range of travel while this inspection is being performed. (6) Any item that comes in contact with the control mechanisms forward of the instrument panel shall be secured clear of the mechanism. (7) Repeat Steps (1) thru (6) for the opposite side of the flight compartment.

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1

3 2

5

A 4

B 2

4

DETAIL

A

1. COPILOT'S CONTROL WHEEL 2. CONTROL COLUMN VERTICAL TUBE 3. PILOT'S CONTROL WHEEL 4. BOB WEIGHT 5. CHAIN AND CABLE

DETAIL

B UC20B 061334AA.AI

Figure 201 Control Column Travel

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STANDARD PRACTICES - AIRFRAME ELECTROSTATIC DISCHARGE SENSITIVITY DESCRIPTION AND OPERATION

20-00-04 00

1. GENERAL Some types of electronic components are easily damaged by electrostatic discharge (ESD), and require special handling and storage procedures. ESD is a release of stored electrostatic charge which has been generated by actions such as contact, rubbing, or separating of materials. A charge of this type can damage electrical and electronic equipment installed in the airplane. In some instances, the damage may not be immediate, but progressive. Components and items of equipment that can be damaged by electrostatic discharge are considered to be electrostatic discharge sensitive (ESDS). Electronic components that are considered to be electrostatic discharge sensitive include integrated circuits, transistors and diodes, monolithic and hybrid microelectronics, MOS capacitors, thin film resistors, and piezoelectric crystals. Any circuit or piece of equipment containing ESDS components is subject to ESD damage if certain handling precautions are not observed. Personnel who remove, inspect, test or install instruments and equipment containing ESDS components must be aware of the possibility of ESD damage, and should handle ESDS components in accordance with procedures covered in this chapter. Proper procedures and policies for the handling of ESDS components and equipment should be adhered to for the following reasons: •

Control of ESD damage, from time of component manufacture to time of actual installation, must be verifiable and must be maintained by use of established industry standards.

•

Established policy dictates that all personnel follow certain procedures to prevent damage to ESDS components and equipment.

•

Personnel in interacting areas of responsibility must be aware of their obligation to maintain proper ESD-controlled environments.

. Table 1 MATERIAL POLARITY MATERIALS

CHARGE (Relative Magnitude and Polarity)

MATERIALS

CHARGE (Relative Magnitude and Polarity)

Air

Positive

Sealing Wax

Negative

Human hands

Positive

Hard Rubber

Negative

Asbestos

Positive

Nickel, Copper

Negative

Rabbit fur

Positive

Brass, Silver

Negative

Glass

Positive

Gold, Platinum

Negative

Mica

Positive

Sulfur

Negative

Human hair

Positive

Acetate, Rayon

Negative

Nylon

Positive

Polyester

Negative

Wool

Positive

Celluloid

Negative

Fur

Positive

Orlon

Negative

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Table 1 MATERIAL POLARITY (Continued) MATERIALS

CHARGE (Relative Magnitude and Polarity)

MATERIALS

CHARGE (Relative Magnitude and Polarity)

Lead

Positive

Saran

Negative

Silk

Positive

Polyurethane

Negative

Aluminum

Positive

Polyethylene

Negative

Paper

Positive

Polypropylene

Negative

Cotton

Positive

PVC (Vinyl)

Negative

Steel

Neutral

KRL-F (CTFE)

Negative

Wood

Negative

SILICON

Negative

Amber

Negative

TEFLON

Negative

Table 1 lists several materials and the associated electrostatic charge polarity and magnitude for each. Materials at the top of the list are capable of producing the greatest amount of positive electrostatic charge, while materials at the bottom of the list are capable of producing a similar negative electrostatic charge. Items of dissimilar polarity provide the greatest potential for electrostatic discharge. Numeric values have not been assigned to the listed materials, as static charge levels are not constant, and will vary with ambient conditions. A greater possibility of ESD exists when the positions of listed items in Table 1 are farther apart. For example, an individual using his/her hands to pick up a PVC pipe has more potential for producing ESD than does an aluminum part contacting a steel part. Table 2 identifies some typical electrostatic charge levels and the actions that can produce the electrostatic charge. Table 2 TYPICAL ELECTROSTATIC VOLTAGES Actions of persons

MOST COMMON READING (VOLTS)

HIGHEST READING (VOLTS)

Walking across carpet

12,000

39,000

Walking across vinyl tile floor

4,000

13,000

Seated in polyurethane foam chair

1,800

18,000

Picking up poly bag

1,500

20,000

Inserting paperwork into vinyl envelopes

800

7,000

This data based on an ambient relative humidity of 15 to 36 percent.

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2. CLASSIFICATION OF ELECTROSTATIC DISCHARGE SENSITIVITY Three levels of sensitivity classification are established for electrostatic discharge sensitivity devices. Classification is used to aid the manufacturer or supplier in providing packaging and handling requirements that protect the ESD sensitive item, device or component through all phases of handling and packaging of the device during its service life. The three classes of ESD sensitivity are as follows: Class 1 - Sensitivity range is from 0 to 1,999 volts. Class 2 - Sensitivity range is from 2,000 to 3,999 volts. Class 3 - Sensitivity range is from 4,000 to 15,999 volts.

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200200

STANDARD PRACTICES - AIRFRAME ELECTROSTATIC DISCHARGE SENSITIVITY MAINTENANCE PRACTICES 1. ESDS EQUIPMENT A. Removal/Installation of ESDS Equipment Observe the following procedures when removing or installing ESDS equipment: CAUTION: Tools with plastic or insulated handles should not be used around ESDS devices. These tools can carry a static charge which does not readily discharge during the grounding process. Insulated tools should be used only during power-on testing of aircraft systems to prevent electrical shock to maintenance personnel performing the tests. Some circuit board assemblies may be protected by plastic covers. These covers can store an electrostatic charge. Use a static control work station to neutralize any electrostatic charge on the covers before touching a printed circuit board. Store the covers a safe distance from the work area. (1) When using test equipment, discharge all test leads to the ground prior to connection to the ESDS circuit under test. (2) Use a portable static control work station when removing ESDS circuit boards from card cages and enclosures at the airplane. (3) Place removed ESDS equipment on the static dissipative surface of the work station before opening the static shielding container holding the replacement ESDS equipment. (4) Just prior to engaging a cable connector with its mating receptacle, touch the connector shell to the receptacle shell to neutralize any electrostatic charge on the connector or the installer’s body. (5) Maintain protective coverings on stored ESDS equipment.

B. Handling of ESDS Components and Equipment All personnel handling ESDS components and equipment should receive instruction in the proper handling of such items. Observe the following handling rules to prevent damage to ESDS components and equipment: (1) Keep ESDS components and equipment inside ESD protective packaging until opened at a static control work station. (2) Before unsealing ESD protective packages, place the packages on the work surface of a static control work station. (3) Do not use pressure air nozzles to remove dust from ESDS printed circuit boards. Rapid movement of air, combined with airborne dust particles, can create an electrostatic charge that will destroy ESDS components. (4) Always wear a grounding wrist strap when opening any ESD protective package. (5) Avoid touching circuit components or connector pins when handling ESDS components or equipment.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (6) Never place any ESDS component, before or after assembly, on a nonconductive surface or in a container not specifically designed for storage of ESDS devices. (7) Protect ESDS components and equipment with protective containers, conductive caps, and/or pin-shorting devices. (8) Store and transport ESDS components and equipment in ESD protective containers. Seal all protective containers with an ESD warning label prior to shipment. (9) Place all loose ESDS components and equipment into ESD protective containers BEFORE removing a grounding wrist strap. (10) Keep the work station free of any material not required to accomplish the assigned task. (11) Follow established ESD protection rules and procedures. (12) Always use a static control work station, either permanent or portable, when removing ESDS components and equipment from protective packaging. (13) Use only grounded, electrically isolated, and temperature controlled soldering irons that have been rated for use with ESDS components and equipment. Use only hand tools that have conductive or static dissipative handles or grips. Test equipment, such as scopes and meters, must be rated for use around ESDS components and equipment. (14) Avoid exposing ESDS components and equipment to large electromagnetic or electrostatic fields such as transformers or transmitting antennas.

C. Controlling Static Charge Buildup Four basic techniques are employed in ESD control. These are: (1) MINIMIZE THE CHARGE BUILDUP - Minimize electrostatic charge buildup by using conductive or static dissipative flooring and static-dissipative work surfaces. Wear leather shoes, cotton socks, and a grounding ankle strap to dissipate body charge buildup. Wear cotton clothing instead of wool or synthetics. Use an ionized air blower to dissipate charges from nonconductive items. (2) DRAIN OFF THE CHARGE TO GROUND - The human body is a good electrical conductor and for that reason electrostatic charges on the body can be dissipated by skin contact with a grounding device such as a wrist or ankle strap. Always wear a grounding wrist strap when opening ESD containers or handling exposed ESDS components and equipment. (3) NEUTRALIZE THE CHARGE - Nonconductors, such as polystyrene coffee cups, plastic bags, and some clothing develop electrostatic charges that cannot be neutralized by grounding. Ionized air flow will neutralize an electrostatic charge on a nonconductor as long as the ionized air blower puts out both positive and negative ions. (4) MINIMIZE THE EFFECTS OF ELECTROSTATIC FIELDS - The immediate environment surrounding ESDS components and equipment must be free of electrostatic fields or must have suitable static shielding to minimize induced effects from electrostatic fields.

D. Permanent Static Control Workstation A static control work station provides for static-free handling of ESDS components and equipment by diverting, to ground, electrostatic charges on conductive objects (Ref. Figure 201).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL A permanent static control work station consists of the following items: CAUTION: Never wear a grounding wrist strap over clothing. The strap must be in contact with the wearer’s skin to adequately dissipate any electrostatic charge. Under certain conditions, personnel using a grounding wrist strap may need to use a lotion-type skin moisture enhancer to provide a low-resistance connection between the wrist and the wrist strap. (1) GROUNDING WRIST STRAP - Each person that handles ESDS components and equipment must wear a grounding wrist strap to dissipate bodily electrostatic charges. The wrist strap must fit firm against the skin and release quickly in case of an emergency. The wrist strap incorporates a 1-megohm current-limiting resistor, in series with the ground cord, to protect the wearer from electrical shock hazards. (2) STATIC-DISSIPATIVE WORK SURFACE - Conductive mats on the work bench surface are designed to remove electrostatic charges from conductive items placed on the mat. (3) CONDUCTIVE FLOORING - Conductive flooring is used when additional control of ESD is required. To maintain total control over ESD, use conductive chairs, a grounding heel strap, and conductive shoes. Conductive flooring in ESD control areas must be free of all wax or other nonconductive coatings. (4) HARD GROUND CONNECTION - Grounding of the static control work station is accomplished through one or more copper ground rods driven into moist earth to a depth sufficient to provide a low resistance path from the work station to ground. All work station connections to ground are made through a one megohm resistor to protect work station personnel from electrical shock hazards by limiting current flow to ground. NOTE: Check building grounds to ensure that there is no current looping from other nearby grounds. Ensure that the source of current is external and not static. (5) IONIZED AIR BLOWER - The ionized air blower provides a constant flow of positive and negative ions over the work station surface to neutralize electrostatic charges on nonconductive materials in the air flow path. The use of an ionized air blower, in combination with a static control work station, provides additional protection for ESDS components and equipment. Since it is not always possible to eliminate all static charge accumulators (Styrofoam, plastic, etc.) from a work area, the ionized air blower is used to provide additional protection by flooding the work area with balanced negative/positive ionized air. Static charge accumulators should always be kept away from static-free areas, but inadvertent static is difficult to control, especially when developed by such common items as clothing, footwear, combs, and pens. An ionized air blower will help control some of this inadvertent buildup. (6) STATIC DISSIPATIVE SEATING - Chairs used at ESD protected work stations must be conductive, and if padded, must be covered with static dissipative material. (7) CONDUCTIVE CONTAINERS - ESDS devices must be transported in approved containers to prevent ESD damage. These special containers are made of metal or special conductive plastic. Before static-sensitive components and equipment are removed from a static control work station, they must be packaged in containers that provide at least as much protection as that provided by the work station. Conductive boxes, kit trays, and similar types of approved containers provide complete ESD protection to ESDS components and equipment while in transit.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (8) GROUNDING HEEL STRAP - A grounding heel strap can provide additional ESD protection. The heel strap makes contact with the wearer’s skin at the ankle, and extends to the bottom of footwear to make contact with a conductive mat or conductive flooring. The grounding heel strap can be used in combination with a grounding wrist strap to provide maximum ESD protection. (9) ANTISTATIC/CONDUCTIVE CLOTHING - Many types of clothing generate electrostatic charges. To remove some of this buildup, work station personnel should wear outer garments that help dissipate electrostatic charges. Cotton ranks among the best fabrics for antistatic protection. Do not wear synthetic or wool fabrics around ESDS devices, as these fabrics retain electrostatic charges.

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Figure 201 Permanent Static Control Work Station

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E. Portable Static-Control Workstation A portable static control work station provides for static-free handling of ESDS components and equipment during maintenance operations at the airplane. The typical portable work station is available as a field service kit that is used to dissipate electrostatic charges before the charges can damage ESDS components and equipment (Ref. Figure 202). A typical portable static control work station consists of the following items: (1) GROUNDING WRIST STRAP - Each person who handles ESDS components and equipment must wear a grounding wrist strap to dissipate bodily electrostatic charges. The wrist strap must fit firm against the skin and should release quickly in case of emergency. The wrist strap incorporates a 1-megohm current-limiting resistor, in series with the ground cord, to protect the wearer from electrical shock hazards. (2) STATIC-DISSIPATIVE WORK SURFACE - A conductive mat is an integral part of the portable work station, and is designed to remove electrostatic charges from conductive items when those items contact the mat. (3) HARD GROUND CONNECTION - Ground the portable work station to the airframe or to a common ground as shown. All portable work station connections to ground are made through 1-megohm current-limiting resistors to protect maintenance personnel from electrical shock hazards.

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Figure 202 Portable Static Control Work Station

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F. Humidity and Dust Effects on ESDS Components and Equipment Humidity is a factor in the control of ESD. The lower the humidity, the greater the chance of damage to ESDS components and equipment. Humidity at the work station should be maintained between 30 and 65 percent. Repair of ESDS circuit boards, including replacement of ESDS components, should be performed in a dust-free environment.

G. Packaging of ESDS Components and Equipment All ESDS components and equipment require special ESD protective packaging. Seal all ESDS packages with an appropriate cautionary label as shown (Ref. Figures 203, 204, 205, 206, 207 and 208). CAUTION: Do not use clips or staples when sealing any ESDS package. Do not use carbon-filled, conductive bags. Remove ESDS components and equipment from protective, static-shielded containers only at a static-control work station after attaching a grounding wrist strip and verifying that ESD producing items are not on the static-dissipative work surface. ESD protective packaging requirements, unless otherwise defined by specification, shall conform to the following: •

Class 1 - Package in multi-layer conductive type bags consisting of an inner and outer layer of antistatic (surface resistivity of 109 to 1014 ohms per square inch) or static dissipative (surface resistivity of 105 to 109 ohms per square inch) material with a middle layer of conductive material (surface resistivity of 10 ohms or less).

•

Class 2 - Package in a static dissipative material possessing a surface resistivity of 105 to 109 ohms per square inch. Materials specified for Class 1 may also be used.

•

Class 3 - Package in an antistatic material possessing a surface resistivity of 109 to 1014 ohms per square inch.

Place all ESDS devices in approved static shielding containers before packing in shipper’s normal exterior containers. Use antistatic cushioning or fill materials. Do not use static generating materials, such as polyethylene, Styrofoam, or paper. Antistatic packaging is generally pink or blue in color. The material differs from common plastic in that an antistatic compound is incorporated into the material during the manufacturing process. This type of packaging DOES NOT provide static shielding, and is generally used to package instruction sheets, data sheets, and other non-ESDS materials prior to introduction into a static-free environment. All non-ESDS items, that are to enter an ESD work station, require repackaging in antistatic materials. Conductive static-shielding packaging differs from antistatic packaging, in that it has the ability to shield the devices, contained within from external static charges. Conductive static-shielding packaging is available in the form of bags and rigid containers.

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Figure 203 Equipment Enclosure Cautionary Placards

Figure 204 ESDS Drawing Note

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H. Marking of ESDS Components and Equipment All ESDS components and equipment should be marked appropriately with an ESDS symbol as shown (Ref. Figure 205). NOTE: ESDS symbols (circle with arrows pointing into the circle from equidistant positions or a hand inside of a triangle with an angling bar across the triangle) are yellow on a black background or black on a yellow background. Mark unit containers with the ESDS caution label on the outside of the package. Mark exterior containers with an ESD caution label as shown (Ref. Figure 208). Apply marks directly to each ESDS printed circuit board, assembly cover, equipment enclosure, or access door that would expose ESDS devices if removed. Mark appropriately using decal transfer, stencil, silk screen, or any other method meeting permanent legibility requirements. Display ESDS symbols in a prominent package location to alert all personnel to the presence of ESDS devices and equipment. The ESDS symbol should be at least 1/4 inch in diameter. ESDS symbols that are attached to circuit boards should contrast with the circuit board base color. Enclosures that contain ESDS circuit boards should be identified by bright orange paint on the outer face of the enclosure.

I. Storage and Transit of ESDS Components and Equipment CAUTION: Never use ordinary plastic containers or packing materials when transporting ESDS components or equipment. When preparing ESDS devices for shipment, ensure all assemblies and equipment have been protected against ESD through appropriate handling at static-controlled work stations. ESDS packages, which have been properly enclosed in protective packages, require proper storage and transfer in conductive static-dissipative, or static-free containers. Shipping information and other instructions, accompanying ESD-protected packages, shall be contained in anti-static materials. ESDS components, that are received in damaged or opened packing containers, are not acceptable, and should be returned for replacement.

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Figure 205 ESD Symbols

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Figure 206 ESDS Unit Container Notice

Figure 207 Protective Container Notice

Page 212 Nov 1/09

20-00-04

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 208 In-House Storage Container Label

20-00-04

Page 213 Nov 1/09

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

STANDARD PRACTICES - AIRFRAME TUBING, HOSE AND FITTINGS DESCRIPTION AND OPERATION

20-00-05 00

1. GENERAL NOTE: Refer to Chapter 29 of the Model 1900 Airliner Series Component Maintenance Manual for procedures to install Cryofit fittings for tubing repair. This chapter contains information to remove, maintain and install hose and tube assemblies and fittings. Although all hoses and tubes may not be specifically identified herein, the basic maintenance practices normally apply. Any handling and installation of individual system hoses, tubes and fittings is identified in the appropriate system chapter. The majority of tube assemblies used in the airplane are aluminum machine or steel machine formed tubing assemblies. Hoses are used in areas of the airplane where a flexible line is more suitable for installations and freedom of movement is necessary.

20-00-05

Page 1 Nov 1/09

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

STANDARD PRACTICES - AIRFRAME TUBING, HOSE AND FITTINGS MAINTENANCE PRACTICES

20-00-05 200200

1. GENERAL Observe all WARNINGS, CAUTIONS and NOTES throughout this maintenance manual when performing maintenance, repair or servicing on any fluid or pneumatic operated system. WARNING: Never perform maintenance on tubing, hoses or fittings while the system is under pressure. Verify that systems which operate using fluids or pneumatics under pressure are fully depressurized before opening or disconnecting a tubing assembly or hose. NOTE: Refer to Chapter 29 of the Model 1900 Airliner Series Component Maintenance Manual for procedures to install Cryofit or Permaswage fittings.

A. Practices The following list of maintenance practices is provided as an aid for handling, removal, installation and repairing of tubing and hoses. (1) Cap or plug all disconnected tubing and hose assembles and fittings immediately to prevent contamination of the system. (2) Visually check for cleanliness, evidence of contamination and obstructions prior to recondition of tube or hose assemblies. (3) Any hose and tube assemblies that did not have protective covers installed must be cleaned and checked for obstruction prior to installation. (4) When connecting tube assemblies, do not force the tube assembly to the installed position. If a mismatch between the male and female fittings should result, check for allowable mismatch (Ref. Figure 201). (5) Never stretch a hose to make a connection. (6) The hose material must be compatible with the applicable system fluids. Substitution of a hose material that is not compatible with the system fluid will contaminate the system.

2. TUBE AND HOSE ASSEMBLIES A. Removal (1) Relieve all system pressure. CAUTION: Failure to use a back-up wrench when loosening or tightening hoses and tubing to fittings may damage the hoses, tubing and fittings. (2) Disconnect both ends of the hose or tube assembly and immediately cap or plug the tube or hose ends and fittings. (3) Remove all clamps securing the hose or tube assembly. (4) Remove the tube or hose assembly and tag identify both ends to aid in reinstallation.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

B. Hose Assembly Installation NOTE: If a new hose assembly is to be installed, the hose assembly must be clean, the correct length, clear of obstructions and made of material compatible with the system fluid. (1) Observe the maintenance practices outlined under the heading TUBE AND HOSE ASSEMBLIES AND FITTINGS - MAINTENANCE PRACTICES in this section. (2) Connect the b-nuts of the hose assembly to the proper fittings. CAUTION: Failure to use a back-up wrench when loosening or tightening hoses to fittings may damage the hoses and fittings. (3) Torque the b-nuts to the fittings using the torque specified in Table 201. (4) After torquing the b-nuts, inspect the hose to ensure that the hose is not under tension and that no indication of twisting is present. (5) Inspect the hose for proper length. (6) Inspect the hose for freedom to expand and contract. (7) Inspect the hose for clearance to all structure. If inadequate clearance exists between the hose and structure, protection must be provided for the hose to prevent damage from chafing.

C. Tube Installation NOTE: If a new tube assembly is to be installed, the tube assembly must be clean, the correct length, clear of obstructions and manufactured of the correct material. (1) Observe the maintenance practices outlined under the heading TUBING, HOSE AND FITTINGS MAINTENANCE PRACTICES in this section. (2) Inspect the tube for damage, particularly at tube ends, fittings and bends. Damaged tube assemblies should be replaced or repaired. (3) Make certain that the fittings are properly installed before connection of the tube assembly. (4) Check alignment and fit of the tube assembly as follows before installation: (a) Place the tube assembly in the proper position and tighten one coupling nut at one end of the tube assembly. (b) The opposite end of the tube must be within two degrees of parallel with the fitting (Ref. Figure 201). (c) The free tubing end must be aligned within 1/32 inch of the fitting per every 10 inches of tube length (Ref. Figure 201). CAUTION: Failure to use a back-up wrench when loosening or tightening tubing to fittings may damage the tubing and fittings. (5) Install the tube assembly on the fittings and tighten the b-nuts to the torque values specified in Table 201.

Page 202 Nov 1/13

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (6) If necessary, apply the proper anti-seize compound to fittings. Any time a high temperature (bleed air) stainless-to-stainless threaded attach point is disconnected, apply lubricant (69, Table 1, Chapter 91-00-00) or equivalent prior to connecting. Table 201 Flared Fitting Torque Table (Inch-Pounds) Hose Size

Tubing O.D. (Inches)

Aluminum Tubing Flare Min. Max.

Steel Tubing Flare Min. Max.

Aluminum Tubing Flareless Min. Max.

Steel Tubing Flareless Min. Max.

Oxygen Line Fitting (Aluminum) Min. Max.

Hose End Fitting Min. Max.

-3

3/16

--- ---

90 to 100

75 to 90

90 to 100

--- ---

70 to 100

-4

1/4

40 to 65

135 to 150

80 to 100

135 to 150

--- ---

70 to 120

-5

5/16

60 to 80

180 to 200

100 to 130

180 to 200

100 125

85 to 180

-6

3/8

75 to 125

270 to 300

100 to 130

270 to 300

--- ---

100 to 250

-8

1/2

150 to 250

450 to 500

200 to 240

450 to 500

--- ---

210 to 420

-10

5/8

200 to 350

700 to 800

360 to 400

700 to 800

--- ---

300 to 480

-12

3/4

300 to 500

900 to 1150

390 to 430

900 to 1150

--- ---

500 to 850

-16

1

500 to 700

1200 to 1400

600 to 900

1200 to 1400

--- ---

700 to 1150

-20

1 1/4

600 to 900

1300 to 1450

600 to 900

1300 to 1450

--- ---

--- ---

-24

1 1/2

600 to 900

1350 to 1500

600 to 900

1350 to 1500

--- ---

--- ---

3. FLUID LINE FITTING A. Installation (1) Lubricate the male threads of the fitting, backup ring and packing sparingly with the system fluid or petrolatum (76, Table 1, Chapter 91-00-00). (2) Install the nut (AN6289) on the fitting until the nut is clear of the thread relief. (3) Install the teflon backup ring in the counterbore of the nut. (4) Install the packing on the thread relief. NOTE: The packing must be compatible with the system fluid. (5) Turn the nut down until the packing is pushed firmly against the lower threaded section of the fitting. (6) Install the fitting into the boss with the nut turning with the fitting until the packing contacts the boss. NOTE: This point can be detected by a sudden increase in torque. (7) Holding the nut with a wrench to prevent it from turning, rotate the fitting in an additional 1 1/2 turns. Position the fitting in the proper direction by turning in no more than one additional turn.

20-00-05

Page 203 Nov 1/13

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (8) Holding the fitting, turn the nut down tightly against the boss. Slight extrusion of the ring around the backup ring is acceptable.

Figure 201 Tubing Installation Mismatch

4. NONPOSITIONING TYPE FITTING A. Installation (1) Lubricate the packing with the system fluid or petrolatum (76, Table 1, Chapter 91-00-00). (2) Install the packing in the fitting thread relief. (3) Thread the fitting into the boss until it bottoms tightly on the boss. (4) Tighten the fitting to the specified torque value.

5. PIPE THREAD FITTING A. Installation Install pipe fittings as follows: (1) Apply teflon tape to the threads as follows: (a) Start tape at or close to narrow end of threads. (b) Wrap the tape around the fitting in the direction of the threads. Wrap clockwise for right hand threaded fittings. Wrap counterclockwise for left hand threaded fittings.

Page 204 Nov 1/13

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (c) Apply tension to the tape to conform the tape to the shape of the threads. (d) The tape should overlap the previous wrap of tape up to one-half inch to seal pipe thread fittings up to two inches in diameter. (2) Thread the fitting into the boss and tighten until it bottoms tightly on the boss.

6. TUBE DAMAGE A. Limits NOTE: Nicks and scratches not exceeding the following limitations may be repaired by polishing out the damaged area, using fine grade of emery cloth and oil. Finish polishing with crocus cloth and oil. Flush and clean all grit from line assembly. (1) Replace steel tubes which have nicks or scratches deeper than 10 percent of tubing wall thickness. (2) Replace any aluminum tube which has nicks or scratches deeper than 20 percent of the tube wall thickness. (3) Replace any tubes which have dents deeper than 5 percent of the tube outside diameter.

7. TUBE MANUFACTURING A new tube must be manufactured from the correct material (Ref. Table 203) and to the specific application for which it is to be used. Before cutting the tube to length, make sure it is long enough to make all bends and any forming that must be made at the ends. Prepare the tubing ends to match the connecting fittings. All bends must be made with an appropriate bending tool and to the limits specified (Ref. Figure 202 and Table 202). TUBE OUTSIDE DIAMETER (O.D.)

TUBE CENTERLINE

BEND RADIUS IS MEASURED TO TUBE CENTERLINE TUBE INSIDE DIAMETER (I.D.)

UC20B 022678AA.AI

Figure 202 Tube Bending

20-00-05

Page 205 Nov 1/13

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Table 202 TUBING BENDING LIMITS RECOMMENDED BEND RADII

TUBE OUTSIDE DIAMETER (D)

3D

ADDITIONAL BEND RADII 4D

6D

INCH

MILLIMETER

INCH

MILLIMETER

INCH

MILLIMETER

INCH

MILLIMETER

1/8

3.175

0.375

9.525

0.500

12.700

0.750

19.050

3/16

4.762

0.563

14.286

0.750

19.048

1.125

28.572

1/4

6.350

0.750

19.050

1.000

25.400

1.5

38.100

5/16

7.937

0.938

23.811

1.250

31.748

1.875

47.622

3/8

9.525

1.125

28.575

1.500

38.100

2.250

57.150

7/16

11.112

1.312

33.336

1.750

44.448

2.625

66.672

1/2

12.700

1.500

38.100

2.000

50.800

3.000

76.200

5/8

15.875

1.875

47.625

2.500

63.500

3.750

95.250

3/4

19.050

2.250

57.150

3.000

76.200

4.500

114.300

7/8

22.225

2.625

66.675

3.500

88.900

5.250

133.350

1

25.400

3.000

76.200

4.000

101.600

6.000

152.400

1 1/8

28.575

3.375

85.725

4.500

114.300

6.750

171.450

1 1/4

31.750

3.750

95.250

5.000

127.000

7.500

190.500

1 3/8

34.925

4.125

104.775

5.500

139.700

8.250

209.550

1 1/2

38.100

4.500

114.300

6.000

152.400

9.000

228.600

1 5/8

41.275

4.875

123.825

6.500

165.100

9.750

247.650

1 3/4

44.450

5.250

133.350

7.000

177.800

10.500

266.700

1 7/8

47.625

5.625

142.875

7.500

189.500

11.250

285.750

2

50.800

6.000

152.400

8.000

203.200

12.000

304.800

2 1/4

57.150

6.750

171.450

9.000

228.600

13.500

342.900

2 1/2

63.500

7.500

190.500

10.000

254.000

15.000

381.000

3

76.200

9.000

228.600

12.000

304.800

18.000

457.200

Page 206 Nov 1/13

20-00-05

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Table 203 COMMON TUBING PART NUMBER

DIMENSIONS

MATERIAL

SPECIFICATION

102-540026-3

1/4 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4

114-580134-1

1/4 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4

114-580020-1

3/8 O.D. X 0.035

5052-0 ALUM TUBE

AMS 4071

114-580021-1

3/8 O.D. X 0.035

5052-0 ALUM TUBE

AMS 4071

114-580023-1

1/4 O.D. X 0.035

5052-0 ALUM TUBE

AMS 4071

114-580024-1

1/4 O.D. X 0.035

5052-0 ALUM TUBE

AMS 4071

114-580025-1

1/4 O.D. X 0.035

5052-0 ALUM TUBE

AMS 4071

114-580032-1

1/2 O.D. X 0.035

SEAMLESS CRES TUBE TYPE 1

MIL-T-6845

114-580033-1

3/8 O.D. X 0.028

SEAMLESS CRES TUBE TYPE 1

MIL-T-6845

114-580034-1

3/8 O.D. X 0.028

SEAMLESS CRES TUBE TYPE 1

MIL-T-6845

114-580035-1

3/8 O.D. X 0.028

SEAMLESS CRES TUBE TYPE 1

MIL-T-6845

114-580038-1

1/4 O.D. X 0.028

SEAMLESS CRES TUBE TYPE 1

MIL-T-6845

114-580039-1

1/4 O.D. X 0.028

SEAMLESS CRES TUBE TYPE 1

MIL-T-6845

114-580040-1

1/4 O.D. X 0.028

SEAMLESS CRES TUBE TYPE 1

MIL-T-6845

114-580041-1

1/4 O.D. X 0.035

5052-0 ALUM TUBE

AMS 4071

114-580042-1

1/2 O.D. X 0.035

SEAMLESS CRES TUBE TYPE 1

MIL-T-6845

114-580044-1

3/8 O.D. X 0.028

SEAMLESS CRES TUBE TYPE 1

MIL-T-6845

114-580045-1

3/8 O.D. X 0.028

SEAMLESS CRES TUBE TYPE 1

MIL-T-6845

114-580046-1

1/4 O.D. X 0.028

SEAMLESS CRES TUBE TYPE 1

MIL-T-6845

114-580047-1

1/2 O.D. X 0.035

SEAMLESS CRES TUBE TYPE 1

MIL-T-6845

114-580048-1

3/8 O.D. X 0.028

SEAMLESS CRES TUBE TYPE 1

MIL-T-6845

114-580049-1

1/2 O.D. X 0.035

6061-T6 ALUM TUBE

MIL-T-7081

114-580050-1

3/8 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4

114-580051-1

1/4 O.D. X 0.028

SEAMLESS CRES TUBE TYPE 1

MIL-T-6845

114-580052-1

1/4 O.D. X 0.028

SEAMLESS CRES TUBE TYPE 1

MIL-T-6845

114-580053-1

1/4 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4

114-580054-1

1/4 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4

114-580055-1

1/4 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4

114-580056-1

1/4 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4

114-580057-1

1/4 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4

114-580058-1

1/4 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4

114-580059-1

1/4 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4

20-00-05

Page 207 Nov 1/13

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Table 203 COMMON TUBING (Continued) PART NUMBER

DIMENSIONS

MATERIAL

SPECIFICATION

114-580060-1

1/4 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4

114-580062-1

1/2 O.D. X 0.035

SEAMLESS CRES TUBE TYPE 1

MIL-T-6845

114-580063-1

1/4 O.D. X 0.028

SEAMLESS CRES TUBE TYPE 1

MIL-T-6845

114-580089-1

1/4 O.D. X 0.035

5052-0 ALUM TUBE

AMS 4071

114-580093-1

1/4 O.D. X 0.035

5052-0 ALUM TUBE

AMS 4071

114-580101-1

1/4 O.D. X 0.035

5052-0 ALUM TUBE

AMS 4071

114-580102-1

1/4 O.D. X 0.035

5052-0 ALUM TUBE

AMS 4071

114-580113-1

1/4 O.D. X 0.035

5052-0 ALUM TUBE

AMS 4071

114-580114-1

1/4 O.D. X 0.035

5052-0 ALUM TUBE

AMS 4071

114-580135-1

1/4 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4

114-580155-1

1/4 O.D. X 0.035

5052-0 ALUM TUBE

AMS 4071

114-580156-1

1/4 O.D. X 0.035

5052-0 ALUM TUBE

AMS 4071

114-580203-1

3/8 O.D. X 0.035

5052-0 ALUM TUBE

AMS 4071

114-580204-1

1/4 O.D. X 0.035

5052-0 ALUM TUBE

AMS 4071

114-580214-1

3/8 O.D. X 0.028

SEAMLESS CRES TUBE TYPE 1

MIL-T-6845

114-580215-1

3/8 O.D. X 0.028

SEAMLESS CRES TUBE TYPE 1

MIL-T-6845

114-580216-1

3/8 O.D. X 0.028

SEAMLESS CRES TUBE TYPE 1

MIL-T-6845

114-580220-1

1/4 O.D. X 0.035

5052-0 ALUM TUBE

AMS 4071

114-580252-1

1/4 O.D. X 0.035

5052-0 ALUM TUBE

AMS 4071

114-580281-1

1/4 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4

114-580303-1

1/4 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4

114-580333-1

3/8 O.D. X 0.028

SEAMLESS CRES TUBE TYPE 1

MIL-T-6845

114-970026-1

5/8 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4

114-970026-3

3/4 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4

114-970031-1

5/8 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4

114-970033-1

5/8 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4

114-970035-1

3/4 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4

114-970036-1

3/4 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4

114-970037-1

5/8 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4

114-970042-1

3/4 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4

114-970043-1

3/4 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4

114-970044-1

5/8 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4

Page 208 Nov 1/13

20-00-05

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Table 203 COMMON TUBING (Continued) PART NUMBER

DIMENSIONS

MATERIAL

SPECIFICATION

114-970045-1

5/8 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4

114-970049-1

1/2 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4

114-970053-1

5/8 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4

114-970054-1

5/8 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4

114-970062-3

5/8 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4

114-970071-1

5/8 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4

114-970085-1

1/2 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4

114-970086-1

1/4 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4

114-970092-1

1/2 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4

114-970096-1

5/8 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4 ANNEALED

114-970105-1

5/8 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4

114-970107-1

5/8 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4

114-970109-1

5/8 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4

114-970111-1

5/8 O.D. X 0.028

SEAMLESS CRES TUBE TYPE 1

MIL-T-8808

117-970051-1

3/8 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4

118-920000-10

3/4 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4

118-920000-100

5/8 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4

118-920000-101

5/8 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4

118-920000-102

5/8 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4

118-920000-103

5/8 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4

118-920000-105

5/8 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4

118-920000-13

3/4 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4

118-920000-14

3/4 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4

118-920000-15

5/8 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4

118-920000-16

5/8 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4

118-920000-19

1/4 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4

118-920000-20

1/4 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4

118-920000-21

1/4 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4

118-920000-22

1/4 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4

118-920000-23

5/8 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4

118-920000-27

5/8 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4

118-920000-29

5/8 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4

20-00-05

Page 209 Nov 1/13

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Table 203 COMMON TUBING (Continued) PART NUMBER

DIMENSIONS

MATERIAL

SPECIFICATION

118-920000-3

3/4 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4

118-920000-31

5/8 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4

118-920000-33

5/8 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4

118-920000-35

5/8 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4

118-920000-5

3/4 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4

118-920000-53

5/8 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4

118-920000-55

1/4 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4

118-920000-56

1/4 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4

118-920000-6

3/4 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4

118-920000-71

3/4 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4

118-920000-72

3/4 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4

118-920000-75

5/8 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4

118-920000-89

5/8 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4

118-920000-9

3/4 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4

118-920000-91

5/8 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4

118-920000-93

1/2 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4

118-920000-95

1/2 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4

118-920000-97

1/2 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4

118-920000-98

1/2 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4

118-920000-99

5/8 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4

129-580042-1

1/4 O.D. X 0.035

6061-T6 ALUM TUBE

MIL-T-7081

129-580043-1

1/4 O.D. X 0.035

6061-T6 ALUM TUBE

MIL-T-7081

129-580044-1

1/4 O.D. X 0.035

6061-T6 ALUM TUBE

MIL-T-7081

129-580045-1

1/4 O.D. X 0.035

6061-T6 ALUM TUBE

MIL-T-7081

129-580055-1

1/4 O.D. X 0.035

5052-0 ALUM TUBE

AMS-4071

129-580056-1

1/4 O.D. X 0.035

5052-0 ALUM TUBE

AMS-4071

129-580061-1

1/4 O.D. X 0.035

5052-0 ALUM TUBE

AMS 4071

129-580062-1

1/4 O.D. X 0.035

5052-0 ALUM TUBE

AMS 4071

129-580070-1

1/4 O.D. X 0.035

6061-T6 ALUM TUBE

MIL-T-7081

129-580075-1

1/4 O.D. X 0.035

5052-0 ALUM TUBE

AMS-4071

129-580076-1

1/4 O.D. X 0.035

5052-0 ALUM TUBE

AMS-4071

129-970000-5

1/4 I.D.

SYN RUBBER HOSE

MIL-H-5593-4

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Table 203 COMMON TUBING (Continued) PART NUMBER

DIMENSIONS

MATERIAL

SPECIFICATION

129-970031-1

5/8 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4

129-970032-1

5/8 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4

129-970033-1

5/8 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4

129-970034-1

5/8 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4

129-970035-5

5/8 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4

129-970035-7

1/2 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4

129-970036-5

5/8 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4

129-970036-7

1/2 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4

129-970038-1

5/8 O.D. X 0.028

SEAMLESS CRES TUBE TYPE 1

MIL-T-8808

129-970039-1

5/8 O.D. X 0.028

SEAMLESS CRES TUBE TYPE 1

MIL-T-8808

129-970040-1

5/8 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4

129-970043-1

5/8 O.D. X 0.028

SEAMLESS CRES TUBE TYPE 1

MIL-T-8808

129-970044-1

5/8 O.D. X 0.028

SEAMLESS CRES TUBE TYPE 1

MIL-T-8808

129-970045-1

5/8 O.D. X 0.028

SEAMLESS CRES TUBE TYPE 1

MIL-T-8808

129-970047-1

5/8 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4

129-970048-1

5/8 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4

129-970051-1

5/8 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4

129-970052-1

5/8 O.D. X 0.035

5052-0 ALUM TUBE

WW-T-700/4

131066-2-0030

1/4 I.D. X 3/8 O.D.

TYGON TUBE TYPE 4040

131066-2-0075

1/4 I.D. X 3/8 O.D.

TYGON TUBE TYPE F-4040

20-00-05

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

8. CONICAL SEAL A. Installation NOTE: Conical seals are used for in-service repair of minor leaks providing the tube end and matching cone are in good condition. Seals can also be used for permanent repairs but must be replaced if the tube end and cone are damaged for any reason. The seal must be made of the same material as the tube it’s being used on. (1) Install conical seal (2) into flare (3) of tubing (5) or onto fitting (1) (Ref. Figure 203). CAUTION: Failure to use a back-up wrench when loosening or tightening tubing to fittings may damage the tubing and fittings. (2) Install tube (5) to fitting (1) and torque b-nut (4) (Ref. Table 204).

Figure 203 Conical Seal Installation

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

AP50

A

12

L

S

OPTIONAL FINISH CODE 'S' INDICATES SILVER PLATING OPTIONAL CLEANING CODE 'L' INDICATES PARTS REQUIRED LOX-CLEANED DASH NO. PER TUBE SIZE '12' = 3/4 IN. MATERIAL CODE 'A' = ALUMINUM ALLOY BASIC PART NUMBER

UC20B 102151AA.AI

Figure 204 Conical Seal Part Number Breakdown Table 204 Recommended Torque Values for Conical Seals Tubing O.D.

Seal Dash No.

All Alum. System Except 6061-T6

All Alum. 6061-T6 System

All CRES and Steel System

All CRES and Steel System

All CRES and Steel System

Alum. B-Nut and Sleeve CRES Fitting

CRES and Steel B-Nut and Sleeve Alum. Fitting

Alum. B-Nut Stainless Sleeve Alum. Fitting

Alum. Seal

Alum. Seal

Copper Seal

Nickel Seal

Stainless Steel Seal

Tin Plated Alum. Seal

Tin Plated Copper Seal

Tin Plated Copper Seal

1/8

-2

30 to 50

35 to 65

80 to 90

90 to 100

90 to 100

35 to 65

70 to 80

70 to 80

3/16

-3

35 to 60

35 to 70

100 to 110

110 to 125

110 to 125

35 to 70

90 to 100

90 to 100

1/4

-4

40 to 65

70 to 120

150 to 165

165 to 190

165 to 190

70 to 120

135 to 150

135 to 150

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Table 204 Recommended Torque Values for Conical Seals (Continued) Tubing O.D.

Seal Dash No.

All Alum. System Except 6061-T6

All Alum. 6061-T6 System

All CRES and Steel System

All CRES and Steel System

All CRES and Steel System

Alum. B-Nut and Sleeve CRES Fitting

CRES and Steel B-Nut and Sleeve Alum. Fitting

Alum. B-Nut Stainless Sleeve Alum. Fitting

Alum. Seal

Alum. Seal

Copper Seal

Nickel Seal

Stainless Steel Seal

Tin Plated Alum. Seal

Tin Plated Copper Seal

Tin Plated Copper Seal

5/16

-5

60 to 80

80 to 130

200 to 220

225 to 250

225 to 250

80 to 130

180 to 200

180 to 200

3/8

-6

75 to 125

130 to 180

300 to 330

335 to 375

335 to 375

130 to 180

270 to 300

270 to 300

1/2

-8

150 to 250

300 to 400

500 to 550

575 to 625

575 to 625

300 to 400

450 to 500

450 to 500

5/8

-10

200 to 350

430 to 550

710 to 770

810 to 875

810 to 875

430 to 550

650 to 700

650 to 700

3/4

-12

300 to 500

650 to 800

990 to 1100

1125 to 1250

1125 to 1250

650 to 800

900 to 1000

900 to 1000

1

-16

500 to 700

900 to 1100

1300 to 1550

1500 to 1750

1500 to 1750

900 to 1100

1200 to 1400

1200 to 1400

1 1/4

-20

600 to 900

1200 to 1450

1650 to 1950

1875 to 2250

1875 to 2250

1200 to 1450

1500 to 1800

1500 to 1800

1 1/2

-24

600 to 900

1550 to 1850

2200 to 2500

2500 to 2850

2500 to 2850

1550 to 1850

2000 to 2300

2000 to 2300

1 3/4

-28

700 to 1000

2000 to 2350

2800 to 3150

3250 to 3600

3250 to 3600

2000 to 2350

2600 to 2900

2600 to 2900

2

-32

800 to 1100

2500 to 2900

3500 to 3950

4000 to 4500

4000 to 4500

2500 to 2900

3200 to 3600

3200 to 3600

Torque values in inch-pounds

Page 214 Nov 1/13

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

STANDARD PRACTICES - AIRFRAME TORQUE WRENCHES MAINTENANCE PRACTICES

20-01-00 200200

1. PROCEDURES When a torque wrench and adapter is used (Ref. Figure 201), compensation must be made for the extra leverage gained. New indicator readings must be calculated before the wrench is used. To figure the desired lower readings which will actually give the torque specified, use the following formula: Wrench length x Specific Torque = Desired Torque

Length of wrench + adaptor

Example:

LxT D=

L+A

D= L= A= T=

Desired reading Length of torque wrench Adaptor length Torque

D= L= A= T=

? 33 inches 11 inches 5000 inch-pounds

33 x 5,000 =

33 + 11

165,000 =

44

= 3,750 in-Ibs

An acceptable method of checking the torque, if a torque wrench is not available, is to attach a spring scale to a conventional flex or T-handle inserted in an adapter (Ref. Figure 202). Force should be applied in a direction perpendicular to an imaginary line extending from the center of the bolt through the spring scale attaching point. To calculate the force in pounds (scale reading) required to obtain the specified torque, divide the torque in inch-pounds by the distance in inches between the center of the bolt and the scale attaching point. For example, if the specified torque is 5,000 inch-pounds and the distance is 25 inches, a pull of 200 pounds must be applied. Bolts to be torqued must be clean and free of all lubricants; otherwise, loss of normal friction allowed for establishing the torque values may result in overtorquing of the bolt.

20-01-00

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

When a torque wrench adapter is used, the length of the adapter must be added to the length of the flex or T handle wrench and a value calculated for that particular combination. The following is a typical example in finding a desired value. Effective length of flex or “T” Handle wrench.....................................

12 inches

Length of adapter.................................

3 inches

Total length..........................................

15 inches

Desired torque on bolt........................... 2,000 inch-pounds

2000 inch-pounds 15 inches

= 133.3 pounds (scale reading)

Figure 201 Torque Wrench and Adaptor

Page 202 Nov 1/09

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 202 Computing Torque with Spring Scale

20-01-00

Page 203 Nov 1/09

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

STANDARD PRACTICES - AIRFRAME LEADING EDGE EROSION PROTECTION DESCRIPTION AND OPERATION

20-04-00 00

1. GENERAL It is imperative to maintain the integrity of all leading edges. Erosion can occur on the leading edges over time and negatively effect the airplane’s lift and performance.

20-04-00

Page 1 Nov 1/09

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

STANDARD PRACTICES - AIRFRAME LEADING EDGE EROSION PROTECTION MAINTENANCE PRACTICES

200200

1. PROCEDURES A. Leading Edge Erosion Protection NOTE: LJF801A 1/2 sealant has good adhesion to rubber deicers. Uralite 3149 does not have good adhesion to rubber, however, both are excellent for erosion protection. Paint all areas to a thickness of 0.20-inch, making sure it is aerodynamically smooth so as not to disrupt airflow. (1) Coat areas indicated using sealant (162 or 163, Table 1, Chapter 91-00-00) (Ref. Figure 201). (2) Install abrasion resistance film (164, Table 1, Chapter 91-00-00) on the vertical stabilizer leading edge as shown.

Figure 201 (Sheet 1 of 3) Leading Edge Erosion Protection

20-04-00

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 201 (Sheet 2 of 3) Leading Edge Erosion Protection

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20-04-00

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 201 (Sheet 3 of 3) Leading Edge Erosion Protection

20-04-00

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

STANDARD PRACTICES BEARINGS MAINTENANCE PRACTICES

20-05-00 200200

1. PROCEDURES A. Hydraulic Press Bearing Removal (1) Remove the bearing housing from the airplane. (2) Place two supports on the hydraulic press table under the bearing housing as shown. The supports should be at least 1/2 inch thicker than the bearing width (Ref. Figure 201). (3) Center a bearing removal and installation tool on the bearing outer race. The bearing and installation tool should be approximately 1/8 inch smaller than the outside diameter of the bearing outer race. CAUTION: The hydraulic press plunger and bearing removal and installation tool should remain in direct alignment with the bearing being removed at all times. (4) Align the bearing and the bearing removal and installation tool with the hydraulic press plunger and apply pressure to force the bearing from the bearing housing.

B. Mechanical Press Bearing Removal (1) Remove the bearing housing from the airplane. (2) Center a bearing removal socket, with an inner diameter larger than the bearing outer race and 1/2 inch deeper than the bearing width, on the bearing housing (Ref. Figure 201). (3) Center a bearing removal socket smaller in diameter than the bearing outer race. (4) Install a washer and bolt through one of the sockets, through the center of the bearing and then through the opposite socket as shown. (5) Install a washer and nut on the bolt threads. (6) Tighten the nut on the bolt until the pressure is sufficient to release the bearing from the bearing housing.

C. Bearing Housing Inspection Inspect the bearing housing for any grooves, cracks, warpage or hole elongation. The bearing housing bearing contact surface must be smooth and uniform.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

D. Bearing Installation Using Retaining Compound CAUTION: When cleaning bearing surfaces never allow solvent to enter the bearing. Never touch the bearing or bearing housing surfaces with bare hands. Use a clean cloth to cover the bearing to prevent contamination after they have been cleaned. (1) Clean the outer surface of the bearing race with solvent (14, Table 1, 91-00-00) and wipe dry. (2) Coat the surfaces where a retaining compound is to be applied with primer (92, Table 1, 91-00-00). This includes the bearing outer surface, bearing housing mating surface and the bearing housing retention flange if applicable. CAUTION: Ensure that no primer is applied to the bearing oil grooves or lubrication ports. NOTE: All cadmium, zinc, corrosion-resistant and anodized steel, including plastic items, must be primed to assure proper adhesion of the retaining compound (194, Table 1, 91-00-00). (3) Allow the primer (92, Table 1, 91-00-00) to air dry for at least 30 minutes at room temperature. NOTE: The retaining compound (194, Table 1, 91-00-00) may be applied before or after bearing installation in the bearing housing. (4) Apply a thin coat of the retaining compound (194, Table 1, 91-00-00) to the bearing and the bearing housing mating surfaces where the primer (92, Table 1, 91-00-00) was applied. (5) Center the bearing on the bearing housing. (6) Using the hydraulic or mechanical pressure method shown, apply pressure to the bearing until it is firmly seated in the bearing housing. Pressure on the bearing must be applied in direct alignment to the bearing housing for the bearing to seat properly (Ref. Figure 201). (7) Apply retaining compound (194, Table 1, 91-00-00) to the area between the bearing and the bearing housing. (8) The retaining compound must cure before the bearing is put into service. Curing may be accomplished with one of the following: (a) Allow the bearing and bearing housing to remain at room temperature for 24 hours without any movement of the parts. (b) Heat the bearing and bearing housing to 275° ± 10°F and maintain that temperature for 15 minutes only. (c) After the bearing has been secured in the housing by the use of a retaining compound, lubricate with the proper lubrication and install in the airplane.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

E. Bearing Installation by Staking (1) Center the bearing on the bearing housing (Ref. Figure 202). (2) Using the hydraulic or mechanical pressure method shown, apply pressure to the bearing until it is firmly seated in the bearing housing. Pressure must be applied in direct alignment to the bearing housing for the bearing to seat properly (Ref. Figure 201). (3) Place the bearing and bearing housing on two supports, if both sides of the bearing are to be staked. The inner bearing race must not touch the supports (Ref. Figure 202). (4) If the bearing housing was previously staked, the new stakes should be centered between the existing stakes. If a new bearing housing is being used, the stake pattern should be the same as the one on the old bearing housing. NOTE: When a ring stake is used, combined total length should be 25% of the bearing circumference. (5) Pin stakes should be located 0.030 ± 0.010 inch from the outer diameter of the bearing on the bearing housing. (6) Pin stakes should only be 0.010 to 0.032 inch deep to retain bearings when the bearing housing is staked on both sides. (7) The number of pin stakes around a bearing housing should be as indicated in Table 201. CAUTION: If the bearing should slip or move in the bearing housing, the bearing must be removed and recleaned. The bearing housing must be recleaned. Examine the bearing for any damage and reinstall in the bearing housing. Table 201 Recommended Stakes for Bearings Bearing OD

Number of Stakes

Up to 0.734 inch

4

0.735 inch to 0.984 inch

6

0.985 inch to 1.234 inches

8

(8) After the bearing has been secured in the housing by staking, lubricate with the proper lubrication and install in the airplane.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 201 Bearing Removal OUTER BEARING RACE HOUSING 0.010 TO 0.025 R

0.010 TO 0.032 INCH

SUPPORT

NOTE:

INNER BEARING RACE

DO NOT SUPPORT AGAINST INNER BEARING RACE DURING STAKING.

Figure 202 Bearing Staking

Page 204 May 1/12

20-05-00

STAKE

UC33B 111561AA.AI

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

STANDARD PRACTICES - AIRFRAME LOCKING DEVICES DESCRIPTION AND OPERATION

20-07-00 00

1. GENERAL Except where specific instructions may be required to satisfy certain applications, the following procedures are standard methods for installing the various locking devices used on bolts, screws and nuts:

A. Self-Locking Nuts Where self-locking nuts are used, the following procedure applies: (1) For each item, note the torque necessary to turn the nut on the bolt before seating the nut. (2) Add the above torque to the value detailed in the assembly instruction for the application. Use this new value as the total applied torque. NOTE: Before reinstallation of self-locking nuts, nuts should be checked for effectivity of the self-locking feature. Reject suspect nuts as necessary.

B. Lockwire and Cotter Pin Requirements When tightening a castellated nut, alignment of the slot must be obtained without exceeding the maximum torque. If this is not possible, replace the nut with another one. After tightening the nut to the recommended torque, the nut must not be loosened to permit insertion of lockwire or a cotter pin. If the slot in the nut or lockwire hole in the bolt or screw is not correctly aligned at the minimum torque value given, the nut, screw or bolt should be further tightened to the next alignment position, but the maximum torque value given must not be exceeded. Should alignment still be impossible without exceeding the maximum torque, back off the nut, screw or bolt one-half turn and retorque.

C. Taper Pins Taper pins are used where movement between two parts is not wanted and other fastener options are not desired.

D. Slotted, Steel Locknuts (Prevailing Torque Type) Effective locking of slotted, steel locknuts on bolts or studs requires full engagement of all locknut threads. The chamfered section of the locknut ID does not exert force on the bolt or stud; therefore, it is not necessary that the bolt or stud be flush with, or protrude from the outer face of the locknut.

E. Standard And Stepped Studs When the torque required to drive a stud to the correct protrusion does not reach the minimum value given, or exceeds the maximum value given, a new stud must be selected.

20-07-00

Page 1 Nov 1/09

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

F. Hose, Tubing and Threaded Couplings NOTE: Refer to Chapter 29 of the Model 1900 Airliner Series Component Maintenance Manual for procedures to install Cryofit or Permaswage fittings. If leakage occurs at a coupling, do not attempt to correct it by overtorquing. Disassemble the fitting and check for nicks, burrs and/or foreign matter. Use new parts to rectify.

Page 2 Nov 1/09

20-07-00

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

STANDARD PRACTICES - AIRFRAME LOCKING DEVICES MAINTENANCE PRACTICES

200200

1. PROCEDURES A. General Lockwire, lock washers, tab locks, tab washers, key washers, cup washers and cotter pins must never be reused. All lockwire and cotter pins must fit snugly into drilled holes in the bolts and studs for locking purposes. Bushings and plugs must be lockwired to boss or casing. Do not lockwire the bushing to the plug. Install cotter pins so that the head fits into the slot of the castellated nut and, unless otherwise specified, bend one end of the pin back over the stud or bolt and the other end flat against the flat on the nut.

B. Lockwire Use the same type and diameter of lockwire as that employed during the initial assembly. Except where otherwise specified, the wire used on the airplane power plant is heat- and corrosion-resistant steel wire of 0.025-inch diameter. (1) BASIC RULES (a) Lockwire must be tight after installation to prevent failure due to rubbing or vibration. (b) Lockwire must be installed in a manner that tends to tighten and keep a part locked in place, thus counteracting the natural tendency of the part to loosen. (c) Lockwire must never be overstressed. It will break under vibrations if twisted too tightly. The lockwire shall be pulled taut when being twisted, but shall have minimum tension, if any, when secured. (d) Lockwire ends must be bent toward the engine, or structure, to avoid sharp or projecting ends which might present a safety hazard or vibrate in the air stream. (e) Internal wiring must not cross over, or obstruct, a flow passage when an alternate method can be used. (2) LOCKWIRE HOLE ALIGNMENT (a) Check the units to be lockwired to make sure that they have been correctly torqued and that the wiring holes are properly positioned in relation to each other. When there are two or more units, it is desirable that the holes in the units be in the same relationship to each other. Never overtorque or loosen units to obtain proper alignment of the holes. It should be possible to align the wiring holes when the units are torqued within the specified limits. However, if it is impossible to obtain a proper alignment of the holes without either over- or undertorquing, select another unit which will permit proper alignment within the specified torque limits.

20-07-00

Page 201 Nov 1/09

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

(3) LOCKWIRE TWISTING (a) To prevent mutilation of the twisted section of the wire when using pliers, grasp the wire at the ends or at a point that will not be twisted. Lockwire must not be nicked, kinked or mutilated. Never twist the wire ends off with the pliers and, when cutting off ends, leave at least three complete turns after the loop, exercising extreme care to prevent the wire ends from falling into areas where they might create a hazard or cause damage. The strength of the lockwire holes is marginal; never twist the wire off with pliers. Cut the lockwire close to the hole, exercising extreme care. (4) LOCKWIRE ILLUSTRATIONS (a) Figure 201 illustrates a typical lockwiring procedure. Although there are numerous lockwiring operations performed on the airplane, practically all are derived from the basic examples shown (Ref. Figure 202).

C. Retaining Rings (Spirolox, etc.) Retaining rings must be installed with approved retaining ring pliers. Internal rings must not be compressed beyond the point where ends of the ring meet. External type rings must be expanded only enough to allow installation without becoming bent. After installation, ensure each retaining ring is completely seated in its groove, without looseness or distortion.

D. Taper Pins (AN386 Only) Refer to Figure 203 and Figure 204 for illustrations relating to torque tube hole location and taper pin installation. Refer to the narrative and Table 201 for the procedure used to correctly size and ream the torque tube hole for the taper pin. NOTE: In practice, to ensure the correct fit of taper pins to the corresponding holes, taper ream the holes to an undersize condition. The finish taper ream is done with a taper ream having the original taper (a taper reamer that has not been sharpened or ground). Sharpened or ground reamers have been known to create a problem in maintaining the proper ratio between the small and large hole. The ratio between the two holes is critical for correct seating of the taper pins. (1) Drill a hole (Ref. Figure 206) as detailed in Table 201 for the appropriate diameter taper pin. (2) Rough and finish the taper ream to the correct diameter and a smooth, consistent surface. The diameter is correct when the small end of the taper pin extends a maximum of 0.062 inch through the part when the pin is seated (Ref. Figure 204). Light tapping with a rawhide mallet may be necessary to seat the taper pin. NOTE: Gaging the depth from the surface of the large end of the hole is not recommended due to the thickness tolerances of the parts when assembled and taper pin diameter tolerances. (3) Deburr as required. (4) Install the taper pin, washer and nut (Ref. Figure 204).

Page 202 Nov 1/09

20-07-00

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 201 Lockwiring Procedures

20-07-00

Page 203 Nov 1/09

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 202 Lockwiring Examples

Page 204 Nov 1/09

20-07-00

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

MACHINIST'S V BLOCK (TYPICAL)

HOMEMADE V BLOCK 3/4" or 1" 4"

WELD

STEEL ANGLES

1. OBTAIN OR MAKE A V BLOCK

DRILL PRESS

3. PUNCH MARK HOLE LOCATION AUTOMATIC PUNCH 2. MOVE V BLOCK TO CENTER DRILL BIT IN V

V BLOCK

TUBE/ROD/CYLINDER, ETC.

4. LOAD TUBING WITHOUT MOVING V BLOCK JIG

5. DRILL THROUGH COMPLETELY

UC20B 050246AA.AI

Figure 203 Torque Tube Hole Location

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THE LARGE END OF THE TAPER PIN MUST NOT TIGHTEN UP BELOW THE SURFACE OF THE TUBE.

SMALL END OF PIN FLUSH TO .062 INCH BEYOND END OF TUBE. COLLAR REF.

NUT REF. UC20B 050247AA.AI

Figure 204 Taper Pin Installation

Table 201 Taper Pin Information Taper Pin Size

Recommended Maximum Pilot Hole Size

Taper Pin Size

Recommended Maximum Pilot Hole Size

AN386-1

0.201

AN386-7

0.600

AN386-2

0.251

AN386-8

0.750

AN386-3

0.3135

AN386-9

0.900

AN386-4

0.351

AN386-10

1.045

AN386-5

0.451

AN386-10A

1.147

AN386-6

0.501

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STANDARD PRACTICES - AIRFRAME AIRPLANE FINISH CARE MAINTENANCE PRACTICES

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1. PROCEDURES CAUTION: The urethane finish undergoes a curing process for a period of time after application. During the first month after paint application, some special care is required. Airplane owners should observe the following recommendations in order to preserve the durability and appearance of the airplane paint.

A. Cleaning Airplane Finishes CAUTION: Prior to washing, attach the pitot cover securely and plug or mask off all other openings. Be particularly careful to mask off all static air buttons before washing or waxing. Use special care to avoid washing away grease from any lubricated area. Prior to cleaning, cover such areas as wheels, brakes, etc., and relubricate after cleaning as necessary. Always be sure all maskings and coverings are removed before returning the airplane to service.

B. During the Curing Period (One Month) CAUTION: Prior to washing, attach the pitot cover securely and plug or mask off all other openings. Be particularly careful to mask off all static air buttons before washing or waxing. Use special care to avoid washing away grease from any lubricated area. Prior to cleaning, cover such areas as wheels, brakes, etc., and relubricate after cleaning as necessary. Always be sure all maskings and coverings are removed before returning the airplane to service. (1) Avoid prolonged flights in heavy rain or sleet. Avoid any operating conditions which might cause abrasion or premature finish deterioration. (2) Clean the airplane with mild detergents and water only. Use a clean soft rag, keeping it free from dirt and grime. Rinse with clear water thoroughly. (3) Use no waxes, polishes, rubbing compounds, or abrasive cleaners of any type. The use of such items can permanently damage the surface finish. (4) Stubborn oil or soot deposits on cowlings, wheel wells, etc. may be removed gently with automotive tar removers.

C. After the Curing Period (1) Continue to wash the airplane regularly. Use mild detergents and water only. Rinsing thoroughly with clear water prevents detergent residue buildup that can dull the paint appearance. (2) Normally, waxing is not necessary; however, if waxing is desired, select a high quality automotive or airplane waxing product. Never use rubbing compounds or abrasive cleaners of any type. Do not use a wax containing silicone because silicone polishes are difficult to remove from surfaces.

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D. Environmental Fallout (Acid Rain) After the specified curing period, avoid outside storage when conditions exist where moisture may collect on painted surfaces. Acids which remain in standing water can stain the paint topcoat and cause permanent damage to the finish. Flush off residual moisture with clean tap water and dry the surface. At this time, waxing the surface can provide protection from acid rain damage.

E. Placard Replacement Ascertain that all placards are in place and legible whenever the airplane has been repainted or touched up after repairs. Replace any placards that have been inadvertently defaced after such repainting or repairs.

2. EXTERIOR FINISHES (ALUMINUM SURFACES) A. Urethane Paints The need for an extremely hard finish for protection against sandblast during takeoff and landings led to the development of urethane coatings for airplanes. Urethane paint dries into a high gloss and retains color much better than standard finishes. It is unaffected by the chemicals in hydraulic fluids, deicer fluids and fuels and requires less care and maintenance than standard finishes. NOTE: Anytime an airplane is repainted or touched up, inspect all placards to assure that they are not covered with paint, are easily readable, and are securely attached. Replace any defaced placards.

B. Urethane Paint Repair Procedures NOTE: The time normally required for urethane paint to cure must be extended at temperatures below 70°F. The paint will not cure at temperatures below 60°F. Model 1900 Series Airliners are finished with urethane primer, and a top coat of urethane enamel. The following procedures include cleaning, paint stripping, repaint preparation, priming, applying a urethane topcoat, and an alternate method for small repairs not requiring paint stripping. Careful observance of these procedures shall result in a smooth, hard, glossy finish with firm adhesion for maximum life. NOTE: Precut stripe, numeral, and letter patterns are available through Modagraphics of Kansas, 1720 S. 151 W. Route 1, Wichita, Kansas 67052.

C. Paint Stripping and Cleaning Urethane Paint CAUTION: Never use aluminum foil to mask electrothermal windshields during painting, for most metal brighteners will combine with aluminum to form a hydrogen gas that eats away the stannous oxide used as an antistatic coating on electrothermal windshields. If metal brighteners are used, cover the windshield with paper or pasteboard masking material. Because of their resistance to chemicals and solvents, urethane paints and primers require a special paint stripper. If a urethane stripper is not available, a good enamel stripper may be used. Removing the finish with such a substitute will require several applications while working the stripper in with a stiff brush or wooden scraper. (1) Mask around the edge of the skin or skins containing the damaged area. Use a double thickness of heavy paper to prevent accidental splashes of paint stripper from penetrating the masking.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL WARNING: Urethane strippers usually contain acids that irritate or burn the skin. Wear rubber gloves and eye protection when using the stripper. (2) Apply urethane stripper as indicated by the manufacturer's directions. Try to stay approximately 1/ 8 inch away from the masking tape. This will necessitate a little more cleanup upon finishing, but will prevent damage to the finish on the next skin. The stripper will not attack aluminum during the stripping process and can be neutralized afterwards by rinsing the affected area with water. (3) Rinse the area with water and dry. (4) Wash the stripped area carefully with a solvent such as acetone, methyl propyl ketone (MPK), or lacquer thinner. This will prevent tiny particles of loose paint from adhering to the stripped area. (5) Using a nylon scratch pad or aluminum wool dipped in clean water, clean the surface with a cleanser such as Bon Ami, Ajax, Comet cleaner, etc. A good scouring will leave the surface completely clean. (6) Thoroughly rinse with clean water and carefully dry the affected area. If the stripped area includes several joints or skin laps, let the airplane sit until all moisture has dried. This may be accelerated by blowing the skin laps and seams with compressed air. Wet masking should be replaced.

D. Urethane Primer (1) Mix the urethane primer (3, Table 1, Chapter 91-00-00) and catalyst (4, Table 1, Chapter 91-00-00) in accordance with the manufacturer's instructions when preparing the primer. NOTE: For the best results, these directions must be followed carefully, for some manufacturers require that the primer be allowed to set for 1/2 hour after the catalyst and base have been mixed while others recommend immediate use after mixing. (2) Apply a coat of urethane primer with a spray gun using 25 to 40 psi of air pressure. A dappled appearance only indicates that the coat is thin. (3) If the initial primer coat is allowed to cure for more than 24 hours before the topcoat is applied, sand the primer slightly to roughen the surface and assure adhesion. Wipe off the sanding dust with a cloth dampened with a solvent (such as lacquer thinner), then apply the topcoat. NOTE: The minimum drying time for urethane primer is approximately 2 hours under conditions of low humidity at temperatures of 85 to 90°F. When the primer coat cannot be scratched by fingernail or when sandpapering does not cause the primer to ball up, the urethane topcoat may be safely applied.

E. Urethane Topcoats (1) Mix the paint and catalyst as described by the manufacturer. (2) Apply the topcoat with a spray gun at 35 to 40 psi of air pressure. Two coats are normally required to fully conceal the primer and build up the topcoat film necessary for adequate service life and beauty. The urethane finish will normally cure to approximately 85% of its full hardness in 24 hours at temperatures of 80°F or higher. NOTE: When interior or exterior paint is required, refer to the airplane log book for the part number of the paint used on the airplane as it was delivered new.

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F. Urethane Touch-Up Repair (1) Mask around the skin containing the damaged area. (2) Remove all loose edges of paint by using a high tack adhesive tape around the edge of the damaged area. (3) Using a coarse sandpaper, fair the edge of the damaged area with the metal. (4) When the edge of the paint begins to fair into a smooth joint, use a fine grade of sandpaper to eliminate the scratches left by the coarse paper. Take care to avoid removing any more metal than is absolutely necessary. (5) Wash the sanded area with a solvent, such as lacquer thinner or methyl propyl ketone (MPK). Change the wash cloths used for this purpose often so that all the sanding dirt will be picked up. (6) After the area to be touched up has been cleaned with solvent until all traces of discoloration are gone, apply a thin coat of pretreatment primer to the damaged area. (7) After the urethane primer has cured for 24 hours, sand the area under repair with medium fine sandpaper. Sand the edge of the repair area until the indentation where the metal and old paint meet is gone. If necessary, apply additional urethane primer until the juncture of old paint with metal is no longer visible. (8) Spray on two topcoats.

G. Special Procedures The following special procedures should be followed when repainting the following areas: (1) Battery Box and Lid The interior surfaces of the battery box and battery lid are to be finished as follows: (a) Apply a minimum of two coats of epoxy-polyamide primer (5, Table 1, Chapter 91-00-00). (b) Apply a minimum of three coats of exterior matterhorn white (6, Table 1, Chapter 91-00-00). (2) Rubber Seals Apply one coat of a thoroughly dissolved solution of one part Oakite No. 6 and two parts water to all rubber surfaces that are to come in contact with either metal or other rubber surfaces. (3) Laminated Fiberglass Surfaces Apply sanding surfacer (7, Table 1, Chapter 91-00-00) until the surface to be painted is smooth. (4) NOSE RADOME CAUTION: Radar operation can be adversely affected by abrupt variations in the thickness of the material in the nose radome, which could be caused by excessive and/or uneven buildup of repair material during repair of the damage. It is recommended that a damaged nose radome be replaced rather than repaired. However, if repairs are made, a transmissivity check should be accomplished by a certified repair station.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL NOTE: Information to repair the radome lightning diverter strips is provided in the Model 1900 Airliner Series Structure Repair Manual, Chapter 53-90-15. New radomes purchased from Hawker Beechcraft Corporation are finished complete with primer and anti-static coating, ready for final painting; however, the new radome must be fit, trimmed, drilled and countersunk to the individual airplane. To assure electrical bonding to the fuselage, a small metal tab or strap should be attached to the inner aft edge of radome with two rivets after fitting. After installation of the tab, the rivet heads on the exterior surface of the radome should be painted with one coat of 528-104 or 528-306 flat black anti-static paint (113, Table 1, Chapter 91-00-00). Fabricate and attach the copper strap as follows: (a) Trim, fit, drill and countersink the new radome to fit the airplane. (b) Touch up the anti-static paint in all countersunk screw holes with one coat of 528-104 or 528-306 flat black anti-static paint (113, Table 1, Chapter 91-00-00). Ensure that the paint covers the entire inside of the countersink and makes contact with the anti-static paint or primer on the exterior of the radome. (c) Fabricate a strap out of 1/32-inch to 1/16-inch thick copper strap (Ref. Figure 201). The strap should be 3-inches in length, 3/4-inches in width. Drill one hole of the proper diameter so that a radome mounting screw will fit through it with little play. (d) Attach the strap to the radome with two Aluminum CherryMax rivets (CR3213-4) at the end of the strap, opposite of the mounting screw hole. NOTE: The grip length of the rivets should be based on the total buildup on the radome. Maintain proper Edge Distance and rivet spacing. (e) Paint the top of the rivet heads that are exposed, with the anti-static or primer, on the exterior of the radome. (f) Prior to priming and painting the radome, mask the copper strap off so that it remains clear of paint and provides a good electrical bond from the radome to the fuselage. Repainting the radome will affect the transmissivity requirements of the radome; therefore, a maximum of three thin coats of polyurethane paint only should not be exceeded. Use of any paint other than polyurethane is not recommended. When preparing the radome for repainting, use a fine grit sandpaper, but do not sand into the black anti-static coating beneath the finish paint coat. NOTE: Multiple coats of anti-static coating will change resistance; therefore, anti-static coating should NOT be repaired with additional coats of anti-static coating or flat black paint.

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Figure 201 Nose Radome Strap Installation (5) Radome Protective Boot If the airplane is equipped with a radome protective boot, the following removal, installation and maintenance procedures must be followed precisely to ensure serviceability of the boot installation. (6) Radome Protective Boot Removal (a) Using a razor blade, carefully score the surface of the protective boot and dissect the protective boot into triangular pie-shaped sections. Each section should be approximately four to six inches wide at the base. Do not cut into the radome. (b) Pull the pie-shaped sections from the radome beginning at the top and pulling down. Pull the next section at 180 degrees from the previously pulled section until all sections are removed. CAUTION: Observe the manufacturer's safety recommendations when cleaning the radome. (c) Remove any adhesive residue from the radome using a clean cloth dampened with a 75% solvent (14, Table 1, Chapter 91-00-00) and 25% toluol (18, Table 1, Chapter 91-00-00) solution. Wipe dry using a dry clean cloth. Do not allow the cleaning solution to set on the radome surface, possibly causing damage to the radome. (7) Radome Protective Boot Maintenance Maintenance of the unpainted protective boot includes the application of a paste wax when needed to prevent staining. The protective boot should be replaced at the first indication of damage. Page 206 Nov 1/09

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (8) Radome Protective Boot Installation Install the SJ-8665 3M radome protective boot as follows: (a) Thoroughly wash the refurbished and primed radome with isopropyl alcohol (30, Table 1, Chapter 91-00-00) and wipe dry. (b) Using a marking pen, place a + on the top center of the radome for boot orientation. Place a zero at one end of the + symbol. (c) With the protective liner still in the protective boot, place the boot on the radome. (d) Rotate the protective boot until the best fit is obtained on the radome. (e) Place a + and zero on the protective boot at the same location as the + and zero on the radome. (f) Remove the protective boot from the radome, turn the protective boot inside out and place over the radome. At this time, disregard the orientation marks. (g) Carefully remove the transparent protective liner from the protective boot. (h) Saturate the exposed adhesive surface of the protective boot with a wetting solution of 25% isopropyl alcohol (30, Table 1, Chapter 91-00-00), 75% water and 1 teaspoon of liquid detergent (Ivory or Joy) per gallon of wetting solution to prevent adhesive-to-adhesive bonding. NOTE: The entire adhesive surface of the protective boot must be saturated with the wetting solution. (i) Remove the protective boot from the radome. (j) Saturate the entire surface of the protective boot contact area on the radome with the wetting solution. (k) Place the protective boot, adhesive side in, on the radome with the orientation marks aligned. (l) Using a methyl propyl ketone (MPK) resistant plastic squeegee with rounded edges, squeegee the wetting solution from under the protective boot. Start at the top center of the protective boot and work the blisters down to the end and out from under the protective boot. NOTE: Care must be taken to avoid leaving blisters under the protective boot. (m) If small blisters still remain under the protective boot, pierce them with a safety pin and relieve the entrapped air or wetting solution. Work with the squeegee if necessary. (n) Wrap the edge of the protective boot outer circumference at the desired location with 1/ 2-inch wide masking tape. (o) Using an industrial razor blade, trim the protective boot at the edge of the 1/2-inch wide masking tape. NOTE: Use extreme caution during the trimming of the protective boot. The razor blade must not cut into the radome.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (p) Place the radome in a bake oven at 125° to 150°F for 20 minutes. NOTE: If a bake oven is not available, a heat lamp at the same temperature may be used. If no heat lamp or oven is available, the radome must set for 24 hours at room temperature for the adhesive to cure. (q) Paint the radome and protective boot as follows: 1 Wipe the surface of the radome and protective boot with a clean rag dampened with methyl propyl ketone (14, Table 1, Chapter 91-00-00). Observe the manufacturer's safety recommendations at all times. If methyl propyl ketone is not desirable, use a #3774 Scotch brite pad and lightly scuff the surface of the protective boot and radome surface. Using a clean cloth dampened with ethanol, wipe the protective boot and radome clean. 2 Do not prime the protective boot before painting. 3 Allow the radome and protective boot to dry for approximately 30 minutes. 4 Apply the topcoat paint directly to the radome and protective boot as instructed by the paint manufacturer.

H. Paint Free Areas The following areas shall be kept free from paint. (1) Engine controls. (2) Flight control cables and chains. (3) Control pedals. (4) Exhaust heated air inlet lip (stainless steel). (5) Firewalls and wrought aluminum surfaces forward of the firewall, with the following exceptions: (a) Aluminum parts attached directly to the firewall shall be primed and painted in detail. (6) Aluminum flexible conduit. (7) All tubing (except unplated steel which shall receive two coats of primer) on the exterior and interiors where color scheme must be maintained. (8) Interior of all fluid lines, including oxygen lines and instrument lines. (9) Chromium-plated portions of the landing gear piston tubes. (10) Rubber and rubberlike surfaces. (11) Electrical wiring, unless otherwise noted as a specific requirement. (12) Glide path antenna. (13) Pitot static buttons. (14) Air conditioner condenser and evaporator coils. Page 208 Nov 1/09

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (15) Air conditioner condenser fan. (16) Air conditioner compressor. (17) Avionics honeycomb shelves. (18) Cargo door and airstair door areas as follows: (a) Cargo door only: 1 Surface of slide pins (slide pin carrier rod is to be painted). 2 Gas springs and buffed aluminum bracket. 3 Roller guides on the fuselage door opening frame. (b) Airstair door only: 1 Buffed surface of the airstair door snubber and the handrail post. 2 Airstair door steps. (c) Both cargo and airstair doors: 1 Roll cams. 2 Insert threads and fastener threads. 3 Door seals. 4 Buffed door handle. 5 All cables. 6 All cable drums and cable drum assemblies. 7 All internal and external splines. 8 All shaft bearing surfaces. 9 All internal bearing journal surfaces. 10 All solid film lubricated surfaces. 11 Chromed latch posts. 12 Upholstery panels. NOTE: The door seal retainer may be painted or left paint free, but must be kept free of paint overspray.

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3. MAGNESIUM SURFACES A. Paint Removal From Magnesium Surfaces (1) Mask around the edge of the damaged area with a double thickness of heavy paper to prevent accidental splashes of paint stripper from penetrating the masking. WARNING: Stripping should be accomplished in a well ventilated area since prolonged exposure to high concentrates of stripper vapor may irritate the eyes and lungs. (2) Apply paint stripper (8, Table 1, Chapter 91-00-00) to the skin under repair with a brush or nonatomizing gun. CAUTION: Never use a wire brush for it will damage the magnesium surface. (3) Allow the paint stripper to work for 20 to 30 minutes, then work the remaining paint loose with a bristle brush. (4) Remove the masking paper and wash the affected area thoroughly with water under high pressure. Remove all remnants of paint with lacquer thinner. (5) Sand the repaired area lightly, then apply Dow Treatment Number 19 solution to prevent corrosion. Dow No. 19 solution may be mixed as follows: (a) Place 3/4 gallon of (70°F to 90°F) water in a one gallon stainless steel, aluminum or vinyl polyethylene container. (b) Add 1-1/3 oz. of chromic acid (CrO3), then 1 oz. of calcium sulfate (CaSO4) (obtain chemicals locally). (c) Add 1/4 gallon (70°F to 90°F) of water and stir vigorously for at least 15 minutes. CAUTION: Do not exceed 3 minutes. (d) Brush the solution on the bare area, keeping the area wet for 1 to 3 minutes until a brown film appears. (e) Rinse the treated area in COLD running water and dry in an oven or by exposure to hot air blast. NOTE: If cold running water is not available, rinsing may be eliminated and the area dried as stated above. (f) Brush 2 coats (a minimum of 30 minutes apart) of epoxy primer (5, Table 1, Chapter 91-00-00) on and around the reworked area. Assure adequate penetration of primer into the treated area.

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B. Painting Magnesium Surfaces (1) Prepare the surface to be repainted as indicated under PAINT REMOVAL FROM MAGNESIUM SURFACES. Clean the affected area thoroughly with lacquer thinner or an equivalent solvent. NOTE: Do not apply wash primer to magnesium surfaces. Allow a minimum of four hours drying time between application of the primer and topcoat. (2) Prime the affected area and apply the topcoat as indicated under the Urethane Paint Procedure in this chapter.

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STANDARD PRACTICES - AIRFRAME CORROSION DESCRIPTION AND OPERATION

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1. GENERAL The information contained in this section has been removed, refer to the Model 1900 Airliner Series Corrosion Control Manual, P/N 114-590021-197.

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STANDARD PRACTICES - AIRFRAME CORROSION MAINTENANCE PRACTICES

200200

1. PROCEDURES The information contained in this section has been removed, refer to the Model 1900 Airliner Series Corrosion Control Manual, P/N 114-590021-197. The CORROSION PREVENTION FOR CHAFED FUEL LINES procedure has been moved to Chapter 28-21-00.

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STANDARD PRACTICES - AIRFRAME AIRFRAME PENETRATION INSPECTION MAINTENANCE PRACTICES

20-10-00 200200

1. PROCEDURES The information contained in this section has been removed, refer to the Model 1900/1900C Airliner Structural Inspection Manual, P/N 98-30937, Chapter 20-00-00.

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STANDARD PRACTICES - AIRFRAME CONTROL OF LIFE-LIMITED PARTS MAINTENANCE PRACTICES

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1. PROCEDURES A. Determination Of Serviceability When Part Life Is Unknown 14 CFR Part 43 requires that all persons who remove life-limited parts from airplanes safely control these parts. Acceptable methods to control these parts are defined by 14 CFR Part 43. In certain circumstances the life status of a part may not be known. The following guidelines may be used to assess the life status of an aircraft part. Note that this is not an exhaustive list. WARNING: If the life status of a part cannot be determined, the part shall be considered to be beyond its safe life and must be disposed of in accordance with 14 CFR Part 43. (1) If the part is original to the airplane: (a) For determining compliance with calendar time life limitations, the beginning of the part life shall be the date of issuance of the original aircraft airworthiness certificate. (b) For determining compliance with usage related life limitations, the total aircraft hours, landings, or cycles shall be considered as appropriate. (2) If the part is a replacement: (a) For determining compliance with calendar-time life limitation, consider the calender time since the logbook verified installation date plus any prior usage of the replacement part. (b) For determining compliance with usage related life limitations, consider the airplane hours, landing, or cycles as appropriate, accrued since the logbook verified part installation plus any prior usage of the replacement part. Once the life of the part is determined, the part should be controlled by one of the acceptable methods outlined in 14 CFR Part 43.

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CHAPTER 21 - ENVIRONMENTAL SYSTEMS TABLE OF CONTENTS SUBJECT

PAGE

GENERAL INFORMATION 21-00-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Bleed Air Control . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Distribution . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Pressurization Control . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Heating . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Cooling . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Air Cycle System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Vapor Cycle System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Temperature Control . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

1 1 1 2 2 2 2 3 7 7

BLEED AIR CONTROL 21-10-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General (UA-1 and After, UB-1 and After and UC-1 thru UC-45, Except UC-39) . . . . . . . . . . . . . . . . . . . . . . 1 Troubleshooting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 101 Procedures (UA-1 and After, UB-1 and After and UC-1 thru UC-45, Except UC-39) . . . . . . . . . . . . . . . . . 101 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Procedures (UA-1 and After, UB-1 and After and UC-1 thru UC-45, Except UC-39) . . . . . . . . . . . . . . . . . 201 Precooler Functional Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Temperature Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 203 Pressure Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 205 Pressure Operational Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 206 Temperature Operational Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 206

PRECOOLER - THROUGH AND BYPASS VALVES 21-10-01 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Procedures (UA-1 and After, UB-1 and After and UC-1 thru UC-45, Except UC-39) . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 201 201

PRESSURE REGULATOR/SHUTOFF VALVE 21-10-03 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Janitrol Pressure Regulator/Shutoff Valve Procedures (UA-1 and After, UB-1 and After and UC-1 thru UC-45, Except UC-39) (Without Kit No. 114-5037 Installed) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Filter Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Filter Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Dukes Pressure Regulator/Shutoff Valve Procedures (UA-1 and After, UB-1 and After and UC-1 thru UC-45, Except UC-39) (With Kit No. 114-5037 Installed) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Adjustment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

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201 201 201 202 203 203 204 204 204 204

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ACM OVERPRESSURE SWITCH 21-10-06 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Procedures (UA-1 and After, UB-1 and After and UC-1 thru UC-45, Except UC-39) . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 201 201

ACM OVERTEMPERATURE SENSOR 21-10-07 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Procedures (UA-1 and After, UB-1 and After and UC-1 thru UC-45, Except UC-39) . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 201 201

BLEED AIR CONTROL 21-11-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General - (UC-39, UC-46 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Bleed Air Valves Switch . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Bleed Air Control Module . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Precooler Valves . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Overpressure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2 Over temperature . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2 Troubleshooting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 101 General - (UC-39, UC-46 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 101 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Procedures - (UC-39, UC-46 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Precooler Functional Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Temperature Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 203 Pressure Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 210 Pressure Operational Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 210 Temperature Operational Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 210

PRECOOLER - THROUGH AND BYPASS VALVES 21-11-01 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Procedures - (UC-39, UC-46 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 201 201

TEMPERATURE CONTROLLER SENSE LINE FILTER 21-11-02 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Procedures - (UC-39, UC-46 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

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PRESSURE REGULATOR/SHUTOFF VALVE 21-11-03 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Janitrol Pressure Regulator/Shutoff Valve Procedures (UC-39, UC-46 and After) (without Kit No. 114-5037 Installed) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Filter Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Filter Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Dukes Pressure Regulator/Shutoff Valve Procedures (UC-39, UC-46 and After) (with Kit No. 114-5037 Installed) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Adjustment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 201 201 202 202 203 203 203 203

ACM OVERPRESSURE SWITCH 21-11-06 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Procedures - (UC-39, UC-46 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 201 201

ACM OVERTEMPERATURE SENSOR 21-11-07 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Procedures - (UC-39, UC-46 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 201 201

DISTRIBUTION 21-20-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Vapor Cycle Air Distribution . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Air Cycle System Distribution . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2 Defrost and Flight Compartment Heat . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2 Troubleshooting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 101 Distribution . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 101 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Flapper Valve Rigging, Conditioned Bleed Air . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Flapper Valve Rigging, Nose Ram Air . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Butterfly Valve Rigging, Heat and Defrost . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201

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VENT BLOWER 21-20-01 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 201 201

AIR OUTLET 21-20-02 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Air Outlet Actuator . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 201 201 201 201 201

PRESSURIZATION CONTROL 21-30-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Troubleshooting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 101 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 101 Airborne Functional Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 101 Airborne System Checks . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 101 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Pressurization Check Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Equipment Required . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Prestart Procedure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Preflight Procedure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 During Climb . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202 Cabin Leak Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202 Single Source Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202 Maximum Differential Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 203 Controller Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 203 Cabin Leak Rate - Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 204 Cabin Leak Rate Flightline Ground Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 204 Maximum Differential Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 206 Cabin Leak Rate Ground Test Equipment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 207 Cabin Leak Rate Hangar Ground Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 211

OUTFLOW VALVE 21-30-01 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Cleaning . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

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CABIN PRESSURE CONTROLLER 21-30-02 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Filter Cleaning . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 201 201 201

PNEUMATIC RELAY 21-30-03 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Filter Cleaning . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 201 201 201

VOLUME TANK 21-30-04 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 201 201

CABIN ALTITUDE WARNING PRESSURE SWITCH 21-30-07 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Functional Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Method One . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Method Two . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 201 201 201 201 202

HEATING 21-40-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Troubleshooting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 101 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 101 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201

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COOLING 21-50-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Troubleshooting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 101 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 101 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Vapor Cycle System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Operational Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201

AIR CYCLE SYSTEM 21-51-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Troubleshooting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 101 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 101 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Air Cycle Machine . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Servicing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Water Collector Drain Tube . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Weld Repair . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Oil . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Replacement (Oil Sump With Fill Plug) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Replacement (Oil Sump Without Fill Plug) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202

REFRIGERATION PACKAGE 21-51-01 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 201 201

FOG NOZZLE AND IN-LINE FILTER 21-51-02 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Removal and Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201

RECIRCULATING EJECTOR 21-51-03 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Heater Blanket Operational Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

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VAPOR CYCLE SYSTEM 21-52-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Hot Gas Bypass Valves . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2 Pressure Cutout Switches . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2 Vent Blowers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2 Pressure Relief Valve . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2 Troubleshooting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 101 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 101 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Precautionary Service Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Vapor Cycle System Maintenance Notes . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Refrigerant Leak Detection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202 Detergent Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 203 Electronic Detector Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 203 Red Leak Detector Dye Additive . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 203 Yellow/Green Leak Detector Dye Additive (Preferred) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 203 Vapor Cycle System Component Repair/Replacement . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 204 Compressor Oil Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 204 Depressurizing the Vapor Cycle System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 205 Evacuating the Vapor Cycle System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 205 Cleaning the Vapor Cycle System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 206 Charging the Vapor Cycle System (Airplanes without Kit 129-5020) . . . . . . . . . . . . . . . . . . . . . . . . . . 208 Charging the Vapor Cycle System (Airplanes with Kit 129-5020) . . . . . . . . . . . . . . . . . . . . . . . . . . . . 208 Vapor Cycle System Operational Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 209

EVAPORATOR 21-52-01 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Filter Replacement . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Forward Evaporator . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Aft Evaporator . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 201 201 201 201 202 202 202

COMPRESSOR (UA-1 AND AFTER; UB-1 AND AFTER; UC-1 THRU UC-100 NOT MODIFIED BY SERVICE BULLETIN NO. 2345) 21-52-02 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Compressor . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Compressor Mount . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

21-CONTENTS

201 201 201 202 203 203

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CHAPTER 21 - ENVIRONMENTAL SYSTEMS TABLE OF CONTENTS (CONTINUED) SUBJECT PAGE Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 204 Compressor Mount Support . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 204 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 204 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 205 Compressor Belt . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 205 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 205 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 206 Tension . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 206

CONDENSER AND BLOWER 21-52-03 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Condenser Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Condenser Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Condenser Blower Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Condenser Blower Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 201 201 201 201

RECEIVER - DRYER 21-52-04 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 201 201

COMPRESSOR (UA-1 AND AFTER; UB-1 AND AFTER; UC-1 AND AFTER MODIFIED BY SERVICE BULLETIN NO. 2345) 21-52-05 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Compressor . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Compressor Mount . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Compressor Mount Support . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Compressor Belt . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Alignment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tension . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 201 202 203 203 203 204 204 205 205 205 206 206 207

TEMPERATURE CONTROL 21-60-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Troubleshooting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 101 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 101 Page 8 Nov 1/10

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CHAPTER 21 - ENVIRONMENTAL SYSTEMS TABLE OF CONTENTS (CONTINUED) SUBJECT

PAGE

AIR DUCT TEMPERATURE SENSOR 21-60-01 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Resistance Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 201 202 202

CABIN TEMPERATURE CONTROLLER 21-60-02 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Functional Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Test Equipment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Test Procedure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Indicator Calibration . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

21-CONTENTS

201 201 201 201 201 201 201 202

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List of Effective Pages CH-SE-SU

PAGE

DATE

CH-SE-SU

PAGE

DATE

21-LOEP

1

Nov 1/13

21-50-00

21-CONTENTS

1 thru 8

Nov 1/10

1 101 201

Nov 1/09 Nov 1/09 Nov 1/09

21-00-00

1 thru 7

Nov 1/09

21-51-00

21-10-00

1 101 201 thru 206

May 1/11 Nov 1/09 Nov 1/09

1 and 2 101 and 102 201 and 202

Nov 1/09 Nov 1/09 Nov 1/09

21-51-01

201 and 202

Nov 1/09

21-10-01

201 and 202

Nov 1/09

21-51-02

201

Nov 1/09

21-10-03

201 thru 206

Nov 1/09

21-51-03

201 thru 203

Nov 1/09

21-10-06

201 and 202

Nov 1/09

21-52-00

21-10-07

201 and 202

Nov 1/09

21-11-00

1 and 2 101 201 thru 210

Nov 1/09 Nov 1/09 Nov 1/09

1 thru 4 101 thru 115 201 thru 211

Nov 1/09 Nov 1/09 Nov 1/13

21-52-01

201 thru 204

Nov 1/09

21-52-02

201 thru 211

Nov 1/10

21-11-01

201 and 202

Nov 1/09

21-52-03

201 and 202

Nov 1/09

21-11-02

201 and 202

Nov 1/09

21-52-04

201 and 202

Nov 1/09

21-11-03

201 thru 205

Nov 1/09

21-52-05

201 thru 213

Nov 1/10

21-11-06

201

Nov 1/09

21-60-00

21-11-07

201

Nov 1/09

1 and 2 101

Nov 1/09 Nov 1/09

21-20-00

1 thru 5 101 201 thru 203

Nov 1/09 Nov 1/09 Nov 1/09

21-60-01

201 thru 203

Nov 1/09

21-60-02

201 thru 204

Nov 1/09

21-20-01

201

Nov 1/09

21-20-02

201

Nov 1/09

21-30-00

1 thru 4 101 thru 105 201 thru 216

Nov 1/09 Nov 1/09 Nov 1/09

21-30-01

201

Nov 1/09

21-30-02

201 and 202

Nov 1/09

21-30-03

201

Nov 1/09

21-30-04

201 and 202

Nov 1/09

21-30-07

201 and 202

Nov 1/09

21-40-00

1 101 thru 103 201 thru 203

Nov 1/09 Nov 1/09 Nov 1/09

21-LOEP

Page 1 Nov 1/13

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

ENVIRONMENTAL SYSTEMS GENERAL INFORMATION DESCRIPTION AND OPERATION

21-00-00 00

1. GENERAL The Model 1900 Airliner Series environmental system makes use of engine bleed air for cabin pressurization, cabin heating and providing the motive force for operation of the air cycle machine (ACM), the primary source of cabin cooling (Ref. Figure 1). A vapor cycle system, driven from the RH engine, augments air cycle machine cooling during times of greater heat loads when additional cooling may be required. A system of valves, regulators and temperature and pressure sensors controls all physical aspects of the bleed air flowing into the cabin. Two outflow valves, operated by the pressurization controller and mounted on the aft pressure bulkhead, provide a controlled exit for the pressurization air in order to maintain a preselected pressure differential between the cabin environment and the outside air. Conditioned bleed air is distributed and recirculated by two distinctly different ducting systems. The outlets in the lower cabin sidewalls deliver conditioned bleed air and recirculated cabin air to the cabin while eyeball outlets in the cabin mid sidewall provide for cabin air recirculation and additional cooling by the vapor cycle system. Temperature regulation is provided for by the cabin temperature controller mounted in the center cabin overhead upholstery. Various modes of temperature control are provided for by the controlling circuitry; the system can be controlled manually by the flight crew, automatically by the cabin temperature controller, or the evaporator blowers can be operated independent of temperature control.

A. Bleed Air Control Bleed air from the compressor stages of the engines provides for pressurization, cabin heating and the motive force for driving the air cycle machine. A heat exchanger and two valves for each engine are used to precool the bleed air to 450°F before the bleed air is ducted into the air cycle machine or bypassed into the cabin. A bleed air shutoff valve/ regulator for each engine, downstream of the precooler valves, provides master control of the bleed air flow and pressure. Protective mechanisms are provided to terminate bleed air flow in the event of failures which would allow the bleed air temperature or pressure to increase without control. Time delay circuits on the relay control board will prevent termination of the bleed air flow in cases of instantaneous temperature or pressure changes of short duration. Should specific temperature and pressure limits be exceeded, the precooler bypass valve and the pressure regulator/shutoff valve will close on the affected side. An expanded discussion of the environmental bleed air system and controls will be found in 21-10-00 for all airplanes with single temperature sensors and pressure switches on each side. Serials UC-39, UC-46 and after and any airplanes that have installed modification Kits 114-5018 and 114-5016 will have an additional temperature sensor and pressure switch near the ACM and filters in the sense lines to the pressure regulator/shutoff valve and temperature controller. Refer to 21-11-00 for a detailed description and operation of this newer version.

21-00-00

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B. Distribution Bleed air from the engines is precooled and delivered to the cabin through outlets in the lower cabin sidewalls for pressurization, heating and cooling from the air cycle machine. The vapor cycle system recirculates and further cools the cabin air as required and distributes the air through adjustable eyeball outlets. The vapor cycle system makes use of the two evaporator blowers in order to recirculate the cooled cabin air. Evaporator coils are mounted on the inlet side of the evaporator blowers to facilitate the exchange of heat between the cabin air and the cooling media when the vapor cycle system is operating. Refer to 21-20-00 for a more detailed explanation of the operation of this system.

C. Pressurization Control Cabin pressurization is maintained at a selected level by the pressurization controller mounted in the pedestal. The controller has settings for cabin altitude and rate-of-change. The sole function of the pressurization controller is to control the outflow valves which open or close proportionally according to the degree of pneumatic vacuum being provided by the controller. The controller will automatically maintain cabin pressure at some level below a maximum operating pressure differential that is coincident to the pressure, relative to ambient, of the altitude setting on the controller. Should a fault in the controller cause the controller to fail to maintain cabin pressure differential below the maximum, the outflow valves are calibrated to open and dump the excess cabin pressure. Pneumatic vacuum air is supplied to the pressurization controller through a normally open solenoid type valve (preset solenoid valve) and to the outflow valves through a normally closed solenoid type valve (dump solenoid valve). These two valves are energized through the manual dump switch on the pedestal. When the switch is in the dump mode, both solenoid valves are energized and the pneumatic vacuum air is shunted from the pressurization controller to the outflow valves, opening them and dumping the cabin pressure to the outside. 21-30-00 contains complete information for troubleshooting and maintenance of this system.

D. Heating Bleed air from the engines enters the cabin distribution ducts for heating through two electrically operated rotary valves (the ACM bypass valve and the ejector bypass valve). The two valves operate in sequence rather than simultaneously. The ACM bypass valve opens first when heating is required and modulates the output of the air cycle machine. As the cabin temperature controller continues to ask for heating, the ACM bypass valve opens fully and shunts operating current to the ejector bypass valve, by way of the valve travel limit switches, opening it. When the cabin temperature controller begins asking for cooling, the ejector bypass valve begins closing. When the ejector bypass valve is fully closed, operating current is shunted back to the ACM bypass valve closing it. The valves also serve the purpose of switching the vapor cycle system compressor in or out. For more information on heating, refer to 21-40-00.

E. Cooling All cabin cooling is provided by the air cycle system and, when required, by the vapor cycle system. Vapor cycle system cooling is invoked, as discussed previously, by the switching activity of the ACM and ejector bypass valves. General information and preliminary troubleshooting of the cooling system is in 21-50-00.

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F. Air Cycle System The air cycle machine utilizes engine bleed air to drive a compressor which compresses the bleed air to a point where the excess heat of compression can be removed through the use of heat exchangers; then when the pressure is released, a specific heat deficit exists, relative to source ambient, and cooling is affected. The sequence in the air cycle machine is as follows: (1) Bleed air enters the air cycle machine through the first stage heat exchanger where excess heat is removed. (2) Bleed air travels through the air cycle machine compressor where it is compressed to a higher pressure and temperature. (3) Bleed air passes through the second stage heat exchanger where the excess heat of compression is removed. (4) Bleed air passes through the air cycle machine turbine providing the motive force necessary to drive the compressor and cooling air fan and the pressure is released and the bleed air allowed to expand producing a specific heat deficit in the bleed air relative to cabin ambient. For detailed maintenance information concerning this system, refer to 21-51-00.

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LO HI

LH

RH

UB21B 990801AA.AI

Figure 1 Environmental System Schematic

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G. Vapor Cycle System When heat loads are such that the air cycle system is producing maximum cooling, a signal is transmitted to the temperature controlling circuitry which causes the refrigerant compressor clutch to be engaged, thereby turning the compressor and initiating the cooling cycle of the vapor cycle system. A condensing coil and blower assembly, located in the right forward inboard wing, removes excess heat from the high temperature, high pressure gaseous refrigerant being discharged from the compressor, allowing the refrigerant to condense to the liquid state. The high pressure, low temperature liquid refrigerant passes through a metering device (thermostatic expansion valve) into the evaporator where the pressure is relieved and the refrigerant allowed to evaporate into the gaseous state producing a heat deficit in the gaseous refrigerant. Cabin air is circulated over the evaporator coil where heat is transferred from the cabin air to the gaseous refrigerant. The low pressure, low temperature refrigerant then returns to the compressor and the entire cycle is repeated. Once the vapor cycle system is activated, it will remain in operation until such time as the ACM bypass valve has opened fully, at which time a signal is transmitted to the heat side of the heat/cool relay and vapor cycle cooling is terminated. Certain measures have been taken to protect the vapor cycle system from damage during normal operation. A 40°F outside air temperature switch keeps the system from being activated while in low ambient temperature conditions; this switch closes at 50 ± 5°F and opens at 30 ± 5°F. Overpressure and underpressure switches deactivate the system in the event operating pressures exceed the maximum or minimum safe operating limits. 21-52-00 contains information to aid in troubleshooting and maintaining this system in proper operating condition.

H. Temperature Control Temperature control in the Model 1900 Airliner Series is accomplished by the cabin temperature controller which makes use of temperature sensitive resistance devices to register cabin temperatures. The totally solid state circuitry in the cabin temperature controller issues commands to the air cycle machine bypass and ejector bypass valves to control the amount of cooling provided by the air cycle machine. When the air cycle machine is providing maximum cooling, the cool command from the cabin temperature controller is shunted through limit switches in the bypass valves to the cool side of the heat-cool command relay, energizing the compressor clutch of the vapor cycle system and initiating vapor cycle system cooling. Heat commands, issued by the cabin temperature controller, cause the ACM and ejector bypass valves to open and duct uncooled bleed air into the cabin. When heating is required and the ACM bypass valve opens fully, the heat command from the cabin temperature controller is shunted to the heat side of the heat-cool relay and the vapor cycle system is cycled off. Electrical schematics and troubleshooting Charts in 21-60-00 will aid the technician in properly maintaining this system.

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21-10-00 00

ENVIRONMENTAL SYSTEMS BLEED AIR CONTROL DESCRIPTION AND OPERATION 1. GENERAL (UA-1 AND AFTER, UB-1 AND AFTER AND UC-1 THRU UC-45, EXCEPT UC-39) Bleed air from the P3 stage of the engine is precooled to 450 ± 25°F by the precooler heat exchanger, mounted immediately aft of the engine oil cooler, and pressure regulated to 38 ± 2 psig by the pressure regulator/shutoff valve (Ref. Figure 203, Maintenance Practices section). A precooler-bypass and a precooler-through valve modulate the amount of bleed air passing through the precooler. The two valves operate oppositely: when the precooler-through valve is opening, the bypass valve is closing and vice versa (Ref. Figure 201, Maintenance Practices section). When the bleed air control switch is turned on, all three environmental bleed air valves are energized (precooler-through valve, precooler-bypass valve and pressure regulator/shutoff valve) through the normally closed contacts of the bleed air valve deactivate relay. If bleed air pressure is adequate, a temperature controller assumes control of the two precooler valves and pneumatically controls the valves in varying degrees to maintain the specified bleed air temperature. A small surge tank, attached to the bleed air line, dampens any surges in bleed air pressure and provides a stable reference source for the overpressure limit switch. Should the overpressure limit of 44 ± 1 psig be exceeded due to a malfunction of the regulator/shutoff valve, the overpressure-limit switch will close. This energizes the bleed-air-valve-deactivate relay, thereby removing voltage from the bleed air valves and allowing them to close. When the overpressure switch closes, a signal is simultaneously transmitted to the annunciator system, illuminating the ENVIR FAIL annunciator. A temperature sensor attached to the bleed air line monitors bleed air temperature and transmits that information to the bleed air overtemperature detector (Ref. 21-10-07, Figure 201). The overtemperature detector will energize the bleed air valve-deactivate relay if the precooler system fails and allows the bleed air temperature to increase to 500°F ± 20. When a failure is detected, an ENVIR FAIL annunciator signal is simultaneously generated. A single bleed air overtemperature detector has inputs and outputs for both left and right bleed air control. Any time the bleed-air-deactivate relay is energized and opens the bleed-air-valve power circuits, a latching circuit will hold the relay in that mode. The bleed air-deactivate relay will be reset and the power circuits for the valves restored when the bleed air control switch is placed in the ENVIR OFF position. The instrument bleed air shutoff valves and the ejector heater of the cycle machine are also controlled through the bleed air control switches.

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ENVIRONMENTAL SYSTEMS BLEED AIR CONTROL TROUBLESHOOTING

100100

1. PROCEDURES (UA-1 AND AFTER, UB-1 AND AFTER AND UC-1 THRU UC-45, EXCEPT UC-39) Should an abnormal temperature be encountered, evaluation of the precooler system and controls is required. Abnormal bleed air pressures will be corrected by troubleshooting the pressure regulator/shutoff valve and controls. Table 101 presents a scheme of troubleshooting in a logical sequence of checks and should serve as a guide for the technician in narrowing down faults to a small portion of the overall system. Table 101 Troubleshooting - Bleed Air Control (UA-1 and After, UB-1 and After and UC-1 thru UC-45, except UC-39) PROBLEM

POSSIBLE CAUSE

1. Bleed air temperature out of limits a. Precooler valve(s) (E14), (E15) (Ref. BLEED AIR TEMPERATURE fails to operate properly. CHECK).

2. Bleed air pressure out of limits (Ref. BLEED AIR PRESSURE CHECK).

CORRECTIVE ACTION a. Perform PRECOOLER VALVES FUNCTIONAL CHECK procedure.

b. Temperature controller fails to operate properly.

b. Clean and inspect temperature controller as instructed in 1900 Airliner Series Component Maintenance Manual.

c. Overtemperature sensor (E112), (E113) fails to operate properly.

c. Replace overtemperature sensor.

d. Overtemperature detector module (E243) fails to operate properly.

d. Replace overtemperature detector module (E243).

a. Water in surge tank.

a. Drain surge tank.

b. Pressure regulator/shutoff valve(s) fails to operate properly.

b. If Janitrol valves are installed, replace pressure regulator/shutoff valve(s). If Dukes valves are installed, adjust or replace pressure regulator/shutoff valve(s) as required.

c. Bleed air control module (A317) fails to operate properly.

c. Replace bleed air control module (A317).

d. Overpressure switch(es) (S303, S304) fails to operate.

d. Replace overpressure switch(es) (S303, S304).

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200200

ENVIRONMENTAL SYSTEMS BLEED AIR CONTROL MAINTENANCE PRACTICES 1. PROCEDURES (UA-1 AND AFTER, UB-1 AND AFTER AND UC-1 THRU UC-45, EXCEPT UC-39) Maintenance of the environmental bleed air system will usually be limited to the replacing of valves and performing certain maintenance checks. Bleed air lines should be opened only as a last resort when necessity dictates. The precooler-bypass valves and their interconnecting network should be removed as an assembly. Space restrictions make it extremely difficult to perform any kind of maintenance, other than functional checks on the valves, while they are installed in the airplane.

A. Precooler Functional Check Air pressure, taken directly from the bleed air line, operates the precooler bypass and precooler-through valves. Solenoids on each valve control the valve in order to maintain the on/off states of the valves. A temperature controller modulates the air pressure, providing proportional control of the two precooler valves. On rising bleed air temperatures, the temperature controller reduces the air pressure which causes the precooler-through valve to begin opening and the bypass valve to begin closing; the inverse of this occurs on decreasing bleed air temperature (Ref. Figure 201). In the de-energized state, the through-valve solenoid closes and prevents the servo air from escaping through the temperature controller. The bypass valve opens when de-energized to relieve pressure on the through valve diaphragm, allowing the valve to spring-return to the closed position. A procedure for checking the function of the valves follows: (1) Set both bleed air valve switches to INST ENVIR OFF position. (2) Locate the flexible bleed air hose descending from the LH side of the engine and disconnect it from the manifold inlet of the precooler valve. (3) Fabricate an orifice plate 2.60 inches in diameter with a 0.064 to 0.071-inch orifice in the center of the plate. The orifice plate may be fabricated out of 0.032-inch-thick 2024-T3 Aluminum sheet. (4) Loosen the coupling on the duct just downstream of the temperature controller. (5) Insert the fabricated orifice plate between the flanges and reinstall the coupling; torque the clamp nut only enough to prevent excessive leakage. (6) Notice the valve position indicators on the valve bodies. The bypass valve should be closed and the precooler-through valve open. (7) Connect shop air pressure, regulated to 35 ± 5 psig, to the manifold inlet of the precooler valve. (8) Check the valve position indicators; both valves should be closed. (9) Apply 28 vdc at pin A of each precooler valve simultaneously to simulate the relay control when the system is normally on. The bypass valve should now be open and the precooler-through valve will remain closed. Should the bypass valve remain closed and the precooler-through valve open, the temperature controller is defective and should be replaced.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (10) Loosen the servo air connection on the temperature controller, allowing the servo air to escape. The bypass valve should now close and the precooler-through valve should open. (11) Should either valve fail to operate as indicated, remove the manifold assembly and precooler valve, then, replace the valve; otherwise, reconnect the flexible hose to the manifold inlet, torquing the clamp nut to 35 ± 2 inch-pounds, and replace the temperature controller located in the bleed air line just inboard of the valves. (12) Remove the orifice plate from the coupling downstream of the temperature controller and reinstall the clamp, torquing the clamp nut to 40 ± 2 inch-pounds.

Figure 201 Bleed Air Flow Schematic (With Janitrol Pressure Regulator/Shutoff Valve Shown) (UA-1 and After, UB-1 and After and UC-1 thru UC-45, Except UC-39)

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B. Temperature Check Bleed air temperatures are continuously monitored by the overtemperature detector. Temperature sensors, mounted in the bleed air lines, change resistance proportionate to bleed air temperature. Logic circuits in the overtemperature detector make use of this information and, through relay logic, close the pressure regulator/shutoff valve on the affected side (Ref. Figure 202). The following procedure determines if the bleed air temperature is within normal operating limits. Locate the overtemperature detector module in the subfloor, Zone 163 (or Zone 143 on later models). (1) Remove the connector (P118) from the detector module. (2) Operate the engines at 1550 rpm and 3000 foot-pounds torque. (3) Measure the resistance between pins 2 and 3 of the connector for the left engine or between pins 15 and 16 for the right engine. (4) Resistance should be between 445 and 515 ohms for a normally operating system. Any readings between 200 and 445 ohms or between 515 and 700 ohms indicates a fault in the precooler-bypass valve, precooler-through valve, or temperature controller. Any reading below 200 ohms or above 700 ohms indicates a faulty temperature sensor.

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Figure 202 Bleed Air Control Schematic (UA-1 and After, UB-1 and After and UC-1 thru UC-45, Except UC-39)

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C. Pressure Check Bleed air pressure is maintained at 38 ± 2 psig by the pressure regulator/shutoff valve (Ref. Figure 203). Failure of the valves to maintain proper pressure can be detected by measuring bleed air pressures. Two test ports located in the RH and LH wheel wells have been provided for this purpose. (1) Locate the bleed air test port in the wheel well and remove the cap. (2) Connect a 0-50 psi pressure gage accurate to within ± 1 psi to the test port. (3) Operate the engine at 1550 rpm and 3000 foot-pounds torque. NOTE: There are no adjustments on Janitrol pressure regulator/shutoff valves. For adjustment of Dukes pressure regulator/shutoff valves, refer to DUKES PRESSURE REGULATOR/ SHUTOFF VALVE ADJUSTMENT, in this section. (4) If bleed air pressure does not stabilize at 38 ± 2 psig, the pressure regulator/shutoff valve should be adjusted or replaced.

Figure 203 Janitrol Pressure Regulator/Shutoff Valve (UA-1 and After, UB-1 and After and UC-1 thru UC-45, Except UC-39 Without Kit No. 114-5037 Installed)

21-10-00

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D. Pressure Operational Check The operational overpressure test may be accomplished as follows: (1) Rotate the switch to the P position. (2) Run the engines up to 85% power. (3) The ENVIR FAIL and MASTER WARNING annunciator lights should illuminate within 60 seconds. (4) The ENVIR OFF annunciator light should illuminate.

E. Temperature Operational Check The operational overtemperature test may be accomplished as follows: (1) Rotate the switch to the T position. (2) Run the engines up to 85% power. (3) The ENVIR FAIL and MASTER WARNING annunciator lights should illuminate within 60 seconds. (4) The ENVIR OFF annunciator light should illuminate.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

ENVIRONMENTAL SYSTEMS PRECOOLER - THROUGH AND BYPASS VALVES MAINTENANCE PRACTICES

21-10-01 200200

1. PROCEDURES (UA-1 AND AFTER, UB-1 AND AFTER AND UC-1 THRU UC-45, EXCEPT UC-39) A. Removal (1) Pull the BLEED AIR CONTROL circuit breakers under ENVIRONMENTAL on the circuit breaker panel (Ref. Figure 201). (2) Remove the forward upper and lower cowlings (Ref. Chapter 71-10-00). (3) Remove the oil cooler inlet ducts and the aft lower cowlings (Ref. Chapter 71-10-00). (4) Disconnect the fuel and lubricant drain lines as necessary to gain access to the valves. (5) Disconnect the electrical connectors from the valves. (6) Disconnect the pressure-drive line and temperature control lines. (7) Remove the couplings from each side of the valve. (8) Support the valves and remove four bolts and washers from each support bracket. NOTE: The outboard support brackets may be removed if necessary to aid in removing the valves.

B. Installation (1) Position the valve between the support brackets and loosely secure each with four bolts and washers (Ref. Figure 201). (2) Install couplings over the flanges of the ducts and valves. Tighten the clamping nuts to 35 ± 2 inch-pounds. NOTE: Do not install insulation over clamps. (3) Tighten the valve attachment bolts. (4) Connect the pressure drive-lines and temperature control lines. (5) Connect electrical connectors. (6) Connect any drain lines that were disconnected. (7) Install oil cooler inlet ducts. (8) Install the cowlings (Ref. Chapter 71-10-00). (9) Reset BLEED AIR CONTROL circuit breakers.

21-10-01

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Figure 201 Precooler-Through and Bypass Valves Installation (Typical) (UC-1 thru UC-45, Except UC-39)

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21-10-01

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

ENVIRONMENTAL SYSTEMS PRESSURE REGULATOR/SHUTOFF VALVE MAINTENANCE PRACTICES

21-10-03 200200

1. JANITROL PRESSURE REGULATOR/SHUTOFF VALVE PROCEDURES (UA-1 AND AFTER, UB-1 AND AFTER AND UC-1 THRU UC-45, EXCEPT UC-39) (WITHOUT KIT NO. 114-5037 INSTALLED) A. Removal The right and left pressure regulator/shutoff valves are not symmetrically located in the wings (Ref. Figure 201). Removal of the right valve requires removing the leading edge only. Refer to the Chapter 6-50-00 WING ACCESS PANELS (UA-1 and After; UB-1 and After) illustration in the Dimensions and Areas section. Removal of the left valve requires removing the leading edge first, then the large upper panel and the large lower panel aft of the leading edge. Refer to the Chapter 6-50-00 WING ACCESS PANELS (UC-1 and After) illustration in the Dimensions and Areas section. Any aircraft modified by Kit No. 114-5016 have a filter added in the sense line for the pressure regulator/ shutoff valve. On the right pressure regulator/shutoff valve, the filter is mounted on a bracket attached directly to the valve. (1) Set the aircraft MASTER SWITCH TO OFF. (2) Remove the panels as required to gain access to the valve. (3) Remove the sense line as follows: (a) Remove the sense line between the valve and the P3 air duct. (b) On aircraft with filters installed, remove the sense line between the left valve and the filter. On the right side, remove both sense lines of the filter, then remove and retain the filter and bracket attached to the inboard side of the valve. (4) Disconnect the electrical connector. (5) Disconnect the valve mounting bracket as follows: (a) On the left valve, the mounting bolts are accessed from inside the left wheel well. Remove the three bolts securing the valve mounting bracket to the nacelle bulkhead. (b) On the right valve, remove the two screws securing the mounting bracket to the wing stringer. (6) Pull back the insulation from the P3 air duct and remove the two couplings holding the valve in the duct. (7) Remove the valve with the mounting bracket attached. (8) Remove and retain the mounting bracket from the valve.

21-10-03

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B. Installation (1) Install on the valve the mounting bracket retained from the removal procedures (Ref. Figure 201). (2) Set the valve in the P3 air duct and install the couplings. (3) Secure the mounting bracket as follows: (a) On the left valve, install the three bolts through the nacelle bulkhead. (b) On the right valve, install the two screws to secure the mounting bracket to the wing stringer. (4) Mold the insulation around the edges of the valve. (5) Connect the electrical connector. (6) Install the sense line as follows: (a) Install the sense line between the valve and the P3 air duct. (b) On aircraft with filters installed, connect the sense line between the left valve and the filter. On the right valve, install the filter and mounting bracket on the valve, then connect the sense lines of the filter to the valve and P3 air duct. (7) Replace all panels.

Figure 201 Janitrol Pressure Regulator/Shutoff Valve (UA-1 and After, UB-1 and After and UC-1 thru UC-45, Except UC-39 Without Kit No. 114-5037 Installed)

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

C. Filter Removal The right and left filters are not symmetrically located in the wings. The left filter is mounted on a wing rib adjacent to the left pressure regulator/shutoff valve. The right filter is mounted on a bracket attached directly to the right pressure regulator/shutoff valve. Removal or servicing the right filter requires removing the leading edge only. Refer to the Chapter 6-50-00 WING ACCESS PANELS (UA-1 and After; UB-1 and After) illustration in the Dimensions and Areas section. Removal or servicing the left filter requires removing the leading edge and the large lower panel aft of the leading edge. Refer to the Chapter 6-50-00 WING ACCESS PANELS (UC-1 and After) illustration in the Dimensions and Areas section. (1) Set the aircraft MASTER SWITCH to OFF. (2) Remove panels as required to gain access to the filter. (3) Remove the sense line between the filter and the P3 air duct. (4) Remove the sense line between the filter and the pressure regulator/shutoff valve. (5) Remove the filter as follows: (a) In the left wing, remove the two screws securing the filter to the wing rib. (b) In the right wing, remove the two screws securing the filter to the bracket mounted on the pressure regulator/shutoff valve.

D. Filter Installation (1) Install the filter as follows: (a) In the left wing, secure the filter to the wing rib with two screws. (b) In the right wing, secure the filter to the bracket on the pressure regulator/shutoff valve. (2) Install the sense line between the filter and the P3 air duct. (3) Install the sense line between the filter and the pressure regulator/shutoff valve. (4) Replace all panels.

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2. DUKES PRESSURE REGULATOR/SHUTOFF VALVE PROCEDURES (UA-1 AND AFTER, UB-1 AND AFTER AND UC-1 THRU UC-45, EXCEPT UC-39) (WITH KIT NO. 114-5037 INSTALLED) A. Removal (1) Disconnect the battery and/or remove all electrical power from the airplane. (2) Gain access to the pressure regulator/shutoff valve by removing the appropriate wing panels (Ref. Figure 202). (a) To gain access to the left pressure regulator/shutoff valve, remove the left inboard wing leading edge panel and the lower wing panels. (b) To gain access to the right pressure regulator/shutoff valve, remove the right inboard wing leading edge panel. (3) Remove the two couplings holding the valve in the duct. (4) Remove the pressure regulator/shutoff valve from the airplane.

B. Installation (1) Install the pressure regulator/shutoff valve in the airplane using the two couplings to hold the valve in the P3 air duct (Ref. Figure 202). NOTE: Observe the proper flow direction when installing the valve. If the valve is being installed on the left side of the airplane, the arrow on the side of the valve should be pointing aft. If the valve is being installed on the right side of the airplane, the arrow on the side of the valve should be pointing inboard. (2) Torque the clamp nuts on the couplings to 35 ± 2 inch-pounds. Do not overtorque. (3) Replace all panels. (4) Reconnect the battery and/or restore electrical power to the airplane. (5) Check the system for proper operation. Perform BLEED AIR PRESSURE CHECK and BLEED AIR PRESSURE OPERATIONAL CHECK, in this section.

C. Adjustment If the bleed air pressure is not between 36 and 40 psig, the pressure regulator/shutoff valve must be adjusted as follows (Ref. Figure 202): WARNING: Do not adjust the pressure regulator/shutoff valve while the engine is running. (1) Install a pressure gage and measure bleed air pressure. Refer to the BLEED AIR PRESSURE CHECK procedure, in this section. (2) When the right valve needs to be adjusted, remove the right inboard leading edge wing panel. (3) When the left valve needs to be adjusted, remove the left upper inboard wing panel.

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21-10-03

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (4) Locate the pressure regulator/shutoff valve and remove the safety wire from the jam nut on the bellows housing. (5) While supporting the pilot regulator, loosen the large jam nut on the bellows housing. NOTE: Do not loosen the smaller jam nut located on the opposite end of the pressure regulator/ shutoff valve. If the smaller jam nut is moved, the valve can not be adjusted properly. (6) The bellows housing is rotated to adjust the pressure. One-eighth turn (45°) of the bellows housing changes the regulated pressure approximately 4 to 5 psi. Do not make any adjustment of more than one-quarter turn (90°) at a time. (a) To decrease the pressure, rotate the bellows housing clockwise (screw in). (b) To increase the pressure, rotate the bellows housing counterclockwise (screw out). (7) Tighten the jam nut while supporting the bellows housing to prevent the adjustment from being altered. CAUTION: The right inboard leading edge wing panel must be in place when the right engine is run. The left engine may be run without replacing the left upper inboard wing panel. (8) Operate the engine at 1540 rpm and 3000 foot-pounds of torque. If the bleed air pressure is not between 36 and 40 psig, shut down the engine and make additional adjustments to the valve. (9) If the valve does not respond to this adjustment procedure, it must be replaced. (10) Install safety wire through the jam nuts at each end of the pilot regulator. (11) Replace any wing panels that were removed. Check that the leading edge deicer boot plumbing is properly connected. (12) Remove the pressure test gage and replace the test port cap.

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Figure 202 Dukes Pressure Regulator/Shutoff Valves Adjustment (UA-1 and After, UB-1 and After and UC-1 thru UC-45, Except UC-39 With Kit No. 114-5037 Installed)

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

ENVIRONMENTAL SYSTEMS ACM OVERPRESSURE SWITCH MAINTENANCE PRACTICES

21-10-06 200200

1. PROCEDURES (UA-1 AND AFTER, UB-1 AND AFTER AND UC-1 THRU UC-45, EXCEPT UC-39) A. Removal The right and left overpressure switches are not symmetrically located in the wings, but are adjacent to their respective pressure regulator/shutoff valves (Ref. Figure 201). Removal of the right overpressure switch requires removing the leading edge only. Refer to the Chapter 6-50-00 WING ACCESS PANELS (UA-1 and After; UB-1 and After) illustration in the Dimensions and Areas section. Removal of the left switch requires removing the leading edge and the large lower panel aft of the leading edge. Refer to the Chapter 6-50-00 WING ACCESS PANELS (UC-1 and After) illustration in the Dimensions and Areas section. (1) Set the aircraft MASTER SWITCH to OFF. (2) Remove panels as required to gain access to the overpressure switch. (3) Disconnect the self-locking electrical connector, located in the wire bundle. (4) Remove the sense line between the switch and surge tank. (5) Remove the screw securing the overpressure switch and clamp in the wing. (6) Remove and retain the clamp from the overpressure switch.

B. Installation (1) Set the overpressure switch in the clamp retained from the removal procedures (Ref. Figure 201). (2) Install the overpressure switch in the wing. (3) Install the sense line between the switch and surge tank. (4) Connect the electrical connector. (5) Replace all panels.

21-10-06

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Figure 201 Left Overpressure Switch (UA-1 and After, UB-1 and After and UC-1 thru UC-45, Except UC-39)

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21-10-06

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

ENVIRONMENTAL SYSTEMS ACM OVERTEMPERATURE SENSOR MAINTENANCE PRACTICES

21-10-07 200200

1. PROCEDURES (UA-1 AND AFTER, UB-1 AND AFTER AND UC-1 THRU UC-45, EXCEPT UC-39) A. Removal The right and left temperature sensors are not symmetrically located in the wings, but are adjacent to the respective pressure regulator/shutoff valves (Ref. Figure 201). Removal of the right temperature sensor requires removing the leading edge only. Refer to the Chapter 6-50-00 WING ACCESS PANELS (UA-1 and After; UB-1 and After) illustration in the Dimensions and Areas section. Removal of the left sensor requires removing the leading edge and the large lower panel aft of the leading edge. Refer to the Chapter 6-50-00 WING ACCESS PANELS (UC-1 and After) illustration in the Dimensions and Areas section. (1) Set the aircraft MASTER SWITCH to OFF. (2) Disconnect the electrical connector. (3) Remove the four screws securing the sensor in the P3 air duct. (4) Remove the sensor and gasket. (5) Retain the gasket.

B. Installation (1) Install gasket retained from removal procedures on temperature sensor (Ref. Figure 201). (2) Set the sensor in the P3 air duct and secure with four screws. (3) Connect and safety the electrical connector. (4) Replace all panels.

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Figure 201 Right Temperature Sensor (UA-1 and After, UB-1 and After and UC-1 thru UC-45, Except UC-39)

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ENVIRONMENTAL SYSTEMS BLEED AIR CONTROL DESCRIPTION AND OPERATION 25

21-11-00 00

1. GENERAL - (UC-39, UC-46 AND AFTER) This subchapter covers those bleed air control systems currently built or ones that have been modified with kits as described below: On UC-39, UC-46 and After, and any airplanes with Kit No. 114-5016 and 114-5018 installed, an additional pressure switch and temperature sensor are installed near the ACM. A filter is also added in the pressure regulator/shutoff valve sense line and two smaller filters are added in the pressure line from the temperature controller to the precooler valves for each side. The overpressure and over temperature functions are slightly different in that a new PCB is used for relay control of the pneumatic valves. Bleed air from the P3 stage of the engine is maintained at 450° ± 25°F by the precooler heat exchanger and control valves. Bleed air is pressure regulated to 38 ± 2 psig by the pressure regulator/shutoff valve (PRSOV). Bleed air for instrumentation is controlled by the normally open pneumatic bleed air shutoff valve (Ref. Figures 201 and 202, Maintenance Practices section).

A. Bleed Air Valves Switch Each BLEED AIR VALVE switch controls four valves. The pneumatic bleed air shutoff valve is held shut by voltage when the control switch is in the INST ENVIR OFF position. The pressure regulator/shutoff valve and precooler valves are controlled by relay logic actuated by the control switch and use bleed air to operate with. The ACM heater is turned on by either the left or right switch.

B. Bleed Air Control Module Relay logic is provided by the A317 Bleed Air Control Module PCB. Inputs from the control switch, pressure switches and over temperature detector will shut the valves on the selected side. When a pressure or temperature fault is detected by the pcb, 28 vdc is applied to the warning legend annunciator. LEDs on the pcb light up when a fault has been encountered. Red LEDs for the right or left side indicates a pressure fault, and green LEDs indicates a temperature fault.

C. Precooler Valves A precooler bypass and a precooler-thru valve modulate the amount of bleed air passing through the precooler. The two valves operate oppositely; when the thru valve is opening, the bypass valve is closing and vise versa. If bleed air pressure is adequate, a temperature controller assumes control of the two precooler valves and pneumatically controls the valves in varying degrees in order to maintain the specified bleed air temperature.

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D. Overpressure Three overpressure switches monitor bleed air pressure, one for each side and a third located near the air cycle machine (ACM). The ACM pressure switch has two sets of contacts, one for each side of the system, which close at 40 ± 1 psi on increasing pressure to indicate an overpressure condition. The left and right pressure switches close at 44 ± 1 psi on increasing pressure to indicate an over pressure condition. An overpressure signal to the A317 PCB will occur only when one side and the ACM pressure switch close. When an overpressure is reported, the relay logic removes voltage applied to the precooler valves and the pressure regulator/shutoff valve on the affected side. The same signal and relay logic will illuminate the L/R ENVIR FAIL warning annunciator. Anytime the pressure regulator/ shutoff valve closes, the L/R ENVIR OFF caution/advisory annunciator illuminates. A surge tank between each left and right switch dampens pressure surges and provides a stable bleed air reference.

E. Over temperature Three temperature sensors monitor bleed air temperature and report it to the overtemperature detector. Like the pressure switches, there is one for each side and one near the ACM. Signals from the overtemperature detector actuate the relay logic to remove voltage applied to the precooler valves and the pressure regulator/shutoff valve on the affected side. The same signal and relay logic will illuminate the L/R ENVIR FAIL warning annunciator. Anytime the pressure regulator/shutoff valve closes, the L/R ENVIR OFF caution/advisory annunciator illuminates.

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ENVIRONMENTAL SYSTEMS BLEED AIR CONTROL TROUBLESHOOTING

100100

1. GENERAL - (UC-39, UC-46 AND AFTER) Should an abnormal temperature be encountered, evaluation of the precooler system and controls is required. Abnormal bleed air pressures will be corrected by troubleshooting the pressure regulator/shutoff valve and control (Ref. Figures 201 and 202, Maintenance Practices section). Table 101 presents a scheme of troubleshooting in a logical sequence of checks and should serve as a guide for the technician in narrowing down faults to a small portion of the overall system. Table 101 Troubleshooting - Bleed Air Control (UC-39, UC-46 and After) PROBLEM 1. Bleed air temperature out of limits. (Ref. BLEED AIR TEMPERATURE CHECK)

2. Bleed air pressure out of limits. (Ref. BLEED AIR PRESSURE CHECK).

POSSIBLE CAUSE

CORRECTIVE ACTION

a. Precooler valve(s) (E14),(E15) fails to operate properly.

a. Perform PRECOOLER VALVES FUNCTIONAL CHECK procedure.

b. Temperature controller fails to operate properly.

b. Clean and inspect temperature controller as instructed in 1900 Airliner Series Component Maintenance Manual.

c. Overtemperature sensor (E112), (E113), (E303) fails to operate properly.

c. Replace overtemperature sensor.

d. Overtemperature detector module (E243) fails to operate properly.

d. Replace overtemperature detector module (E243).

a. Water in surge tank.

a. Drain surge tank.

b. Pressure regulator/shutoff valve(s) fails to operate properly.

b. If Janitrol valves are installed, replace pressure regulator/shutoff valve(s). If Dukes valves are installed, adjust or replace pressure regulator/shutoff valve(s) as required.

c. Bleed air control module (A317) fails to operate properly.

c. Replace bleed air control module (A317).

d. Overpressure switch(es) (S303,S304 or S315) fails to operate.

d. Replace overpressure switch(es) (S303,S304 or S315).

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ENVIRONMENTAL SYSTEMS BLEED AIR CONTROL MAINTENANCE PRACTICES 1. PROCEDURES - (UC-39, UC-46 AND AFTER) Maintenance of the environmental bleed air system will usually be limited to the replacing of valves and performing certain maintenance checks. Bleed air lines should be opened only as a last resort when necessity dictates. The precooler-bypass valves and their interconnecting network should be removed as an assembly. Space restrictions make it extremely difficult to perform any kind of maintenance, other than functional checks on the valves, while they are installed in the airplane.

A. Precooler Functional Check Air pressure, taken directly from the bleed air line, operates the precooler bypass and precooler-through valves. Solenoids on each valve control the valve in order to maintain the on/off states of the valves. A temperature controller modulates the air pressure, providing proportional control of the two precooler valves. On rising bleed air temperatures, the temperature controller reduces the air pressure which causes the precooler-through valve to begin opening and the bypass valve to begin closing; the inverse of this occurs on decreasing bleed air temperature (Ref. Figure 201). In the de-energized state, the through-valve solenoid closes and prevents the servo air from escaping through the temperature controller. The bypass valve opens when de-energized to relieve pressure on the through valve diaphragm, allowing the valve to spring-return to the closed position. A procedure for checking the function of the valves follows: (1) Set both bleed air valve switches to INST ENVIR OFF position. (2) Locate the flexible bleed air hose descending from the LH side of the engine and disconnect it from the manifold inlet of the precooler valve. (3) Fabricate an orifice plate 2.60-inches in diameter with a 0.064 to 0.071-inch orifice in the center of the plate. The orifice plate may be fabricated out of 0.032-inch-thick 2024-T3 Aluminum sheet. (4) Loosen the coupling on the duct just downstream of the temperature controller. (5) Insert the fabricated orifice plate between the flanges and reinstall the coupling; torque the clamp nut only enough to prevent excessive leakage. (6) Notice the valve position indicators on the valve bodies. The bypass valve should be closed and the precooler-through valve open. (7) Connect shop air pressure, regulated to 35 ± 5 psig, to the manifold inlet of the precooler valve. (8) Check the valve position indicators; both valves should be closed. (9) Apply 28 vdc at pin A of each precooler valve simultaneously to simulate the relay control when the system is normally on. The bypass valve should now be open and the precooler-through valve will remain closed. Should the bypass valve remain closed and the precooler-through valve open, the temperature controller is defective and should be replaced.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (10) Loosen the servo air connection on the temperature controller, allowing the servo air to escape. The bypass valve should now close and the precooler-through valve should open. (11) Should either valve fail to operate as indicated, remove the manifold assembly and precooler valve, then, replace the valve; otherwise, reconnect the flexible hose to the manifold inlet, torquing the clamp nut to 35 ± 2-inch-pounds, and replace the temperature controller located in the bleed air line just inboard of the valves. (12) Remove the orifice plate from the coupling downstream of the temperature controller and install the clamp, torquing the clamp nut to 40 ± 2-inch-pounds.

Figure 201 Bleed Air Flow Schematic (With Janitrol Pressure Regulator/Shutoff Valve Shown) (UC-39, UC- 46 and After)

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B. Temperature Check Bleed air temperatures are continuously monitored by the overtemperature detector. Temperature sensors, mounted in the bleed air lines, change resistance proportionate to bleed air temperature. Logic circuits in the overtemperature detector make use of this information and, through relay logic, close the pressure regulator/shutoff valve on the affected side (Ref. Figure 202). The following procedure determines if the bleed air temperature is within normal operating limits. Locate the overtemperature detector module in the subfloor, Zone 163. (1) Remove the connector from the detector module. (2) Operate the engines at 1550 rpm and 3000 foot-pounds torque. (3) Measure the resistance between pins 2 and 3 of the connector for the left engine, between pins 15 and 16 for the right engine and between pins 8 and 9 for the ACM. (4) Resistance should be between 445 and 515 ohms for a normally operating system. Any readings between 200 and 445 ohms or between 515 and 700 ohms indicates a fault in the precooler-bypass valve, precooler-through valve, or temperature controller. Any reading below 200 ohms or above 700 ohms indicates a faulty temperature sensor (Ref. Figure 202).

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UC21B 031451AA.AI

Figure 202 (Sheet 1 of 3) Bleed Air Control (UC-39, UC-46 and After)

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UE21B 991314AA.AI

Figure 202 (Sheet 2 of 3) Bleed Air Control (UC-39, UC-46 and After)

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Figure 202 (Sheet 3 of 3) Bleed Air Control (UC-39, UC-46 and After)

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C. Pressure Check Bleed air pressure is maintained at 38 ± 2 psig by the pressure regulator/shutoff valve. Failure of the valves to maintain proper pressure can be detected by measuring bleed air pressures. Two test ports located in the RH and LH wheel wells have been provided for this purpose. (1) Locate the bleed air test port in the wheel well and remove the cap. (2) Connect a 0 to 50 psi pressure gage accurate to within ± 1 psi to the test port. (3) Operate the engine at 1550 rpm and 3000 foot-pounds torque. NOTE: There are no adjustments on Janitrol pressure regulator/shutoff valves. For adjustment of Dukes pressure regulator/shutoff valves (DUKES PRESSURE REGULATOR/SHUTOFF VALVE ADJUSTMENT). (4) If bleed air pressure does not stabilize at 38 ± 2 psig, the pressure regulator/shutoff valve should be adjusted or replaced.

D. Pressure Operational Check The operational overpressure test may be accomplished as follows: (1) Rotate the switch to the P position. (2) Run the engines up to 85% power. (3) The ENVIR FAIL and MASTER WARNING annunciator lights should illuminate within 60 seconds. (4) The ENVIR OFF annunciator light should illuminate.

E. Temperature Operational Check The operational overtemperature test may be accomplished as follows: (1) Rotate the switch to the T position. (2) Run the engines up to 85% power. (3) The ENVIR FAIL and MASTER WARNING annunciator lights should illuminate within 60 seconds. (4) The ENVIR OFF annunciator light should illuminate.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

ENVIRONMENTAL SYSTEMS PRECOOLER - THROUGH AND BYPASS VALVES MAINTENANCE PRACTICES

21-11-01 200200

1. PROCEDURES - (UC-39, UC-46 AND AFTER) A. Removal (1) Pull the BLEED AIR CONTROL circuit breakers under ENVIRONMENTAL on the circuit breaker panel. (2) Remove the forward upper and lower cowlings (Ref. Chapter 71-10-00). (3) Remove the oil cooler inlet ducts and the aft lower cowlings (Ref. Chapter 71-10-00). (4) Disconnect the fuel and lubricant drain lines as necessary to gain access to the valves. (5) Disconnect the electrical connectors from the valves. (6) Disconnect the pressure-drive line and temperature control lines. (7) Remove the couplings from each side of the valve. NOTE: The outboard support brackets may be removed if necessary to aid in removing the valves (Ref. Figure 201). (8) Support the valves and remove four bolts and washers from each support bracket.

B. Installation (1) Position the valve between the support brackets and loosely secure each with four bolts and washers (Ref. Figure 201). NOTE: Do not install insulation over clamps. (2) Install couplings over the flanges of the ducts and valves. Tighten the clamping nuts to 35 ± 2 inch-pounds. (3) Tighten the valve attachment bolts. (4) Connect the pressure drive-lines and temperature control lines. (5) Connect electrical connectors. (6) Connect any drain lines that were disconnected. (7) Install oil cooler inlet ducts. (8) Install the cowlings (Ref. Chapter 71-10-00). (9) Reset BLEED AIR CONTROL circuit breakers.

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Figure 201 Precooler-Through and Bypass Valves Installation (Typical) (UC-39, UC-46 and After)

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ENVIRONMENTAL SYSTEMS TEMPERATURE CONTROLLER SENSE LINE FILTER MAINTENANCE PRACTICES

21-11-02 200200

1. PROCEDURES - (UC-39, UC-46 AND AFTER) A. Removal Two filters (Zones 521 and 621) are located in the sense line of the temperature controller that connect the precooler-bypass and thru valves in each nacelle (Ref. Figure 201). Refer to the Chapter 6-40-00 ZONE DIAGRAMS illustration in the Dimension and Areas section. (1) Remove the inboard nacelle panel. (2) Gain access to the temperature controller as follows: (a) On earlier UC serial airplanes, disconnect the oil and vent lines and remove the plate covering the precooler section. (b) On later UC serial airplanes, a double plate is used. Remove the forward plate. (3) Disconnect the sense line from the temperature controller. (4) Remove the filter from each end of the sense line.

B. Installation (1) Install a filter in each end of the sense line. (2) Install the sense line between the temperature controller and the T-fitting to the precooler bypass and thru valve (Ref. Figure 201). (3) Install the plate that covers the precooler section and connect the oil and drain lines, if removed. (4) Install the nacelle cover.

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Figure 201 Temperature Control Line Filter Installation (Typical) (UC-39, UC-46 and After)

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

ENVIRONMENTAL SYSTEMS PRESSURE REGULATOR/SHUTOFF VALVE MAINTENANCE PRACTICES

21-11-03 200200

1. JANITROL PRESSURE REGULATOR/SHUTOFF VALVE PROCEDURES (UC-39, UC-46 AND AFTER) (WITHOUT KIT NO. 114-5037 INSTALLED) A. Removal The right and left pressure regulator/shutoff valves are not symmetrically located in the wings. Removal of the right valve requires removing the leading edge panel 23. Removal of the left valve requires removing the leading edge panel 24, then the large upper panel and large lower panel aft of the leading edge. Refer to the Chapter 6-50-00 WING ACCESS PANELS (UC-1 and After) illustration in the Dimensions and Areas section. A filter is located in the sense line for the pressure regulator/shutoff valve. On the right pressure regulator/shutoff valve, the filter is mounted on a bracket attached directly to the valve. (1) Set the aircraft MASTER SWITCH TO OFF. (2) Remove the panels as required to gain access to the valve. (3) Remove the sense line as follows: (a) Remove the sense line between the valve and the P3 air duct. (b) Remove the sense line between the left valve and the filter. On the right side, remove both sense lines of the filter, then remove and retain the filter and bracket attached to the inboard side of the valve. (4) Disconnect the electrical connector. (5) Disconnect the valve mounting bracket as follows: (a) On the left valve, the mounting bolts are accessed from inside the left wheel well. Remove the three bolts securing the valve mounting bracket to the nacelle bulkhead. (b) On the right valve, remove the two screws securing the mounting bracket to the wing stringer. (6) Pull back the insulation from the P3 air duct and remove the two couplings holding the valve in the duct. (7) Remove the valve with the mounting bracket attached. (8) Remove and retain the mounting bracket from the valve.

B. Installation (1) Install on the valve the mounting bracket retained from the removal procedures. (2) Set the valve in the P3 air duct and install the couplings. (3) Secure the mounting bracket as follows:

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (a) On the left valve, install the three bolts through the nacelle bulkhead. (b) On the right valve, install the two screws to secure the mounting bracket to the wing stringer. (4) Mold the insulation around the edges of the valve. (5) Connect the electrical connector. (6) Install the sense line as follows: (a) Install the sense line between the valve and the P3 air duct. (b) Connect the sense line between the left valve and the filter. On the right valve, install the filter and mounting bracket on the valve, then connect the sense lines of the filter to the valve and P3. (7) Replace all panels.

C. Filter Removal The right and left filters are not symmetrically located in the wings. The left filter is mounted on a wing rib adjacent to the left pressure regulator/shutoff valve. The right filter is mounted on a bracket attached directly to the right pressure regulator/shutoff valve. Removal or servicing the right filter requires removing the leading edge panel 23. Removal or servicing the left filter requires removing the leading edge panel 24 and the large lower panel aft of the leading edge. Refer to the Chapter 6-50-00 WING ACCESS PANELS (UC-1 and After) illustration in the Dimensions and Areas section. (1) Set the aircraft MASTER SWITCH to OFF. (2) Remove panels as required to gain access to the filter. (3) Remove the sense line between the filter and the P3 air duct. (4) Remove the sense line between the filter and the pressure regulator/shutoff valve. (5) Remove the filter as follows: (a) In the left wing, remove the two screws securing the filter to the wing rib. (b) In the right wing, remove the two screws securing the filter to the bracket mounted on the pressure regulator/shutoff valve.

D. Filter Installation (1) Install the filter as follows: (a) In the left wing, secure the filter to the wing rib with two screws. (b) In the right wing, secure the filter to the bracket on the pressure regulator/shutoff valve. (2) Install the sense line between the filter and the P3 air duct. (3) Install the sense line between the filter and the pressure regulator/shutoff valve. (4) Replace all panels.

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2. DUKES PRESSURE REGULATOR/SHUTOFF VALVE PROCEDURES (UC-39, UC-46 AND AFTER) (WITH KIT NO. 114-5037 INSTALLED) A. Removal (1) Disconnect the battery and/or remove all electrical power from the airplane. (2) Gain access to the pressure regulator/shutoff valve by removing the appropriate wing panels. (a) To gain access to the left pressure regulator/shutoff valve, remove the left inboard wing leading edge panel and the lower wing panels. (b) To gain access to the right pressure regulator/shutoff valve, remove the right inboard wing leading edge panel. (3) Remove the two couplings holding the valve in the duct, (Ref. Figure 201). (4) Remove the pressure regulator/shutoff valve from the airplane.

B. Installation (1) Install the pressure regulator/shutoff valve in the airplane using the two couplings to hold the valve in the P3 air duct (Ref. Figure 201). NOTE: Observe the proper flow direction when installing the valve. If the valve is being installed on the left side of the airplane, the arrow on the side of the valve should be pointing aft. If the valve is being installed on the right side of the airplane, the arrow on the side of the valve should be pointing inboard. (2) Torque the clamp nuts on the couplings to 35 ± 2-inch-pounds. Do not overtorque. (3) Replace all panels. (4) Connect the battery and/or restore electrical power to the airplane. (5) Check the system for proper operation (Ref. 21-11-00, BLEED AIR PRESSURE CHECK and BLEED AIR PRESSURE OPERATIONAL CHECK).

C. Adjustment If the bleed air pressure is not between 36 to 40 psig, the pressure regulator/shutoff valve must be adjusted as follows: WARNING: Do not adjust the pressure regulator/shutoff valve while the engine is running. (1) Install a pressure gage and measure bleed air pressure (Ref. 21-11-00, BLEED AIR PRESSURE CHECK). (2) When the right valve needs to be adjusted, remove the right inboard leading edge wing panel. (3) When the left valve needs to be adjusted, remove the left upper inboard wing panel. (4) Locate the pressure regulator/shutoff valve and remove the safety wire from the jam nut on the bellows housing.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL NOTE: Do not loosen the smaller jam nut located on the opposite end of the pressure regulator/ shutoff valve. If the smaller jam nut is moved, the valve can not be adjusted properly. (5) While supporting the pilot regulator, loosen the large jam nut on the bellows housing. (6) The bellows housing is rotated to adjust the pressure. One-eighth turn (45°) of the bellows housing changes the regulated pressure approximately 4 to 5 psi. Do not make any adjustment of more than one-quarter turn (90°) at a time. (a) To decrease the pressure, rotate the bellows housing clockwise (screw in). (b) To increase the pressure, rotate the bellows housing counterclockwise (screw out). (7) Tighten the jam nut while supporting the bellows housing to prevent the adjustment from being altered. CAUTION: The right inboard leading edge wing panel must be in place when the right engine is run. The left engine may be running without replacing the left upper inboard wing panel. (8) Operate the engine at 1540 rpm and 3000 foot-pounds of torque. If the bleed air pressure is not between 36 and 40 psig, shut down the engine and make additional adjustments to the valve. (9) If the valve does not respond to this adjustment procedure, it must be replaced. (10) Install safety wire (178, Table 1, Chapter 91-00-00) through the jam nuts at each end of the pilot regulator. (11) Replace any wing panels that were removed. Check that the leading edge deicer boot plumbing is properly connected. (12) Remove the pressure test gage and replace the test port cap.

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Figure 201 Dukes Pressure Regulator/Shutoff Valves Adjustment (UC-39, UC-46 and After With Kit No. 114-5037 Installed)

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ENVIRONMENTAL SYSTEMS ACM OVERPRESSURE SWITCH MAINTENANCE PRACTICES

21-11-06 200200

1. PROCEDURES - (UC-39, UC-46 AND AFTER) A. Removal The ACM overpressure switch is located in the ACM P3 air duct. Removal of the switch requires removing the leading edge panel 24 and then the large upper panel 18. Refer to the Chapter 6-50-00 WING ACCESS PANELS (UC-1 and After) illustration in the Dimensions and Areas section. (1) Set the aircraft MASTER SWITCH to OFF. (2) Remove the panels as required to gain access to the overpressure switch. (3) Disconnect the self-locking electrical connector, located in the wire bundle. (4) Remove the sense line between the switch and the ACM P3 air duct. (5) Remove the screw securing the overpressure switch and clamp in the wing. (6) Remove and retain the clamp from the overpressure switch.

B. Installation (1) Set the overpressure switch in the clamp retained from the removal procedures. (2) Install the overpressure switch in the wing. (3) Install the sense line between the switch and the ACM P 3 air duct. (4) Connect and safety the electrical connector. (5) Replace all panels.

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ENVIRONMENTAL SYSTEMS ACM OVERTEMPERATURE SENSOR MAINTENANCE PRACTICES

21-11-07 200200

1. PROCEDURES - (UC-39, UC-46 AND AFTER) A. Removal The ACM overtemperature sensor is located in the ACM P3 air duct. Removal of the sensor requires removing the leading edge panel 24 and then the large upper panel 18. Refer to the Chapter 6-50-00 WING ACCESS PANELS (UC-1 and After) illustration in the Dimensions and Areas section. (1) Set the aircraft MASTER SWITCH to OFF. (2) Disconnect the electrical connector. (3) Remove the four screws securing the sensor in the ACM P3 air duct. (4) Remove the sensor and gasket. (5) Retain the gasket.

B. Installation (1) Install gasket retained from removal procedures on temperature sensor. (2) Set the sensor in the ACM P3 air duct and secure it with four screws (3) Connect and safety the electrical connector. (4) Replace all panels.

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ENVIRONMENTAL SYSTEMS DISTRIBUTION DESCRIPTION AND OPERATION

21-20-00 00

1. GENERAL Air produced to cool and heat the airplane cabin is distributed through insulated ducting to outlets located on the cabin floor, adjacent to the armrests, and in the flight compartment. Distribution ducting for the air cycle system is independent of the ducting for the vapor cycle cooling system (Ref. 21-10-00).

A. Vapor Cycle Air Distribution Cool air produced by the vapor cycle cooling system is distributed by the vent blowers through insulated ducts to the passenger compartment and the flight compartment (Ref. Figure 1). The eyeball outlets can be adjusted to control the direction and amount of airflow and are located above the passenger compartment armrests and in the flight compartment overhead panel. The blowers for the vapor cycle cooling system are located immediately forward of each evaporator and will operate in high speed or low speed when the cabin temperature mode switch is set to each position (Ref. Figure 2). When the blower control switch is set to AUTO, the blowers will operate in the LO speed when the mode switch is set to each position except OFF. When the blower control switch is set to HI or LO, the blowers will operate in the selected speed regardless of the position of the mode switch. The low speed relays are energized through the VENT BLOWER CONTROL circuit breaker, the mode switch, and the blower control switch. The circuit breaker is located on the circuit breaker panel. When the mode switch is set to OFF, the low speed relays are energized through the LO position of the blower control switch. The low speed relays can be energized through the mode switch when the blower control is set to AUTO. When the vent blower control switch is set to LO or AUTO, only the low speed relays are energized and power is supplied through the 30-ampere limiters to the LO wire of each vent blower. The vent blower control switch is located on the copilot's inboard subpanel. The relays are located adjacent to the blowers. A thermal protection device in the low speed circuit (LO terminal) of the ventilation blower (modified by Kit No. 630-203-1 or Kit No. BC80A-901-3) provides thermal protection to the vent blower assembly. The thermal protection device will interrupt power to the low speed circuit resistor at a specified temperature setting. This protects the vent blower from overheating due to shorted brush lead-wires or deteriorated armature bearings which can cause smoke or fumes to enter into the vent system. Refer to Table 1, VENTILATION BLOWER ASSEMBLY, for new part number after modification kit is installed. The vent blower control switch is located on the copilot’s inboard subpanel. The relays are located adjacent to the blowers (Ref. Figure 2). The high speed relays are energized through the blower control switch when it is set to the HI position. When the high speed relays are energized, power is supplied through the closed contacts of the high speed relays and the low speed relays to the HI wire of each vent blower. Flapper valves installed on each blower assembly provide air to the evaporators to help prevent moisture from freezing on the evaporators when a large percentage of the eyeball outlets are closed.

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Table 1 Ventilation Blower Assembly

Part Number

New Part Number upon incorporation of Mod Kit

Effectivity UA-1 and After UB-1 and After UC-61 thru UC-126

Description Cockpit or Cabin Ventilation Blower Assembly

114-380028-31 or BC80A-12

UC-127 and After

Cockpit or Cabin Ventilation Blower Assembly

114-380028-91 or BC80A-92

114-380028-51 or EM630-23

UA-1 and After UB-1 and After UC-127 and After

Cockpit or Cabin Ventilation Blower Assembly

114-380028-111 or EM630-33

114-380028-71 or BC80A-32

UA-1 and After UB-1 and After UC-127 and After

Cockpit or Cabin Ventilation Blower Assembly

114-380028-91 or BC80A-92

114-380028-11 or EM6303

114-380028-111 or EM630-33

1

Part number of Hawker Beechcraft Corporation, P.O. Box 85, Wichita, KS, 67201-0085. Part number of Advanced Industries Inc., 4550 Southeast Blvd., Wichita, KS, 67210. 3 Part number of Electromech Technologies, 2600 S. Custer, Wichita, KS, 67217. 2

B. Air Cycle System Distribution Conditioned bleed air, produced by the air cycle machine cooling system, is distributed through insulated ducting to floor outlets in the passenger compartment and the flight compartment. The floor outlets include protective coverings and are located adjacent to the sidewalls on the floor of the passenger compartment and the flight compartment. A flapper valve is installed in the ducting and is controlled by a push-pull cable located on the copilot's inboard subpanel. The flapper valve is located at the junction of the ejector tube and the main distribution duct and is in line with the third cabin window. The control cable can be pulled to divert most of the conditioned bleed air forward to the flight compartment (Ref. Figure 1). Ambient air can be supplied to the cabin through the air cycle system ducting when the airplane is depressurized. A manually controlled valve located in the nose ram air duct can be opened to allow air to enter the airplane through the ram air door solenoid valve and the manual valve when the cabin pressure control switch is set to DUMP. The control for the manual shutoff valve is a push-pull cable mounted under the copilot's inboard subpanel.

C. Defrost and Flight Compartment Heat Conditioned bleed air, produced in the air cycle machine, is distributed through insulated ducting to outlets located forward of the instrument panel for heating and windshield defrost. Push-pull cables allow the pilot and copilot to control butterfly valves that control the amount of conditioned bleed air flow from the outlets. The pilot's control is located on the pilot's outboard subpanel and the copilot's control is located on the copilot's outboard subpanel. Conditioned bleed air for windshield defrost is controlled by a push-pull cable control located on the pilot's inboard subpanel. The control cable is pulled to allow conditioned bleed air flow through a butterfly valve to the windshield.

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Figure 1 Environmental System Air Distribution

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This Page Intentionally Left Blank

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Figure 2 Vent Blower Electrical Schematic

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ENVIRONMENTAL SYSTEMS DISTRIBUTION TROUBLESHOOTING

100100

1. DISTRIBUTION Improper rigging of the flapper valve in the conditioned bleed air ducting may cause improper operation of the air cycle system. Refer to the procedures listed in this section to rig the flapper valve. Figure 101 outlines troubleshooting of the vent blower control.

Figure 101 Troubleshooting - Vent Blower Control

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ENVIRONMENTAL SYSTEMS DISTRIBUTION MAINTENANCE PRACTICES

200200

1. PROCEDURES A. Flapper Valve Rigging, Conditioned Bleed Air When the flapper valve in the main ducting is closed to decrease the airflow to the aft outlets, the valve should not shutoff the flow of air to the aft outlets completely (Ref. Figure 201). If the flapper valve or control cable is removed or adjusted, the flapper assembly should be rigged to close with approximately 0.3-inch clearance of the main duct when the control cable on the copilot's inboard subpanel is pulled to its stop.

B. Flapper Valve Rigging, Nose Ram Air The valve installed in the nose ram air inlet duct closes the inlet completely. The valve should be rigged completely closed when the VENT manual control is pushed in to its stop.

C. Butterfly Valve Rigging, Heat and Defrost The butterfly valves in the pilot's heat outlet, the copilot's heat outlet, and the defrost duct, shut off the flow of air to the outlets (Ref. Figure 202). Each valve should be rigged completely closed when the applicable cable is pushed in to its stop.

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Figure 201 Flapper Valve Rigging - Condition Bleed Air

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Figure 202 Butterfly Valve Rigging - Heat and Defrost

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ENVIRONMENTAL SYSTEMS VENT BLOWER MAINTENANCE PRACTICES

21-20-01 200200

1. PROCEDURES A. Removal (1) Remove the center aisle carpet (Ref. Chapter 25-20-01, CARPET REMOVAL and INSTALLATION). (2) Remove the floorboards covering the vent blower (Ref. Chapter 6-50-00, FLOORBOARD ACCESS PANELS). (3) Disconnect the electrical leads from the terminals on the blower. (4) Remove the attaching parts securing the blower to the main distribution ducts. (5) Remove the attaching parts securing the blower to the mounting bracket. (6) Loosen the nut on the clamp securing the blower to the evaporator and remove the blower.

B. Installation (1) Install the blower on the main distribution duct and secure with the attaching parts. (2) Install the blower on its mounting bracket and secure with the attaching parts. (3) Tighten the nut on the clamp to secure the blower to the evaporator. (4) Connect the electrical leads to the blower terminals. (5) Install the carpet (Ref. Chapter 25-20-01, CARPET REMOVAL and INSTALLATION). (6) Install the floorboards in the center aisle (Ref. Chapter 6-50-00, FLOORBOARD ACCESS PANELS).

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ENVIRONMENTAL SYSTEMS AIR OUTLET MAINTENANCE PRACTICES

21-20-02 200200

1. PROCEDURES A. Removal The air outlet may be removed as follows: (1) Fabricate a tool for the air outlet removal and installation as follows: (a) Use a pipe with a 2.80 in. outer diameter and 0.20 in. thickness. (b) Drill two No. 54 holes in line in the pipe wall, 2.70 in. apart. (c) Insert a 0.06-in. diameter pin into each hole. (d) Cut the pin so that only 0.05 in. of the pin protrudes above the pipe wall. (2) Using the fabricated tool, remove the air outlet by inserting the pins of the tool into the two holes in the air outlet and unscrewing it counterclockwise.

B. Installation (1) Lubricate the threads on the side of the air outlet with lubricant, petrolatum (90, Table 1, Chapter 91-00-00). (2) Using the fabricated tool under AIR OUTLET REMOVAL procedure in this section, install the air outlet by inserting the pins of the tool into the two holes of the air outlet and screwing it in clockwise. NOTE: If the air outlet is installed cross threaded, damage may result to the outlet requiring a new air outlet.

2. AIR OUTLET ACTUATOR A. Removal (1) Remove the air outlet as described under AIR OUTLET REMOVAL procedure in this section. (2) Remove and retain the spring washer at the back of the air outlet to release the actuator. (3) Note the position of the actuator and remove it.

B. Installation (1) Set the actuator protruding out of the back of the air outlet. (2) Secure the actuator in place with spring washer. (3) Install the air outlet as instructed under AIR OUTLET INSTALLATION procedure in this section.

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ENVIRONMENTAL SYSTEMS PRESSURIZATION CONTROL DESCRIPTION AND OPERATION

21-30-00 00

1. GENERAL Pressurization of the airplane cabin is controlled by the cabin pressurization controller, a pneumatic relay, solenoid valves, and outflow valves. The pressure controller, pneumatic relay, preset and dump solenoids, volume tank and orifices are installed within the pedestal in the flight compartment. The cabin pressure control switch, the climb rate indicator, the cabin differential pressure indicator, and the pressure controller selectors are located on the pedestal. Orifices installed in the vacuum and vacuum control lines, in the pressure controller, and in the pneumatic relay help limit the amount of change in cabin pressurization. Filters are incorporated in the cabin air ports on the pressure controller, and the pneumatic relay. Outflow valves sense atmospheric pressure through plumbing that connects the valves to static ports located on the LH side and the RH side of the fuselage aft of the aft cabin door. These static ports are separate from the static ports connected to the instruments mentioned in Chapter 34-10-00. The outflow valves are installed on dump vent ducts located on the aft pressure bulkhead. On serials UC-1 and after, a low point drain is located at F.S. 163, on the right side, under the cabin flooring. The drain is used to relieve any water that may collect in the control lines between the dump valve solenoid and the outflow valves (Ref. Figures 1 and 2). Vacuum pressure for the pressure controller is controlled by a vacuum regulator located immediately aft of the forward pressure bulkhead. This regulator is used only in the pressurization system and maintains a vacuum pressure between 3.75 and 4.75 in. Hg. The regulator is connected to the vacuum manifold. For further information on the vacuum manifold (Ref. Chapter 37-00-00). The inflow of air for cabin pressurization is environmental bleed air from the LH and RH engines. The inflow pressure is controlled by bleed air regulator/shutoff valves installed in the bleed air lines connected to each engine compressor. For information on the valves and the control of bleed air from each engine (Ref. 21-10-00). When the cabin pressure control switch is set to PRESS (pressurize) prior to takeoff, the energized preset solenoid shuts off the regulating vacuum to the pressurization controller (Ref. Figures 1 and 2). The energized pressure dump solenoid opens, making vacuum control available to the outflow valves. The vacuum control is amplified by the pneumatic relay and allows the diaphragm in the outflow valves to open so that the cabin pressure altitude and the rate of change of cabin pressure can be set on the pressure controller. In flight (when the squat switch is opened), the preset solenoid (normally open) is de-energized and allows vacuum to be applied to the controller. The dump solenoid valve (normally closed) de-energizes and shuts off direct vacuum to the outflow valves. The controller then controls the outflow valves through the pneumatic relay. The pneumatic relay amplifies the vacuum control from the controller to allow the outflow valves to be controlled in tandem. The function of each valve is identical. Each outflow valve contains a relief valve that relieves cabin pressure according to the rate-of-change and altitude set on the pressure controller unless the valve relieves pressure when the cabin-to-atmosphere pressure differential exceeds a maximum differential of 4.8 psi ± 0.1. The volume tank provides a pressure memory to allow the controller an accurate rate-of-change control. If atmospheric pressure exceeds cabin pressure, a negative pressure relief diaphragm in each outflow valve opens the valves to allow atmospheric pressure to relieve cabin negative pressure. The cabin pressure control switch can be set to the TEST position to de-energize the preset and dump solenoids and allow the pressure control system to function as though the airplane were in flight.

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Ambient air can enter the cabin when the cabin pressure differential is minimal and the ram air door solenoid latch is opened (de-energized) by setting the cabin pressure control switch to DUMP. If the VENT manual control is pulled to open the butterfly in the ram air inlet, ambient air is allowed to flow into the ram air inlet, through the solenoid controlled door on the forward pressure bulkhead, and into the conditioned bleed air duct in the flight compartment. When the cabin pressure control switch is set to the PRESS position, the ram air door solenoid latch is energized to the closed position and magnetically holds the ram air door closed. When the airplane lands (the LH squat switch is in the down position), the preset solenoid and the dump solenoid are energized and the cabin is allowed to depressurize through the opened outflow valves. Cabin pressure altitude and the cabin-to-atmosphere pressure differential are indicated on the differential pressure indicator. The pressure differential is expressed in psig and the pressure altitude is expressed in thousands of feet. The climb-rate-indicator allows monitoring of the rate of change of cabin pressurization. Cabin pressure altitude and the rate of change of cabin pressure altitude can be controlled by using the appropriate adjustment knob located on the cabin pressure controller. However, the differential pressure indicator and the climb rate indicator are not directly connected to the controller. Turning the rate-of-change selector clockwise will increase the rate of change of cabin pressurization and turning the selector counterclockwise will decrease the rate of change. If cabin pressure altitude exceeds 12,500 +0 -500 feet, the cabin altitude warning pressure switch closes, and the warning annunciator light labeled CABIN ALT will illuminate. The cabin altitude warning circuit is separate from the pressurization control circuit. The pressure switch in the altitude warning circuit is installed on the fuselage electrical equipment panel located beneath the cabin floorboard in line with the aft edge of the cabin forward door (Ref. Chapter 39-20-00).

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Figure 1 Pressurization Control Mechanical Schematic

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Figure 2 Pressurization Control Electrical Schematic

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ENVIRONMENTAL SYSTEMS PRESSURIZATION CONTROL TROUBLESHOOTING

100100

1. PROCEDURES This troubleshooting section contains a simple AIRBORNE FUNCTIONAL TEST to determine that the cabin is pressurizing properly. The AIRBORNE SYSTEM CHECKS is also included to check that the entire system is functioning properly during flight. In addition, there are two methods to check the CABIN LEAK RATE. The FLIGHTLINE GROUND TEST is a simplified version which requires running an engine and using the airplane’s pressurization controls. The HANGAR GROUND TEST does not require starting the engines, uses test equipment to monitor cabin pressure, and includes three configurations for attaching test equipment for pressurizing the cabin. For information on troubleshooting (Ref. Figure 101). Figure 101 is used in conjunction with the PRESSURIZATION CHECKS WORKSHEET (Ref. Table 101).

A. Airborne Functional Test (1) Prior to takeoff, start at least one engine and set N1 to 85 percent. Set the applicable bleed air switch to OPEN. Set the controller at an altitude approximately 1000 feet below field pressure altitude or to sea level on the dial, whichever is the higher altitude, and hold the control switch in the TEST position for at least 45 seconds. The climb rate indicator and the cabin pressure altitude should show a decrease. If the controller is set above field pressure altitude, the climb rate indicator will drop, then stabilize at zero when the control switch is held in the TEST position. Pressurization will not occur unless the controller is set below field pressure altitude. (2) After performing the procedures outlined in Step (1), operate the system with the control switch in the PRESS position. Prior to takeoff, set the desired cabin pressure altitude and the rate-of-climb on the pressure controller. As the airplane leaves the ground, the cabin will pressurize at the rate set on the cabin rate-of-climb selector until the preset cabin altitude is reached, or until the maximum pressure differential is reached. The cabin will then maintain pressurization according to the maximum differential pressure of the outflow valves (4.8 psi ± 0.1). The maximum differential pressure for each valve is identical. Cabin altitude will continue to climb at a slower rate than the airplane until cruise altitude is reached. Cabin pressure will then stabilize and maintain maximum differential control.

B. Airborne System Checks Troubleshooting of the pressurization system is best carried out in two phases: data collection and data analysis. Data is collected as a result of a series of operational checks and recorded on the PRESSURIZATION CHECKS WORKSHEETS. The data is then applied to the troubleshooting flowchart where certain fault possibilities and appropriate corrective actions are suggested (Ref. Table 101).

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Data Analysis for Pressurization Check Procedures In PRESTART PROCEDURE Step (2) Were both outflow/safety valves closed?

NO Replace open valve.

NO Is dump solenoid opening?

NO

YES

NO

Did either valve open?

Did both valves open during PREFLIGHT PROCEDURE Step (2)

YES YES

Is dump solenoid receiving current through squat switch? NO

YES

Replace dump solenoid.

Correct open circuit.

NO

Replace closed valve.

YES Are valves receiving vacuum?

Correct vacuum leaks (1.).

NO Replace preset solenoid

Is preset solenoid opening?

NO

YES

Replace faulty valve. Is dump solenoid stuck open? YES

NO

Did outflow valves close during PREFLIGHT PROCEDURE Step (7)

Replace dump solenoid.

YES Is there vacuum at test port #1?

NO Replace pneumatic relay.

YES

YES

Fault in controller (2.).

Did outflow valves close during PREFLIGHT PROCEDURE Step (7)

NO

Fault in controller.

NO

YES

Magnetic latch malfunctioning.

NO

Is ram air door closed? YES

Cabin leak rate too high (3.).

Did cabin pressurize during PREFLIGHT PROCEDURE Step (8) YES

Is cabin rate adjustable during PREFLIGHT PROCEDURE Step (9) YES NO

NO Damaged controller.

Squat switch not opening preset and dump solenoids circuits.

YES NO

Seal cabin leaks (3.).

Did airplane pressurize during CLIMB Step (2)

Did cabin leak rate check good during CABIN LEAK CHECK Step (1) YES GO TO NEXT PAGE. UC21B 071359AB.AI

Figure 101 (Sheet 1 of 2) Pressurization Control Troubleshooting Flowchart

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Data Analysis for Pressurization Check Procedures NO

Are SINGLE SOURCE CHECKS good in Steps (2) and (5)?

YES

Troubleshoot bleed air control. NO

Are MAXIMUM DIFFERENTIAL CHECKS good in Steps (3) and (5)?

Replace out of tolerance valve (5.).

YES NO

NO

Controller out of adjustment.

Are rates unbalanced (4.)?

Are CONTROLLER CHECK minimum rates within tolerance in Steps (1) and (4)?

YES YES

Correct vacuum leaks (1. or 2.). NO

Did cabin pressure altitude fall within tolerance during CONTROLLER CHECK Steps (3) and (6)?

Controller out of adjustment.

YES

System is operating as designed.

1. Vacuum leaks may occur in any of the lines supplying vacuum to the outflow valve or may occur in the pneumatic relay or the outflow valve head. Refer to the vendor's maintenance instructions (Report no. 4-356, paragraphs 7-23) in the 1900 Airliner Series Component Maintenance Manual for vacuum leak check procedures. 2. Excessive vacuum leaks may occur in any of the lines supplying vacuum to the pneumatic relay or may occur internally in the controller. Refer to the vendor's maintenance instructions (Report no. 4-356, paragraphs 7-22) in the 1900 Airliner Series Component Maintenance Manual for vacuum leak check procedures. 3. Perform the checks under CABIN LEAK RATE FLIGHTLINE GROUND TEST or CABIN LEAK RATE HANGAR GROUND TEST found in this subchapter and seal excessive leaks until the cabin leak rate is reduced to an acceptable level. 4. While determining the minimum up and down rates of the controller, should the minimum down rate be out of tolerance and high, and the minimum up rate be low (either in or out of tolerance), the minimum rates are unbalanced. 5. These outflow valves are not field adjustable. Any attempt to adjust a valve will void the warranty. UC21B 071360AB.AI

Figure 101 (Sheet 2 of 2) Pressurization Control Troubleshooting Flowchart

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL The pressurization checks are performed in four stages: prestart, preflight, climb and cruise. Occasionally, at certain points during the checks, the procedure suggests terminating the entire check procedure until the currently revealed fault is corrected. Beyond these points, any further data collection would most likely yield unreliable results. It is suggested that no attempt to analyze data be made during the data collection phase of the troubleshooting. It is also suggested that, when performing the pressurization checks and analyzing the data, the technician not deviate from the logical progression that has been established by these procedures. The PRESSURIZATION CHECKS WORKSHEET may be copied by maintenance personnel; such copying will not constitute a violation of Hawker Beechcraft Corporation copyrights. Table 101 PRESSURIZATION CHECKS WORKSHEET DATA COLLECTION PRESTART PROCEDURE From Step (2): Were both outflow valves closed? Yes __________ No __________ PREFLIGHT PROCEDURE From Step (2): Did both outflow valves open? Yes __________ No __________ From Step (3): Did both outflow valves close? Yes __________ No __________ From Step (4): FPA from copilot’s altimeter __________ CPA from test altimeter __________ From Step (7): Did outflow valve begin closing? Yes __________ No __________ From Step (8): Did cabin rate indicate a descent within 45 seconds? Yes __________ No __________ From Step (9): Can cabin rate be changed? Yes __________ No __________ From Step (10): Cabin Pressure Altitude (CPA) from test altimeter __________ CPA from cabin altimeter __________ CPA Set on Controller __________ From Step (12): Did cabin indicate a climb? Yes __________ No __________ Can cabin rate be changed? Yes __________ No __________ Did CPA stabilize at Field Pressure Altitude (FPA)? Yes __________ No __________ From Step (13): Cabin controller setting __________ (FPA + 2000 ft.) DURING CLIMB From Step (1): Did cabin pressurize during climb? Yes __________ No __________ From Step (2): Cruising pressure altitude __________ Outside air temperature __________ CPA from test altimeter __________ Tol: 850 to 1250 ft. CABIN LEAK CHECK From Step (1): Cabin leak climb rate __________ Tol: 2400 fpm or less

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Table 101 PRESSURIZATION CHECKS WORKSHEET (Continued) SINGLE SOURCE CHECK From Step (2): RH Engine N1 __________ Tol: Not more than 85% N1 From Step (5): LH Engine N1 _________ Tol: Not more than 85% N1 MAXIMUM DIFFERENTIAL CHECK From Step (1): Altitude from copilot’s altimeter __________ (Same as DURING CLIMB Step (2)) WARNING During the following differential pressure checks, carefully monitor the cabin differential pressure gage and the test altimeter. A cabin altitude less than 850 ft. may exceed the maximum cabin pressure differential and will necessitate terminating this check. Place the pressurization control switch in DUMP immediately and reconnect the atmospheric reference lines to the disabled valve. From Step (3): CPA from test altimeter __________ (LH outflow valve disabled) Differential pressure gage reading __________ Tol: 4.7 to 4.9psi From Step (5): CPA from test altimeter __________ (RH outflow valve disabled) Differential pressure gage reading __________ Tol: 4.7 to 4.9psi CONTROLLER CHECK From Step (1): Minimum up-rate __________ (Controller set at 8000 ft.) (less than 300 fpm) From Step (2): Maximum up rate __________ (greater than 1500 fpm) From Step (3): CPA from test altimeter __________ Tol: 8000 ft. ± 400ft From Step (4): Minimum down-rate __________ (Controller set at 2000 ft.) (less than 300 fpm) From Step (5): Maximum down-rate __________ (greater than 1500 fpm) From Step (6): CPA from test altimeter __________ Tol: 2000 ft. ± 400ft

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ENVIRONMENTAL SYSTEMS CABIN PRESSURIZATION SYSTEM MAINTENANCE PRACTICES 1. PRESSURIZATION CHECK PROCEDURES A. Equipment Required (1) Test Altimeter (Certified Accurate) (2) Pressurization Checks Worksheet

B. Prestart Procedure (1) Remove the upper aft upholstery panel covering the outflow valves. Check all components and lines for security of attachment. (2) Notice the positions of the outflow valves: both valves should be closed. Replace either valve that is not closed.

C. Preflight Procedure (1) Start the engines according to the procedure outlined in the appropriate Pilot’s Operating Handbook. (2) Turn on the instrument air only and time the outflow valves’ opening cycles: it may take up to 30 seconds for the valves to open fully. Replace either valve that does not open. (3) Turn the instrument air off and time the closing of the valves: closing time for these valves should be between 15 and 30 seconds. Replace either valve for excessive closing time. NOTE: An outflow valve which operates too quickly or too slowly may cause “bumps” when transitional between pressurized and non pressurized operation. (4) Set the copilot’s altimeter and the test altimeter to 29.92 in. Hg. and record the field pressure altitude (FPA) from the copilot’s altimeter and the cabin pressure altitude (CPA) from the test altimeter. If a discrepancy exists between the two altimeter readings, set the copilot’s altimeter to the same altitude setting as the test altimeter. (5) Turn on both environmental bleed air switches and increase engine rpm to high idle. (6) Set the controller cabin altitude dial to 1000 feet below FPA or to sea level on the dial, whichever is the higher altitude. NOTE: At low elevation airports during high barometric pressure days, when the FPA is below sea level, this check cannot be performed and should be no cause to condemn any component in the pressurization system. (7) While monitoring the outflow valves, hold the pressurization control switch in the TEST position. (8) The cabin rate-of-change indicator should indicate a descent within 45 seconds. If the cabin will not pressurize at this point, discontinue these checks; the cabin leak rate is too great or the cabin ram air door may not be closed.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (9) When the cabin rate of change stabilizes, turn the rate knob to minimum (fully CCW) and notice that the cabin rate of change decreases. Replace a controller that shows no change in rate response. (10) When the cabin altitude stabilizes (rate-of-change indicator reads zero), record the cabin altitude from the test altimeter and the airplanes cabin altitude indicator; record cabin altitude setting on the pressurization controller and the reading from the cabin differential pressure gage. (11) While still holding the test switch in TEST, turn the cabin altitude select dial to 500 feet above FPA. (12) The cabin rate-of-change indicator should indicate a rate of climb. Turn the rate-of-change knob to the maximum rate and notice an increase in the rate of climb: the cabin should stabilize at FPA. Replace a controller that shows no change in rate response. (13) Set cabin altitude on the controller to 2000 feet above FPA and the rate knob to the 12 o’clock position. (14) Turn off both environmental bleed air sources according to the guidelines listed in the Pilot’s Operating Handbook.

D. During Climb (1) Turn both environmental bleed air sources on. NOTE: Should the cabin fail to begin pressurizing after passing through the cabin altitude selected on the controller, discontinue these checks: the left landing gear squat switch is not opening the preset and dump solenoids circuit. (2) Select SL on the controller and climb the airplane to 12,000 feet pressure altitude according to the copilot’s altimeter: record this altitude on the worksheet. Cabin altitude should stabilize between 850 and 1250 feet.

E. Cabin Leak Check (1) Turn off RH and LH environmental bleed air and record the cabin rate of climb when the rate stabilizes: the cabin rate of climb will initially indicate a rate above the actual leak rate, stabilize, then begin to decrease gradually. Record the cabin climb rate at the point where the cabin rate begins to decrease gradually. NOTE: If the cabin leak rate exceeds 2400 fpm, terminate these checks and correct the cabin leak rate. (2) Turn on both bleed air sources and permit the cabin altitude to stabilize once again.

F. Single Source Check (1) Turn the left environmental bleed air off. (2) Reduce power towards high idle on the RH engine and record the engine N1 at the point where cabin altitude begins to climb. Do not reduce power below high idle. (3) Restore power to the RH engine and turn the left bleed air on. (4) Permit the cabin altitude to stabilize again and turn the right environmental bleed air off.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (5) Reduce the power towards high idle on the LH engine and record the engine N1 at the point where cabin altitude begins to climb. Do not reduce the power below high idle. (6) Restore power to the LH engine and turn the RH bleed air on.

G. Maximum Differential Check (1) Before beginning the next check ensure that the airplane is at the pressure altitude selected earlier, 12,000 feet. WARNING: Carefully monitor the cabin pressure differential during the following differential pressure checks to ensure that cabin pressure does not exceed the maximum allowable limit. Should the altitude on the test altimeter descend below 850 feet, discontinue this check immediately. Placing the pressurization control switch in DUMP will allow cabin pressure to return to normal limits. (2) Disconnect the atmospheric reference lines from the LH outflow valve and seal the reference lines to prevent loss of cabin pressure through the reference ports: this evaluates the maximum differential pressure function of the RH outflow valve. (3) Record the cabin altitude from the test altimeter (Tol: 880 to 1265 ft), cabin differential from the airplanes differential pressure gage (Tol: 4.7 to 4.9 psi), and cabin altitude from the airplanes cabin altimeter. CAUTION: An outflow valve which fails to maintain cabin altitude within acceptable tolerance must be replaced. These valves are not field adjustable and any attempt to adjust a valve will void the warranty. (4) Reconnect the atmospheric reference lines to the LH outflow valve and remove the atmospheric reference lines from the opposite valve, sealing the reference lines as before: this will evaluate the maximum differential pressure function of the LH valve. (5) Record cabin altitude from the test altimeter (Tol: 880 to 1265 ft), cabin differential from the airplane’s differential pressure gage (Tol: 4.7 to 4.9 psi), and cabin altitude from the airplane’s cabin altimeter; then reconnect the atmospheric reference lines to the outflow valve. (6) Restore power to the LH engine and turn the RH bleed air on.

H. Controller Check (1) Select MIN rate and 8000 feet cabin altitude on the controller, allow the rate of climb to stabilize and time the amount of change in cabin pressure altitude, according to the test altimeter, for sixty seconds. Record this minimum up-rate on the worksheet. (2) Select MAX rate, allow the rate of climb to stabilize and time the amount of change in cabin pressure altitude for sixty seconds. Record this maximum up-rate on the worksheet. (3) When the cabin altitude stabilizes and the rate-of-change indicator has returned to zero, record the cabin pressure altitude from the test altimeter. (4) Select MIN rate and 2000 feet cabin altitude on the controller, allow the rate of descent to stabilize and time the amount of change in cabin pressure altitude for sixty seconds. Record this minimum down-rate on the worksheet.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (5) Select MAX rate, allow the rate of descent to stabilize and time the amount of change in cabin pressure altitude for sixty seconds. Record this maximum down-rate on the worksheet. (6) When the cabin altitude stabilizes and the rate-of-change indicator has returned to zero, record the cabin pressure altitude from the test altimeter. (7) Refer to the troubleshooting Charts in this section for a complete analysis of the data collected during these pressurization check procedures.

2. CABIN LEAK RATE - MAINTENANCE PRACTICES A. Cabin Leak Rate Flightline Ground Test This test uses engine bleed air to pressurize the cabin, bypasses the airplane’s pressurization controller to guarantee an effective climb rate, and uses the airplane’s instrumentation to monitor the differential pressure and cabin climb rate. A Chart is also provided to relate cabin climb rate to maximum allowable leakage (Ref. Figure 201). (1) Station one person outside the airplane during the test to prevent anyone from approaching the airplane or trying to open a door or access panel. WARNING: DO NOT under any circumstances, open the cabin door, cargo door, cockpit windows, escape hatches, or access panels while the cabin has ANY pressure applied during this test. Always have at least two people inside the airplane during this test. Workers inside the airplane during this test must be free of obesity, heart disorders, respiratory or ear infections, and must be emotionally stable. (2) Attach a sign on the outside of the cabin door that reads: DANGER - AIRPLANE IN PRESSURE TEST. (3) Obtain access to the orange 3/8-inch test cap on the cabin pressurization control plumbing under the pedestal to bypass the pressurization controller and allow the pressurization system to establish a climb rate even though the airplane may be sitting at sea level. (4) Remove the cap and install a suitable adjustable needle valve. Start with the valve shut. (5) Set the altimeter to 29.92 in. Hg. and note the Field Pressure Altitude (FPA). Use the FPA and the graph in Figure 201 to establish the maximum climb rate equivalent to maximum allowable leakage. Even though the airplane will not leave the ground, the climb rate will decrease as the cabin is pressurized and will increase as the cabin leaks. It is the positive climb rate that can be used as a measure of maximum allowable leakage when the cabin is pressurized to 4.3 psi differential. In the example given on the graph in Figure 201, the FPA was 3800 feet, therefore, the airplane instrumentation “feels” like it is at 3800 feet. The 3800 feet value is converted to the cabin climb rate of 2210 ft/min. When the cabin is pressurized to a 4.3 psi differential, the climb rate should not exceed 2210 ft/min; otherwise, the cabin is leaking more than is acceptable. WARNING: DO NOT allow the cabin climb rate to exceed 1000 ft/min as discomfort will be imposed on the pilots or maintenance personnel. (6) Set the CABIN CONTROLLER rate selector knob to full clockwise position.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (7) Secure all windows, doors, escape hatches, and access panels that are part of the pressure vessel. CAUTION: DO NOT exceed engine operating limits. (8) Start either engine and set N1 to 85%. (9) Set the applicable bleed air switch to OPEN. (10) Set the MODE CONTROL switch to AUTO. (11) Set the CABIN PRESSURE switch on the pedestal to TEST. (12) Slowly open the needle valve on the plumbing under the pedestal and use it to control the climb rate and limit cabin pressure. (13) Monitor the CABIN CLIMB indicator and establish a climb rate at or below 1000 ft/min. (14) Monitor the CABIN ALT/DIFF PRESS indicator to establish a cabin differential of 4.3 psi. NOTE: It may be necessary to increase engine N1 after the test has begun in order to reach 4.3 psi cabin differential. Adjust the needle valve after any engine N1 change to maintain the rate of climb at or below 1000 ft/min. (15) When 4.3 psi cabin differential is reached, set the bleed air switches to INST & ENVIR OFF. Observe the indicated cabin rate of climb: the value will initially indicate a higher rate, stabilize, then begin to decrease gradually. Record the cabin climb rate at the point where it begins to decrease gradually. If the climb rate exceeds the limit determined in Step (5), the maximum allowable leakage is excessive and should be corrected.

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Figure 201 Equivalent Maximum Allowable Leakage

B. Maximum Differential Check A maximum differential check may be performed at this time if the operator has any reason to question the operation of the outflow valves. WARNING: Carefully monitor the cabin differential pressure during the following checks to ensure that it does not exceed the maximum allowable limit of 4.9 psi. If the limit is exceeded, relieve cabin pressure by closing the needle valve or by shutting off the bleed air valves. (1) Close the needle valve installed in Section 2. A. Step (4). Set the bleed air switches to OPEN. Slowly open the needle valve and establish a climb rate at or below 1000 ft/min. The outflow valves should begin to limit maximum differential as it approaches 4.7 psi.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (2) Remove the upper cover on the aft pressure bulkhead to gain access to the outflow valve pneumatic plumbing. (3) Disconnect and seal the atmospheric reference lines from the left outflow valve to evaluate the maximum differential function of the right outflow valve. (4) Record the cabin differential from the CABIN ALT/DIFF PRESS indicator. The operating tolerance is 4.5 to 4.9 psi. CAUTION: An outflow valve which fails to maintain cabin differential within the specified tolerance must be replaced. These valves are not adjustable. Any attempt to adjust a valve will void the warranty. (5) Clear and connect the atmospheric reference lines of the left outflow valve. (6) Disconnect and seal the atmospheric reference lines from the right outflow valve to evaluate the maximum differential function of the left outflow valve. (7) Record the cabin differential from the CABIN ALT/DIFF PRESS gage. The operating tolerance is 4.7 to 4.9 psi. (8) Clear and connect the atmospheric reference lines of the right outflow valve. (9) Replace the upper cover on the aft pressure bulkhead. (10) Gradually decrease cabin pressure by closing the needle valve or shutting off the bleed air valves. (11) Set the CABIN PRESSURE switch to DUMP. (12) Watch the CABIN ALT/DIFF PRESS indicator and wait for it to read zero. WARNING: DO NOT under any circumstances, open the cabin door, cargo door, cockpit windows, escape hatches, or any access panels while the cabin has ANY pressure applied during this test. (13) Open the storm window in the cockpit to verify that the cabin pressure is completely bled off. (14) Shut down the engines. (15) Remove the needle valve installed at the test port. Replace and tighten the test port cap.

C. Cabin Leak Rate Ground Test Equipment Test units used to ground test the cabin for pressurization leaks must include air inlet filters and a pressure relief valve to prevent damage to the airplane. The typical test unit used to check the cabin leak rate will consist of an air blower, a dry filter, pressure relief valve, aircraft pressure indicator, climb rate indicator, air flow indicator, unit pressure indicator, unit air temperature indicator, unit air control/ dump valve, shop air regulator, and a shop air pressure indicator. The method for interfacing the test equipment to the airplane will vary. Some test equipment splices into the bleed air plumbing in the nacelles with a “Y” pressure hose or an optional single 2-inch pressure hose attached to an adapter plate in the fuselage belly.

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The test units and adapter plate available for the CABIN LEAK RATE HANGAR GROUND TEST on Model 1900 Airliner Series airplanes that include the safety features and instrumentation described above are: (1) Cabin Pressurization Unit: TRONAIR, South Eber Road, Holland, Ohio, 43528; Phone (413) 866-6301. (a) Model 15-7600-1000 (60 Hz) (b) Model 15-7602-1000 (50Hz) (c) Adapter K-1285 (Ref. Figure 203) NOTE: When it is desirable to pressurize the cabin with the single 2-inch pressure hose and the adapter plate required is not supplied, the following adapter plate may be obtained from Hawker Beechcraft Corporation: (d) Cabin Pressurization Adapter Plate TK1794-5/939 (Ref. Figure 203) NOTE: The safety net required to secure the cabin door and the safety straps required to secure the cargo door may also be purchased from Hawker Beechcraft Corporation: (e) Safety Net - Cabin Door - P/N 97-00000/939-1 (f) Safety Straps - Cargo Door - P/N 97-00000/939-2

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Figure 202 Bleed Air Plumbing Access

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Figure 203 Cabin Pressurization Adapter Plates

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D. Cabin Leak Rate Hangar Ground Test Carefully review the technical data provided with the test unit prior to initiating the test procedure. (1) Connect ground power and set the MASTER SWITCH to ON. (2) Disconnect the vacuum (relay) line (9) from the T-fitting (12) that connects outflow valves (10). Cap the vacuum (relay) line (9) while leaving T-fitting (12) uncapped (Ref. Figure 205). (3) Set both BLEED AIR VALVES switches to INST & ENVIR OFF to close the instrument pneumatic shutoff valves. (4) Three configurations are given to cover the different requirements for preparing the airplane for the CABIN LEAK RATE HANGAR GROUND TEST. (a) Configuration 1 - Use this configuration when using the TRONAIR Cabin Pressurization Test Unit. 1 Remove the control cable access panel located just aft of the nose wheel well. 2 Install and seal the adapter plate K-1285 where the access panel was removed. 3 Connect the 2-inch pressure nose and sense line from the test unit to the adapter plate. (b) Configuration 2 - Use this configuration when using test equipment that requires splicing into the bleed air plumbing at each nacelle to pressurize the cabin. 1 Open each inboard nacelle cover. 2 Disconnect the bleed air hose (1) from the bleed air manifold fitting (2) (Ref. Figure 202). 3 Attach the “Y” pressure hose from the test unit to the bleed air fitting (2) in each nacelle. 4 Set the MODE CONTROL switch to MAN. Hold the MAN TEMP in the INCR position until the ACM bypass valve and the ejector bypass valve are completely open, then release it. 5 Set the rate control knob on the CABIN CONTROLLER to the full clockwise position. 6 Bypass the airplane’s circuitry to energize the solenoids of the pressure regulator/ shutoff valves and the precooler valves as follows: a Remove the floorboards to gain access to the PCB rack (Items 7, 8 and 9) (Ref. Figure 2, Chapter 6-50-00, Sheet 2 of 3). b On aircraft with the A317 Bleed Air Control Module PCB, apply 28 vdc to pin 16 of the edgeboard connector P517 and P518 and leave the airplane’s cables connected to the PCB. c On aircraft with the A121 Relay Panel, apply 28 vdc to pins 4, 8, 13 and 17 of the edgeboard connector P275 and leave the airplane’s cables connected to the PCB.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (c) Configuration 3 - Use this configuration when using test equipment with a 2-inch pressure hose and an adapter plate. 1 Remove the control cable access panel located just aft of the nose wheelwell. 2 Install and seal the TK1794-5/939 adapter plate where the access panel was removed. 3 Open the right avionics compartment door. 4 Locate the AN fitting mounted in the pressure bulkhead immediately below the tube assembly of the vacuum regulator. 5 Remove both caps on the fitting, one inside the cockpit and one inside the avionics compartment. 6 Connect the sense line from the test unit to the fitting on the pressure bulkhead in the avionics compartment to sense cabin pressure. (5) Exit all personnel from the airplane. (6) Close and latch the cabin door, cockpit storm windows, cargo door, escape hatches, and any access panels that would contribute to cabin pressurization. (7) Secure the cabin door with the safety net and secure the cargo door with two safety straps called out under CABIN LEAK RATE GROUND TEST EQUIPMENT or use heavy strength nylon web safety straps. The minimum breaking force of the straps should be 26,000 pounds. If making structural repairs to the pressure vessel, attach a safety net large enough to cover the entire pressure vessel. WARNING: DO NOT under any circumstances, open the cabin door, cargo door, cockpit windows, escape hatches, or access panels while the cabin has ANY pressure applied during this test. DO NOT leave the test area during the test. (8) Attach a sign on the outside of the cabin door that reads: DANGER - AIRPLANE IN PRESSURE TEST (9) Connect shop air to the test unit. (10) Adjust the shop air on the test unit to 15 psi. (11) Turn the test unit ON to start the blower. (12) Open the pressure hose shutoff valve. CAUTION: DO NOT exceed a 2000 ft/min climb rate, as damage to the airplane’s CABIN CONTROLLER will result. DO NOT pressurize the cabin more than 4.3 psi. (13) Slowly open the airflow valve, monitor the climb rate and cabin pressure indicators. The climb rate indicator will initially show a decent as the pressure increases. As the cabin pressure reaches 4.3 psi, the cabin will begin to leak and the climb indicator will show a climb.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (14) When the cabin pressure reaches 4.3 psi, decrease the air flow valve until the climb rate indicator reads zero. At this point the amount of airflow into the cabin is equal to the amount of air escaping the cabin. NOTE: Standard cubic feet per minute (SCFM) is the true value of cubic feet per minute taking into consideration the effects of temperature and pressure on a gas, in this case, air. (15) Read the air flow indicator and convert the reading to SCFM by using temperature and pressure conversions supplied with the test unit technical data. If there are no conversion graphs provided with the test unit, and the air flow indicator reads cubic feet per minute directly, use the Flow Rate Correction Graph (Ref. Figure 204) to compensate for ambient temperature and pressure in the cabin and obtain a SCFM reading. (16) If the 4.3 psi is maintained with a maximum flow rate of 58 SCFM, the pressure vessel of the airplane is satisfactorily air tight. (17) If the flow rate is exceeded, listen for pressure leaks in the following areas: NOTE: Leaks can be detected by sound and pinpointed by feel. It may be necessary to attain maximum pressure, then shut down the test unit so that the sound of the blower will not interfere with the detection of leaks in the pressure vessel. (a) Nose wheel well. (b) Hoses, plumbing, and wire bundles piercing the pressure vessel. (c) Windshield seal and attachment screws. (d) Seal around storm window. (e) Seal around emergency exit hatches. (f) Seals around cabin doors and cargo doors. (g) Access panel seals. (h) Outflow valve mounting and gaskets. (i) Control cable seals in aft pressure bulkhead and in the fuselage adjacent to the inboard wing. (j) Fuselage belly drain plugs. (k) Fuselage Skin laps. (l) Antenna doublers and mounting areas. (18) Refer to Chapter 91-00-00 for the proper materials to be used to pressure seal the airplane. (19) Close the airflow valve on the test unit. (20) Close the shutoff valve on the test unit.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL WARNING: DO NOT under any circumstances, open the cabin door, cargo door, cockpit windows, escape hatches, or access panels while the cabin has ANY pressure applied. (21) Depressurize the airplane completely. On the TRONAIR Cabin Pressurization Unit, turn the air control valve to full decrease position to dump pressure. When the rate-of-climb indicator is at the horizontal position and the pressure gage reads zero, the cabin is depressurized. (22) Disconnect the pressure hose(s) from the airplane. (23) Disconnect the sense line from the airplane. (24) Remove the safety net from the cabin door. (25) Remove the safety straps from the cargo door. (26) Connect all bleed air plumbing, if removed. (27) Remove the test leads from the PCB’s, if installed. (28) Replace floorboards, if removed. (29) Remove the adapter plate, if used, install and seal the adapter plate with sealant tape (89, Table 1, Chapter 91-00-00).

Figure 204 Flow Rate Correction Graph

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Figure 205 Pressurization Control Components - Tail Cone

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Figure 206 Pressurization Control Components - Flight Compartment

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ENVIRONMENTAL SYSTEMS OUTFLOW VALVE MAINTENANCE PRACTICES

21-30-01 200200

1. PROCEDURES CAUTION: While performing maintenance on the pressurization system, NEVER use teflon tape on tube fittings as a thread seal. Teflon tape may have a tendency to shred and possibly contaminate the components of the pressurization system.

A. Removal CAUTION: An outflow valve that fails to maintain cabin differential pressurization within acceptable limits must be replaced. Field adjustment of these valves is not authorized and will void the warranty. (1) Remove the upholstery panel on the aft pressure bulkhead to gain access to the valves. (2) Remove the static port plumbing lines and the vacuum line from the valve. Tag the lines to aid in installation. (3) Remove the clamp securing the valve to the fitting on the aft pressure bulkhead and remove the valve.

B. Installation CAUTION: Never use teflon tape on pressurization fittings as a thread seal. (1) Install the valve on the fitting on the aft pressure bulkhead and secure with the clamp. (2) Ensure the packing is in place and connect the vacuum and static port lines. (3) Install the upholstery panel on the aft pressure bulkhead.

C. Cleaning Remove the outflow valve according to the procedure under OUTFLOW VALVE REMOVAL and clean the poppet seats with water and a mild soap or solvent (30, Table 1, Chapter 91-00-00).

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ENVIRONMENTAL SYSTEMS CABIN PRESSURE CONTROLLER MAINTENANCE PRACTICES

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1. PROCEDURES CAUTION: While performing maintenance on the pressurization system, NEVER use teflon tape on tube fittings as a thread seal. Teflon tape may have a tendency to shred and possibly contaminate the components of the pressurization system.

A. Removal (1) Remove the sidepanels from the pedestal. (2) Remove the top panel from the pedestal to gain access to the controller mounting screws. (3) Disconnect the lighting leads from the controller. CAUTION: Cutting of the wire leads in order to remove the electrical connector will necessitate a possibly unnecessary complete overhaul of the controller at the repair facility. (4) Disconnect the plumbing lines that connect the controller to the volume tank, the pneumatic relay and the vacuum regulator. (5) Remove the attaching screws and the controller from the pedestal.

B. Installation CAUTION: Never use teflon tape on pressurization fittings as a thread seal. (1) Install the controller on the top panel of the pedestal and secure with the attaching screws. Install the top panel on the pedestal. (2) Connect the plumbing lines to the controller. (3) Connect the lighting leads to the controller. (4) Install the sidepanels on the pedestal.

C. Filter Cleaning The pressurization controller filter is located in the controller assembly and should be cleaned as follows: (1) Remove the LH and the RH sidepanels from the pedestal. (2) Remove the filter from the controller housing. (3) Wash the screens and filter element in cleaning solvent (2, Table 1, Chapter 91-00-00). Dry with shop air.

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(4) Ensure that the orifice in the filter housing is free of foreign material. CAUTION: Be careful not to enlarge or damage the orifice in the housing. Proper operation of the system is dependent on the correct size of the orifice. (5) Install the filter element on the controller finger tight. (6) Install the LH and the RH sidepanels on the pedestal.

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ENVIRONMENTAL SYSTEMS PNEUMATIC RELAY MAINTENANCE PRACTICES

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1. PROCEDURES CAUTION: While performing maintenance on the pressurization system, NEVER use teflon tape on tube fittings as a thread seal. Teflon tape may have a tendency to shred and possibly contaminate the components of the pressurization system.

A. Removal (1) Remove the LH and RH sidepanels from the pedestal. (2) Remove the attaching screws securing the relay to the mounting brackets. (3) Disconnect and cap the plumbing lines from the relay. (4) Remove the relay from the pedestal.

B. Installation CAUTION: Never use teflon tape on pressurization fittings as a thread seal. (1) Connect the plumbing lines to the capped tees and elbow fitting on the relay. (2) Install the relay on the mounting brackets and secure with the attaching parts. (3) Install the sidepanels on the pedestal.

C. Filter Cleaning The pneumatic relay filter is located in the pedestal and should be cleaned as follows: (1) Remove the LH and the RH sidepanels from the pedestal. (2) Remove the filter from the underside of the relay. (3) Wash the screens and filter element in cleaning solvent (2, Table 1, Chapter 91-00-00). Dry with shop air. (4) Install the filter in the relay finger tight. (5) Install the LH and RH sides on the pedestal.

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ENVIRONMENTAL SYSTEMS VOLUME TANK MAINTENANCE PRACTICES

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1. PROCEDURES CAUTION: While performing maintenance on the pressurization system, NEVER use teflon tape on tube fittings as a thread seal. Teflon tape may have a tendency to shred and possibly contaminate the components of the pressurization system.

A. Removal (1) Remove the RH sidepanel from the pedestal (Ref. Figure 201). (2) Disconnect and cap the plumbing line from the volume tank. (3) Remove the attaching screws and the volume tank from the pedestal frame.

B. Installation CAUTION: Never use teflon tape on pressurization fittings as a thread seal. (1) Install the volume tank on the pedestal frame and secure with the attaching screws (Ref. Figure 201). (2) Connect the plumbing line to the volume tank. (3) Install the RH sidepanel on the pedestal.

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Figure 201 Pressurization Control Components - Flight Compartment

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ENVIRONMENTAL SYSTEMS CABIN ALTITUDE WARNING PRESSURE SWITCH MAINTENANCE PRACTICES

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1. PROCEDURES CAUTION: While performing maintenance on the pressurization system, NEVER use teflon tape on tube fittings as a thread seal. Teflon tape may have a tendency to shred and possibly contaminate the components of the pressurization system.

A. Removal (1) Remove electrical power from the airplane. (2) Obtain access and remove cabin floor panel 7 in Zone 143 (Ref. Chapter 6-50-00). (3) Identify, tag and disconnect the wires from the Cabin Altitude Warning Pressure Switch (Ref. Chapter 39-20-00, FORWARD LOWER CABIN ELECTRICAL EQUIPMENT). (4) Remove the attaching hardware and the Cabin Altitude Warning Pressure Switch from the Forward Lower Cabin Electrical Equipment panel.

B. Installation (1) Install the Cabin Altitude Warning Pressure Switch, with the attaching hardware, to the Forward Lower Cabin Electrical Equipment panel. (2) Connect the wires to the pressure switch and remove identification tags. Ensure that the wires are connected to the COMM and NO terminals of the switch. WARNING: Failure to connect the wires to the correct terminals of the pressure switch may result in false or no indication of CABIN ALTITUDE annunciator. (3) Install cabin floor panel 7. (4) Restore electrical power to the airplane.

2. FUNCTIONAL TEST This procedure verifies that the Cabin Altitude Warning Pressure Switch System is operating correctly. Method one involves the removal and installation of the switch. Method two requires the airplane to be flown to check the functionality of the pressure switch.

A. Method One (1) Perform the CABIN ALTITUDE WARNING PRESSURE SWITCH REMOVAL procedure in this section. (2) Place the pressure switch inside a vacuum chamber with the an ohmmeter connected to the COMM and NO terminals of the switch. If the vacuum chamber does not have a scale for determining degree of vacuum achieved, a certified altimeter must be used. Set the altimeter to 29.92 in. Hg. Place the altimeter in the vacuum chamber in a position so that it can be clearly read.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (3) Increase the vacuum in the chamber and check for onset of continuity at an increasing altitude of 12,000 to 12,500 feet. (4) After continuity is acquired, slowly decrease the vacuum in the chamber and check for termination of continuity at a decreasing minimum altitude of 10,500 feet. (5) The switch must be replaced if it does not meet these specifications. (6) Remove the pressure switch from the vacuum chamber. (7) On the aircraft, connect an insulated 22 gauge jumper wire between the exposed pressure switch wires on the Forward Lower Cabin Electrical Equipment panel. Note the CABIN ALTITUDE annunciator illuminates. (8) Remove jumper wire. (9) Perform the CABIN ALTITUDE WARNING PRESSURE SWITCH INSTALLATION procedure in this section.

B. Method Two WARNING: Comply with standard FAA regulations for oxygen usage when performing this test. (1) Fly the airplane at an altitude of 13,000 to 15,000 feet. (2) The pilot should select 13,000 feet cabin altitude on the Cabin Pressure Controller. (3) Set a hand held certified altimeter to 29.92 in. Hg. As cabin altitude rises, the pilot will note the cabin altitude at which the CABIN ALTITUDE annunciator light illuminates. The annunciator light shall illuminate at 12,000 to 12,500 feet altitude. (4) The pilot shall then select a cabin altitude below 10,000 feet. (5) As the cabin altitude decreases, the pilot will note the cabin altitude at which the annunciator light extinguishes. The annunciator light shall extinguish before 10,500 feet altitude. (6) The pressure switch must be replaced if it does not meet these specifications. Refer to CABIN ALTITUDE WARNING PRESSURE SWITCH REMOVAL in this section.

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ENVIRONMENTAL SYSTEMS HEATING DESCRIPTION AND OPERATION

21-40-00 00

1. GENERAL Bleed air from the engines enters the cabin ductwork through the ACM bypass valve and the ejector bypass valve. The ACM bypass valve begins opening first upon receiving a heat command from the cabin temperature controller. When the ACM bypass valve is fully open, the intergral open limit switch in the valve shunts the heat command from the cabin temperature controller to the ejector bypass valve, opening it. As the cabin temperature controller starts to issue cool commands, the ejector bypass valve begins closing. When the ejector bypass valve is fully closed, the cool command is shunted to the ACM bypass valve, closing it.

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ENVIRONMENTAL SYSTEMS HEATING TROUBLESHOOTING

100100

1. PROCEDURES Faults in the heating system will usually result in the faulty operation of other environmental systems as well. Since the functions of the bypass valves are closely interfaced with the vapor cycle system and temperature controls and dependent upon the proper function of the environmental bleed air system, troubleshooting of the bypass valves is the primary concern. The troubleshooting section will help the technician to differentiate between heating malfunctions which are most likely to be traced to the faults in other environmental systems and heating malfunctions which are the direct fault of one of the bypass valves (Ref. Charts in Figures 101 and 102).

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Figure 101 Heating Troubleshooting - Inadequate Heating

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Figure 102 Heating Troubleshooting - Excessive Heating

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ENVIRONMENTAL SYSTEMS BYPASS VALVE MAINTENANCE PRACTICES

200200

1. PROCEDURES Access to the bypass valves can be gained through the upper inboard left wing panel. In the event of a malfunction in one of the valves, it will be necessary to remove the valve for overhaul or repair. The two bypass valves are similar in appearance and located near the air cycle machine. The ejector bypass valve is attached to the check valve T assembly just outboard of the dual heat exchanger. The ACM bypass valve is located immediately to and outboard of the air cycle machine.

A. Removal (1) Locate the appropriate valve and disconnect the electrical connector from the valve motor (Ref. Figures 201 and 202). (2) Push back the insulation covering the couplings and remove the coupling clamps. (3) Remove the mounting bolts attaching the valve stabilizing bracket to the wing structure (ACM bypass valve only). (4) Lift the valve and stabilizing bracket out of the wing.

B. Installation (1) Position the valve in the wing (Ref. Figures 201 and 202). (2) Attach the stabilizing bracket to the wing structure using the same bolts removed earlier (ACM bypass valve only). (3) Install the coupling clamps and torque the clamp nuts to 35 ± 2 inch-pounds. (4) Position the insulation around the couplings. (5) Connect the electrical connector.

21-40-00

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Figure 201 ACM Bypass Valve

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21-40-00

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 202 Ejector Bypass Valve

21-40-00

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ENVIRONMENTAL SYSTEMS COOLING DESCRIPTION AND OPERATION

21-50-00 00

1. GENERAL Cabin cooling in the Model 1900 AIrliner Series airplanes is accomplished in part by the air cycle system. Air cycle system cooling is augmented by cooling from the vapor cycle system. Automatic control of both systems is accomplished through the cabin temperature controller and it’s associated secondary controls. The intent of the chapter is to aid the technician in determining to which cooling system a fault can be traced in the event of cooling failure. Due to the complex interplay between systems, it will not always be apparent which system is malfunctioning. The troubleshooting portion of the chapter will assist the technician in narrowing down faults and will direct the technician to troubleshoot a particular system. 21-51-00 and 21-52-00 deal with troubleshooting and maintenance of the air cycle system and the vapor cycle system respectively. 21-60-00 contains troubleshooting and maintenance information for the temperature controlling system.

21-50-00

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ENVIRONMENTAL SYSTEMS COOLING TROUBLESHOOTING

100100

1. PROCEDURES Differential troubleshooting of the cabin cooling system is primarily a matter of ruling out a malfunction in one system or the other. Assessment of vapor cycle system operation was determined to be the simplest approach. The troubleshooting chart graphically illustrates the troubleshooting sequence (Ref. Figure 101).

Figure 101 Cabin Cooling Troubleshooting

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ENVIRONMENTAL SYSTEMS COOLING MAINTENANCE PRACTICES

200200

1. PROCEDURES The following procedure can be used to evaluate the operational efficiency of the vapor cycle system. This check should be performed only when the outside air temperature is above 50°F to insure that the 40°F OAT cutout switch is closed.

2. VAPOR CYCLE SYSTEM A. Operational Check (1) Place thermometers, known to be accurate, in a forward eyeball outlet and in an aft eyeball outlet. (2) Start the right hand engine and increase the engine rpm to 62% N1. Do not turn on the environmental bleed air on. (3) Place the cabin temperature mode switch in MANUAL and hold the manual temperature control switch in DECREASE for 60 seconds. (4) Take the cabin air temperature reading near the floor of the cabin and after the temperature of the air coming out of the eyeball outlets has stabilized, 5 to 10 minutes should be sufficient, record the temperatures of the eyeball outlets air. (5) Subtract each of the temperature readings at the eyeball outlets from the cabin air temperature recorded earlier. A differential of at least 20°F at both eyeball outlets would be indicate of a properly operating system.

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ENVIRONMENTAL SYSTEMS AIR CYCLE SYSTEM DESCRIPTION AND OPERATION

21-51-00 00

1. GENERAL The Model 1900 Series Airliner air cycle system consists of an air cycle machine (ACM), dual heat exchanger, water collector, fog nozzle, inline filter assembly and a recirculating ejector. The amount of cooling supplied by the air cycle system is controlled by varying the amounts of hot air which bypass portions of the air cycle machine. Control over this bypass flow is a function of the temperature control system. The heat exchanger and the air cycle machine are mounted together to form an integral unit (Refrigeration Package). Hot engine bleed air becomes partially cooled as it flows through the primary core section of the heat exchanger. The primary cooled air enters the air cycle machine compressor which raises the air pressure and temperature. The compressed air returns to the secondary core of the heat exchanger where the heat of compression is removed. The bleed air then enters the turbine section of the air cycle machine where it is cooled by expansion across the turbine nozzles. This expanding air also drives the air cycle machine turbine. The driven turbine in turn drives the compressor rotor and the fan rotor. The rotating air cycle machine fan draws cooling air from the ambient air inlet, located on the under side of the left wing, through the cooling air passageways in both the primary and secondary cores of the heat exchanger and discharges the cooling air overboard. The ambient air flowing through the primary and secondary cores of the heat exchanger absorbs heat from the bleed air circuits (Ref. Figure 1). As a result of cooling, moisture condensation occurs in the secondary core of the heat exchanger upstream of the air cycle machine turbine section. The condensed moisture is collected by a water collector, routed to a fog nozzle and is sprayed into the ambient air inlet of the secondary core of the heat exchanger where it assists in cooling bleed air temperature by evaporation. The recirculating ejector is a welded aluminum tube with two duct ends. The inlet duct accepts conditioned bleed air from the turbine outlet of the air cycle machine. The outlet duct provides for the entry into the cabin of the mixed conditioned and recirculated cabin air. An electrical heater surrounds the bleed air duct at the turbine air inlet discouraging any ice accumulation inside the cold duct.

21-51-00

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Figure 1 ACM Ambient Air Ducts

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ENVIRONMENTAL SYSTEMS AIR CYCLE SYSTEM TROUBLESHOOTING

100100

1. PROCEDURES Troubleshooting of the air cycle system is a fairly simple task as generally air cycle system cooling problems can be traced to faults in other environmental subsystems. The troubleshooting chart presents a logical approach to troubleshooting of the air cycle system (Ref. Figure 101).

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INADEQUATE COOLING ARE THE AMBIENT AIR INLET AND HEAT EXCHANGER CORE CLEAN AND FREE OF OBSTRUCTIONS?

NO

YES

CORRECT PROBLEM

NO

IS THE FOG NOZZLE OR INLINE WATER FILTER CLOGGED?

YES

NO

ARE THE BYPASS VALVES CLOSING? (REF. CHART 2 EXCESSIVE HEATING, 21-40-00)

YES

CLEAN NOZZEL AND INLINE FILTER

CORRECT PROBLEM

NO

ARE BLEED AIR TEMPERATURE AND PRESSURE CORRECT? (REF. 21-10-00 BLEED AIR CONTROL TROUBLESHOOTING)

YES

CORRECT PROBLEM

FAULT IS IN AIR CYCLE MACHINE

UC21B 071774AA.AI

Figure 101 Cabin Cooling Troubleshooting

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ENVIRONMENTAL SYSTEMS AIR CYCLE SYSTEM MAINTENANCE PRACTICES

200200

1. PROCEDURES Maintenance of the air cycle system should usually be a process of cleaning the heat exchanger inlet and the fog nozzle-filter assembly; however, should a fault develop in the air cycle machine or heat exchanger, removal of the air cycle machine and heat exchanger assembly (refrigeration package) would be required.

2. AIR CYCLE MACHINE A. Servicing Check the quantity of oil contained in the see-thru oil sump and add oil as required to maintain a full sump. Manufacturer's recommendation for lubricating oil to be used in the air cycle machine is Exxon 2389 or any oil conforming to MIL-L-7808 (72, Table 1, Chapter 91-00-00).

3. WATER COLLECTOR DRAIN TUBE A. Weld Repair Information to conduct this repair is provided by HAMILTON SUNDSTRAND WATER COLLECTOR P/N 775642 DRAIN TUBE WELD REPAIR LETTER OF INSTRUCTIONS, in Chapter 21 of the Model 1900 Series Component Maintenance Manual.

4. OIL A. Replacement (Oil Sump With Fill Plug) (1) Loosen the leading edge (UA-1 and After, UB-1 and After: Wing Access Panel 55; UC-1 and After: Wing Access Panel 24) as necessary to remove the lower access panel (UA-1 and After, UB-1 and After: Wing Access Panel 63; UC-1 and After: Wing Access Panel 18) and remove the lower access panel from the left center wing (Ref. Chapter 6-50-00). (2) Locate the oil sump on the bottom of the air cycle machine. (3) Cut the lockwire and remove the oil sump fill plug. If a lubrication instruction tag is attached to the fill plug, remove and discard the tag. (4) Remove all the oil from the oil sump using a syringe-type suction device. NOTE: Do not mix different types of oil when filling. (5) Fill the oil sump to overflowing with lubricating oil, preferably Exxon 2389 or any oil conforming to MIL-L-7808 (72, Table 1, Chapter 91-00-00). (6) Install a new packing on the fill plug and install the plug in the sump, tightening the plug to 10 ± 2 inch-pounds. (7) Lockwire the plug to the oil sump.

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B. Replacement (Oil Sump Without Fill Plug) (1) Loosen the leading edge (UA-1 and After, UB-1 and After: Wing Access Panel 55; UC-1 and After: Wing Access Panel 24) as necessary to remove the lower access panel (UA-1 and After, UB-1 and After: Wing Access Panel 63; UC-1 and After: Wing Access Panel 18) and remove the lower access panel from the left center wing (Ref. Chapter 6-50-00). (2) Locate the oil sump on the bottom of the air cycle machine. (3) Remove and retain the screws and washers that attach the sump to the air cycle machine turbine housing. (4) Lower the sump away from the turbine housing and pour the oil from the sump. (5) Wipe the oil residue from the sump with a dry, lint-free cloth. (6) Install a new M83248/1-033 packing on the sump. NOTE: Do not mix different types of oil when filling. (7) Fill the sump to the fill line with lubricating oil, preferably Exxon 2389 or any oil conforming to MIL-L-7808 (72, Table 1, Chapter 91-00-00). (8) Install the sump on the turbine housing with the screws and washers retained during removal. (9) Torque the screws 12 to 18 inch-pounds plus running torque (running torque must fall between 2 to 13 inch-pounds or replace screw). (10) Install the leading edge and the lower access panel on the left center wing.

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ENVIRONMENTAL SYSTEMS REFRIGERATION PACKAGE MAINTENANCE PRACTICES

21-51-01 200200

1. PROCEDURES A. Removal (1) Remove the left inboard wing leading edge and the upper and lower wing access panels (Ref. Figure 201). (2) Drain the air cycle machine oil sump (Ref. 21-51-00, AIR CYCLE MACHINE SERVICING). (3) Disconnect and remove the cooling air exhaust plenum and a section of bleed air line along the spar cap. (4) Disconnect the bleed air duct at the inlet of the first stage heat exchanger. (5) Disconnect and remove the mufflers at the ACM bypass inlet. (6) Disconnect the bleed air ejector duct at the turbine outlet. (7) Disconnect the water drain tube at the water collector. (8) Remove the subspar opening stiffener screws. (9) Remove the two mounting bolts which secure the heat exchanger to the wing structure. (10) Remove the two mounting fasteners from the housing mounting bracket on the air cycle machine under the compressor. (11) Remove the air cycle machine through the lower wing access opening.

B. Installation (1) Place the air cycle machine into the opening in the wing, being careful not to damage the see-thru oil sump on the lower side of the air cycle machine (Ref. Figure 201). (2) Install the mounting fastener to the housing mounting bracket on the air cycle machine. (3) Install the two mounting bolts which secure the heat exchanger to the wing structure, replacing the shims as required to vertically align the fasteners at the forward mount. (4) Connect the water drain tube to the water collector. (5) Connect the mufflers at the ACM bypass inlet. (6) Connect the bleed air duct at the inlet of the first stage heat exchanger. (7) Install the plenum and bleed air duct section removed earlier. (8) Service the air cycle machine oil sump (Ref. 21-51-00, AIR CYCLE MACHINE SERVICING). (9) Install the left inboard upper and lower wing panels and the wing leading edge.

21-51-01

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Figure 201 Refrigeration Package

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21-51-01

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

ENVIRONMENTAL SYSTEMS FOG NOZZLE AND IN-LINE FILTER MAINTENANCE PRACTICES

21-51-02 200200

1. PROCEDURES A. Removal and Installation (1) Locate the fog nozzle and in-line filter assembly immediately adjacent to the first stage heat exchanger inlet (Ref. Figure 201). (2) Release the holding bracket and remove the fog nozzle. (3) Disconnect the water line. (4) Disassemble the filter and clean the filter mesh with solvent (86, Table 1, Chapter 91-00-00). Do not remove the filter mesh from the filter housing. (5) Clean the nozzle and orifice with solvent (86, Table 1, Chapter 91-00-00), being careful not to enlarge the orifice during cleaning. (6) Assemble the in-line filter and install the filter and nozzle on the bulkhead bracket. (7) Connect the water drain tube.

Figure 201 Fog Nozzle Filter Assembly

21-51-02

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

ENVIRONMENTAL SYSTEMS RECIRCULATING EJECTOR MAINTENANCE PRACTICES

21-51-03 200200

1. PROCEDURES A. Removal (1) Open the LEFT and RIGHT BLEED AIR CONTROL circuit breakers on the right circuit breaker panel. (2) Remove the carpet (Ref. Chapter 25-20-01, CARPET REMOVAL AND INSTALLATION) and floorboards 16D and 9 (UA-1 and After) or floorboards 16D and 10 (UB-1 and After and UC-1 and After) (Ref. Chapter 6-50-00). (3) Perform the REFRIGERATION PACKAGE REMOVAL procedure (Ref. 21-51-01). (4) Disconnect the recirculating ejector heater electrical connector (5) (Ref. Figure 201). (5) Remove bolts (11) attaching the outer retaining ring (10) and flexible diaphragm seal (12) to the fuselage pressure plate (13) (Ref. Detail B). (6) Loosen clamp (4) at the inboard end of the recirculating ejector. Slide the clamp inboard onto the flexible sleeve (3) (Ref. Detail C). (7) Remove the ejector attach bolt (9) securing the recirculating ejector to the support bracket (8) on the inboard side of the left seat track (6). (8) Remove the recirculating ejector (1) from the airplane through the refrigeration package cavity. The inboard end of the ejector should slide out of flexible sleeve (3), which should remain clamped to the plenum inlet (2).

B. Installation (1) Position the recirculating ejector (1) in the airplane through the refrigeration package cavity and insert the inboard end into the flexible sleeve (3) (Ref. Figure 201). (2) Install the recirculating ejector attach bolt (9) attaching the recirculating ejector (1) to support bracket (8) on the inboard side of the left seat track (6). (3) Install clamp (4) over the flexible sleeve (3), clamping it to the beaded end of the recirculation ejector (1). (4) Install the bolts (11) attaching the outer retaining ring (10) and diaphragm seal (12) to the fuselage pressure plate (13). (5) Connect the recirculating ejector heater electrical connector (5). (6) Perform the REFRIGERATION PACKAGE INSTALLATION (Ref. 21-51-01). There should be a gap of 0.12 inch between the air cycling system outlet and the ejector inlet (Ref. Detail B). Adjust the gap by changing the number of shims (7) between the seat track (6) and the recirculating ejector support bracket (8) (Ref. Detail C).

21-51-03

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (7) Close the LEFT and RIGHT BLEED AIR CONTROL circuit breakers on the right circuit breaker panel.

C. Heater Blanket Operational Check (1) Remove wing access panel 57 (UA-1 and After; UB-1 and After) or panel 18 (UC-1 and After) from the top of the left wing (Ref. Chapter 6-50-00, WING ACCESS PANELS). (2) Disconnect the ejector heater (E184) connector from the airplane harness electrical connector (P493). Refer to the applicable Model 1900/1900C Airliner Wiring Diagram Manual, Chapter 21-31-01. (3) Place the BATT switch to the ON position. (4) Place the LEFT BLEED AIR VALVE switch to the ON position. (5) Using a multimeter, perform the following Steps: (a) Check for battery voltage at Pin A on the airplane harness connector (P493). (b) Check for ground at Pin B on the airplane harness connector (P493). (6) Place the LEFT BLEED AIR VALVE switch to the OFF position. (7) Place the BATT switch to the OFF position. (8) Using a multimeter, measure the heater element resistance between Pins A and B. Replace the ejector heater if the resistance reads other than 10.37 to 11.46 ohms. (9) Connect the ejector heater (E184) connector to the airplane harness connector (P493). (10) Install wing access panel 57 (UA-1 and After; UB-1 and After) or panel 18 (UC-1 and After) (Ref. Chapter 6-50-00, WING ACCESS PANELS).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

2

3 4

C A 1

RECIRCULATED CABIN AIR

1. RECIRCULATING EJECTOR 2. AIR DISTRIBUTION PLENUM 3. FLEXIBLE SLEEVE (CONNECTOR) 4. CLAMP 5. ELECTRICAL CONNECTOR 6. LEFT SEAT TRACK 7. SHIM 8. SUPPORT BRACKET 9. EJECTOR ATTACH BOLT 10. OUTER RETAINER RING 11. BOLTS 12. DIAPHRAGM SEAL 13. FUSELAGE PRESSURE PLATE

DETAIL

B

A 5

ACM CONDITIONED AIR

7 13 8 6

0.12 INCH GAP

9 DETAIL AIR CYCLE MACHINE OUTLET

C

12 11 10 DETAIL

B

UC21B 071078AA.AI

Figure 201 Recirculating Ejector Installation

21-51-03

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

ENVIRONMENTAL SYSTEMS VAPOR CYCLE SYSTEM DESCRIPTION AND OPERATION

21-52-00 00

1. GENERAL The vapor cycle cooling system utilizes refrigerant to provide cooling when the airplane is on the ground and to provide supplemental cooling when the air cycle cooling system is performing at maximum levels. On airplanes without Kit 129-5020 installed, use refrigerant R-12. On airplanes with Kit 129-5020 installed, use refrigerant R-134a. A compressor, a condenser with a 40,000-BTU capacity, and two 12,500-BTU evaporators are utilized to cycle the refrigerant from a gas to a liquid state to provide cooling of the passenger compartment and flight compartment. Adjustable outlets, located adjacent to the cabin armrests and in the flight compartment overhead panel, distribute cool air produced by the vapor cycle cooling system. The temperature control switches used to control the air cycle cooling system are used to control the vapor cycle cooling system. The cabin temperature mode switch, the cabin temperature selector, and the manual temperature control switch are located on the copilot's inboard subpanel. Refer to 21-60-00 for further information on the operation of the temperature controls. The compressor is equipped with an electric clutch and a relief valve and is installed in the accessory gear box section of the RH engine. The discharge line of the compressor is connected to the condenser, located adjacent to its blower in the RH inboard wing. The receiver-dryer removes moisture from the refrigerant and is located under the center aisle floor in line with the third cabin window. The two evaporators and the corresponding blowers are located under the center aisle floorboards. The forward evaporator and blower are in line with the first cabin window and the aft evaporator and blower are in line with the sixth cabin window. The blowers distribute cool air to the adjustable outlets in the cabin and flight compartment. Refer to 21-20-00 for further information on the blowers and vapor cycle cool air distribution. A sight gage, an overpressure cutout switch and an underpressure cutout switch, and service valves are installed in the refrigerant lines located in a service box that is accessible through a door in the underside of the fuselage. The service box is in line with the wing flaps. Reset switches for the pressure cutout switches are located in the service box (Ref. Figure 1). When the low ambient temperature limit switch is closed and RH engine speed exceeds 62% N1, the vapor cycle cooling system can be activated by the air cycle cooling system (Ref. Figure 2). The limit switch will be closed when ambient temperature, as sensed in the condenser inlet, is high enough to allow cooling of the airplane cabin (50°F ± 5°). When the ACM bypass valve and the ejector bypass valve are closed and the low ambient switch is closed, the cool coil of the heat-cool command relay is energized, and power is supplied to the cool command input of the N1 speed sensor PCB. The compressor clutch is then allowed to engage 10 seconds after the RH engine speed increases above 62% N1. When the RH landing gear downlock switch is closed (the airplane is on the ground), a ground circuit is completed and energizes the coil of the condenser blower relay. Power is then supplied to the condenser blower and the blower moves air across the condenser and out the vents in the wing underside. When the airplane is in flight and the downlock switch is not closed (the landing gear is up), the condenser is cooled by the flow of ram air that enters through the inlet in the RH wing leading edge. When the cabin has been cooled to a preset temperature and the ACM bypass valve reaches full open, the cool coil of the heat-cool command relay is de-energized and the heat coil of the relay is energized. Power is then removed from the compressor clutch and it disengages. If the N1 speed of the RH engine falls below 62% while the vapor cycle cooling system is operating, the annunciator light labeled AIR COND N1 LOW (green) will illuminate, power will be removed from the compressor clutch and it will disengage (Ref. Figure 2). This compressor cutout avoids overloading the engine when it is operating at low speeds. If RH engine speed is above 62% N1 but ambient temperature falls below 30°F ± 5°, the limit switch in the condenser inlet will open and the clutch will disengage to prevent icing of the system components.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL The condenser blower relay is located in the RH nacelle. The heat-cool command relay is located under the cabin floorboards immediately aft of the main wing spar. The N1 speed sensor is located under the cabin floorboards in line with the second cabin window. Refer to Chapter 39-20-00 for further information on the relays and the N1 speed sensor PCB. Refer to 21-40-00 for further information on the ACM bypass valve and the ejector bypass valve. When the electric clutch on the compressor is engaged, the refrigerant is compressed into a high pressure, high temperature gas. As this gas is pumped through the condenser, cooling air is vented across the condenser and removes heat from the gas, condensing it into a liquid state (Ref. Figure 1). This liquid refrigerant passes through the receiver-dryer where moisture and impurities are removed. The refrigerant is then metered by the thermostatic expansion valves (TEV) and flows into the evaporators. Heat from recirculating cabin air is absorbed by the refrigerant in the evaporators and the liquid refrigerant evaporates into a gas. Vent blowers draw the recirculating air through the coils in the forward and aft evaporator assemblies. This gas is returned to the compressor where it is compressed into the higher pressure discharge gas.

A. Hot Gas Bypass Valves A temperature sensing switch is installed on the outlet line of each evaporator. If the temperature of the refrigerant outlet line is 33°F or below, the switch opens a solenoid valve in a hot gas line containing discharge gas from the compressor. This hot gas bypasses the condenser and the expansion valves and flows directly into the evaporator to prevent moisture from freezing on the evaporator. When the temperature in the evaporator outlet line reaches 45°F, the thermoswitch closes the hot gas bypass valve and refrigerant flow is returned to its normal flow through the condenser and the expansion valves.

B. Pressure Cutout Switches If the discharge pressure in the system increases beyond a safe limit, the overpressure cutout switch will open, removing power from the compressor clutch. If the suction pressure in the refrigerant line falls below a preset limit, the underpressure cutout switch will close and cause the thermal relay to open. When the thermal relay opens, power is removed from the compressor clutch. If power has been removed from the clutch by an open cutout switch, the switch must be reset to allow the compressor clutch to engage. The reset switches are located in the service box in the underside of the fuselage.

C. Vent Blowers The vent blowers are controlled by the vent blower switch and the cabin temperature mode switch. Both switches are located on the copilot's inboard subpanel. The blowers will operate in the high speed or the low speed when the cabin temperature mode switch is set to each position. When the blower switch is set to AUTO, the vent blowers will operate in the low speed when the mode switch is set to each position except OFF. The control of these blowers is independent of all other components of the environmental system. Refer to 21-20-00 for further information on the blowers and their relays.

D. Pressure Relief Valve A pressure relief valve is installed on the discharge service port of the compressor to protect the compressor from an overpressure condition. The valve relieves pressure at 450 psig.

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21-52-00

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 1 Vapor Cycle System Refrigerant Flow Schematic

21-52-00

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Figure 2 Vapor Cycle Cooling System Electrical Schematic

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ENVIRONMENTAL SYSTEMS VAPOR CYCLE SYSTEM TROUBLESHOOTING

100100

1. PROCEDURES Refer to Figures 101 thru 104 to troubleshoot the vapor cycle cooling system. Refer to Figure 105 to troubleshoot the condenser blower control system. Refer to Table 101 thru Table 108 and Figures 106 thru 114 for sample fault conditions and suggested corrective actions.

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Figure 101 Vapor Cycle System Refrigerant Flow Schematic

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21-52-00

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 102 Vapor Cycle Cooling System Electrical Schematic

21-52-00

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Figure 103 Troubleshooting - Inadequate Cooling (Compressor Operating)

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Figure 104 Troubleshooting - Inadequate Cooling (Compressor Not Operating)

21-52-00

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Figure 105 Troubleshooting - Condenser Blower Control

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21-52-00

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL The following are some sample fault conditions and suggested corrective actions. For approximate normal operating pressures for R-134a refrigerant in the Model 1900C Airliner vapor cycle system (Ref. Figure 203) R-134a Vapor Cycle System Suction & Discharge Pressures Versus Ambient Temperatures illustration in the MAINTENANCE PRACTICES section. The charted pressures are based on conditions in which the cabin door(s) are open and the cabin is heat soaked to near ambient temperature. All the cabin eyeball air outlets are open and the right engine is operated at about 70% N1. Pressures may then be noted after they stabilize (about 5 minutes).

CONNECT TO HIGH (DISCHARGE) SIDE OF SYSTEM

CONNECT TO LOW (SUCTION) SIDE OF SYSTEM

UC21B 023314AA.AI

Figure 106 A Typical Air Conditioner Service Set

21-52-00

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UC21B 023318AA.AI

Figure 107 Normally Functioning System @ 85°F

Table 101 Vapor Cycle System - Normally Functioning System @ 85°F GAUGE READINGS

OTHER INDICATIONS

Low Side Gauge - Normal - 38 psi

Sight glass - Clear

High Side Gauge - Normal - 185 psi

Discharge air - Cold

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UC21B 023319AA.AI

Figure 108 Moisture in System

Table 102 Vapor Cycle System - Moisture in System PROBLEMS Low Side Gauge - Normal, then sometimes drops to below zero. (and) High Side Gauge - Normal, then sometimes goes high.

PROBABLE CAUSE

CORRECTIVE ACTION

Moisture in system freezes, temporarily stopping cycle. However, normal system operation returns when ice melts.

1. Reclaim refrigerant from system. 2. Replace receiver-dryer. Replace any oil removed with receiver-dryer. 3. Remove moisture by evacuating system. 4. Charge system with R-134a. 5. Operate system and check performance.

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Figure 109 Low R-134a Charge

Table 103 Vapor Cycle System - Low 134a Charge PROBLEM Low Side Gauge - Normal or Low (and) High Side Gauge - Low

PROBABLE CAUSE System slightly low on R-134a, due to leak or incorrect charge.

Sight glass - bubbles may be continuously visible. Return lines warm. Low pressure cutout switch tripped.

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21-52-00

CORRECTIVE ACTION 1. Leak test system. 2. Reclaim refrigerant from system if necessary. 3. Repair system leaks. 4. Charge system with R-134a. 5. Operate system and check performance.

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

UC21B 023321AA.AI

Figure 110 Poor Refrigerant Circulation

Table 104 Vapor Cycle System - Poor Refrigerant Circulation PROBLEM Low Side Gauge - zero to negative (and) High Side Gauge - Low Receiver-dryer - frost on tubes from receiver-dryer to evaporator units.

PROBABLE CAUSE Refrigerant flow obstructed by debris. Receiver-dryer clogged.

CORRECTIVE ACTION 1. Reclaim refrigerant from system. 2. Replace receiver-dryer or remove debris. Replace any oil removed. 3. Charge system with R-134a. 4. Operate system and check performance.

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UC21B 023322AA.AI

Figure 111 No Refrigerant Circulation

Table 105 Vapor Cycle System - No Refrigerant Circulation PROBLEM Low Side Gauge - zero to negative (and) High Side Gauge - Low Receiver-dryer - frost or moisture on tubes before and after receiver-dryer.

PROBABLE CAUSES Refrigerant flow restricted by debris or moisture or refrigerant flow obstructed by gas leakage from expansion valve heat sensing tube.

CORRECTIVE ACTION 1. Reclaim refrigerant from system. 2. Check heat sensing tube at expansion valve. Replace expansion valve if necessary. 3. Remove expansion valve and attempt removal of debris. If debris cannot be removed, replace expansion valve. 4. Replace receiver-dryer and replace any oil removed. 5. Charge system with R-134a. 6. Operate system and check performance.

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Figure 112 Air in System

Table 106 Vapor Cycle System - Air in System PROBLEM Low Side Gauge - High (and) High Side Gauge - High Sight glass - bubbles visible during system operation.

PROBABLE CAUSE Air is present in system possibly from inadequate evacuation procedure.

CORRECTIVE ACTION 1. Evacuate system. 2. Charge system with R-134a. 3. Operate system and check performance.

Pipes - low pressure pipes are hot to the touch.

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Figure 113 Expansion Valve Improperly Mounted or Heat Sensing Tube Defective

Table 107 Vapor Cycle System - Expansion Valve Improperly Mounted or Heat Sensing Tube Defective PROBLEM Low Side Gauge - High (and) High Side Gauge - High

PROBABLE CAUSE Excessive refrigerant in low side pipes possibly from expansion valve being opened too wide.

Large amount of frost or moisture on low side pipes.

CORRECTIVE ACTION 1. Check heat sensing tube for proper installation. 2. If heat sensing tube is properly positioned, evacuate system. 3. Check expansion valve and replace if defective. 4. Charge system with R-134a. 5. Operate system and check performance.

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Figure 114 Compressor Malfunction

Table 108 Vapor Cycle System - Compressor Malfunction PROBLEM Low Side Gauge - High (and) High Side Gauge - Low

PROBABLE CAUSE Internal compressor leak or compressor mechanically broken or belt loose or broken.

CORRECTIVE ACTION 1. Evacuate system as necessary. 2. Repair or replace compressor or belt. 3. Charge system with R-134a as necessary. 4. Operate system and check performance.

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1. PROCEDURES A. Precautionary Service Procedures Before attempting maintenance that requires opening of refrigeration lines or compressor fittings, maintenance personnel should be thoroughly familiar with the pertinent instructions. These instructions should be followed carefully to ensure that the system functions properly. If moisture is allowed to enter refrigerant lines, ice can form in the lines and hydrofluoric acid can form, causing damage to system components. Contamination of the system with dirt can cause blockage and damaging wear in the compressor. All replacement subassemblies for the vapor cycle system are sealed and dehydrated. They should remain sealed until immediately prior to making connections. Refrigerant lines and other components should be at room temperature before uncapping to prevent the condensation of moisture from entering the system. If a connection is not made immediately after uncapping a component, it should not remain unsealed for more than 15 minutes. If the time period is longer, reseal the connections. New compressors are provided with four ounces of Oil (117, Table 1, Chapter 91-00-00) and are charged with a mixture of refrigerant (116, Table 1, Chapter 91-00-00) and dry nitrogen to provide an internal pressure that is slightly above atmospheric pressure. For airplanes without Kit 129-5020 installed, the oil must be drained from the new compressor and replaced with four ounces of 500-viscosity oil (70, Table 1, Chapter 91-00-00) and recharged with refrigerant (71, Table 1, Chapter 91-00-00). For airplanes with Kit 129-5020 installed, the new compressors do not require exchanging the oil and refrigerant. Care should be taken to prevent damage to all fittings and connections. Minute damage to a connection could cause it to leak. Any fittings contaminated with grease or dirt should be cleaned with a cloth dampened with alcohol. Do not use a chlorinated solvent such as trichloroethylene as a cleaning agent because it adds contaminants. If dirt, grease or moisture inside lines cannot be removed, the line must be replaced. On airplanes without Kit 129-5020 installed, use compressor oil (70, Table 1, Chapter 91-00-00). On airplanes with Kit 129-5020 installed, use air conditioning oil (117, Table 1, Chapter 91-00-00). Apply a small amount of clean refrigeration oil (as previously noted) to all line connections and dip packings in the oil to help make a leak-resistant connection.

B. Vapor Cycle System Maintenance Notes WARNING: The vapor cycle system is a high pressure system. Before disconnecting a refrigerant line, depressurize the system with a recycle servicing unit, then purge the entire system with a vacuum pump operating at a 125-micron level. A face shield should be worn when performing maintenance on the lines because refrigerant coming in contact with the eyes can cause loss of sight. Also wear gloves and avoid breathing any vapors. Do not smoke when servicing the system because the refrigerant converts into a highly toxic gas when exposed to an open flame.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL CAUTION: On airplanes without Kit 129-5020 installed, use only air conditioning refrigerant (71, Table 1, Chapter 91-00-00) in this system. On airplanes with Kit 129-5020 installed, use only refrigerant (116, Table 1, Chapter 91-00-00) in this system. All other refrigerants, particularly those containing methyl chloride, will cause deterioration of the aluminum components. Insufficient torque can result in loose joints and excessive torque can result in deformed connecting parts. Either condition can result in refrigerant leakage. NOTE: Due to air quality-control regulations being enacted in the United States, R-12 and R-134a refrigerant cannot be vented into the atmosphere. When performing maintenance on the vapor cycle system where refrigerant R-12 can escape from the system, evacuate the system with a recovery or recycle servicing unit that will salvage the refrigerant. For information on connecting aluminum fittings in the system, refer to Table 5 in Chapter 91-00-00. The receiver-dryer should be the last component connected. It should be connected last to ensure maximum protection of the vapor cycle system against moisture.

C. Refrigerant Leak Detection Table 201 Special Tools and Equipment TOOL NAME

PART NO.

SUPPLIER

USE

1. Detector Dye Injector (R-134a System)

TP-3886A or equivalent

Tracer Products 956 Brush Hollow Road, Westbury, NY 11590

To inject detector dye into R-134a vapor cycle systems.

2. Detector Dye Injector (R-12 System)

TP-3880 or equivalent

Tracer Products 956 Brush Hollow Road, Westbury, NY 11590

To inject detector dye into R-12 vapor cycle systems.

3. Ultraviolet Lamp

TP-1200P or equivalent

Tracer Products 956 Brush Hollow Road, Westbury, NY 11590

Fluoresces ultraviolet dye.

A reduction of system cooling ability and declining pressures, significantly below nominal pressures shown on the graphs may be an indication of leakage (Ref. Figures 201 and 202). For the R-12 system only, the continual presence of bubbles in the refrigerant may indicate a partial loss of refrigerant. The sight glass should be checked during system operation at maximum available ambient and cabin temperatures. The sight glass is located in a service box in the underside of the fuselage in line with the flaps. Streams of bubbles or foam seen in the glass indicate the refrigerant quantity is low. Colored fluid in the sight glass may indicate the system contains a leak detector dye additive. NOTE: Do not use bubbles in the sight glass as an indication of undercharge with R-134a refrigerant. A system fully charged with R-134a may still exhibit bubbles or cloudiness in the sight glass. Use the pressure graph as an indication of proper charge.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL If a connection does not seal by the normal torque procedures, the use of soft flare gaskets in the flare fitting is permissible. If a leak occurs at a flare fitting containing a soft seal, replacement of the fitting is recommended. Copper gaskets should be used in copper and brass fittings and aluminum gaskets should be used in aluminum fittings. Minor leakage of less than 2 ounces per year is permissible at the compressor shaft seal. If a loss of refrigerant is suspected, the system plumbing should be inspected to determine the source of the leak. Large leaks can be located by the appearance of oily spots where oil has been carried out by escaping refrigerant. Smaller leaks can be detected by the following tests:

D. Detergent Test The system must contain a partial charge in order to detect leaks. The detergent test is accomplished by applying soap solution to an area suspected of leaking. Bubbles may form if leaks are present.

E. Electronic Detector Test An electronic detector includes a probe that is moved along the plumbing to detect escaping refrigerant. The probe should be held below the line because refrigerant is heavier than air. The probe should be capable of detecting leaks equal to 1/2 ounce per year and will emit a flashing light or a high-pitched sound when escaping refrigerant is detected.

F. Red Leak Detector Dye Additive NOTE: Do not inject red leak detector dye into a system containing Fluoro-Lite leak detector dye. R-12 systems (Airplanes without Kit 129-5020 Installed) It is permissible to add 1/4 ounce of leak detector dye (78, Table 1, Chapter 91-00-00) for every 10 ounces of oil into the low pressure service port. A red film will appear in the area where leaks are present. R-134a systems (Airplanes with Kit 129-5020 Installed) It is permissible to add 1/4 ounce of leak detector dye (132, Table 1, Chapter 91-00-00) for every 10 ounces of oil into the low pressure service port. A red film will appear in the area where leaks are present.

G. Yellow/Green Leak Detector Dye Additive (Preferred) Check the quantity of oil contained in the see-thru oil sump and add oil as required to maintain a full sump. Manufacturer's recommendation for lubricating oil to be used in the air cycle machine is Exxon 2389 or any oil conforming to MIL-L-7808 (72, Table 1, Chapter 91-00-00). CAUTION: For installation of the Fluoro-Lite leak detector dye into a system containing some other type of dye already, perform the CLEANING THE VAPOR CYCLE SYSTEM procedures prior to installing the Fluoro-Lite dye. Other prepackaged forms of Fluoro-Lite dye are available. However, these other forms of dye may not be of the same concentration. Contact your supplier for the proper quantity to use.

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R-12 SYSTEMS (Airplanes without Kit 129-5020 Installed) A leak detector dye (130, Table 1, Chapter 91-00-00) may be added by either premixing with the oil (70, Table 1, Chapter 91-00-00) or directly inserting it slowly into the low pressure service port by an injector (2, Table 201, 21-52-00). Add 1/4 ounce of dye for every 10 ounces of oil in the system. For leak inspection, scan all fittings and components using a hand-held ultraviolet or UV/blue lamp (3, Table 201, 21-52-00). All exposed leaks will appear bright yellow/green. R-134a systems (Airplanes with Kit 129-5020 Installed) A leak detector dye (129, Table 1, Chapter 91-00-00) may be added by either premixing with the oil (117, Table 1, Chapter 91-00-00) or directly inserting it slowly into the low pressure service port by an injector (1, Table 201, 21-52-00). Add 1/4 ounce of dye for every 10 ounces of oil in the system. For leak inspection, scan all fittings and components using a hand-held ultraviolet or UV/blue lamp (3, Table 201, 21-52-00). All exposed leaks will appear bright yellow/green.

H. Vapor Cycle System Component Repair/Replacement If a component is being repaired or replaced that, when removed, causes loss of pressurization, complete the following Steps: (1) Depressurize the system as needed. Perform the DEPRESSURIZING THE VAPOR CYCLE SYSTEM procedure. (2) Repair/replace components as needed. NOTE: The receiver-dryer should be the last component connected. It should be connected last to ensure maximum protection of the vapor cycle system against moisture. (3) Install a new receiver-dryer (Ref. 21-52-04, RECEIVER-DRYER INSTALLATION). (4) Restore any recovered oil or an equal amount of fresh oil to the system. (5) Clean all exposed oil. Use dye cleaner/remover (131, Table 1, Chapter 91-00-00) for systems containing a leak detector dye additive. Scan with a UV lamp (3, Table 201, 21-52-00) to check for any remaining dye residue. (6) Perform the EVACUATING THE VAPOR CYCLE SYSTEM procedure. (7) Perform the CHARGING THE VAPOR CYCLE SYSTEM procedure.

I. Compressor Oil Check If a refrigerant leak or a compressor oil leak has occurred, if it is suspected that damage has occurred to the compressor, or if the vapor cycle system has been recharged, the oil level should be checked as follows: (1) Set the right engine speed to 65% N1 and set the temperature mode and control switches to operate the charged system at maximum cooling levels for at least 10 minutes. (2) Turn off the refrigerant system, shutdown the engine, and evacuate the system with a recycle/recovery servicing unit (this salvages the refrigerant for future use).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (3) When the pressure is between 0 and 5 psi, remove the oil fill plug from the compressor and check the oil level by inserting a rod into the compressor crankcase. It may be necessary to rotate the compressor pulley to ensure that the rod bottoms on the crankcase and not on the compressor shaft. The level should be between 3 and 3.5 inches below the top of the compressor case. (4) On airplanes without Kit 129-5020 installed, add compressor oil, (70, Table 1, Chapter 91-00-00), if necessary, to obtain the correct level. On airplanes with Kit 129-5020 installed, add refrigerant oil (117, Table 1, Chapter 91-00-00) if necessary to obtain the correct level. New compressors are provided with four ounces of oil (117, Table 1, Chapter 91-00-00) and are charged with a mixture of refrigerant (116, Table 1, Chapter 91-00-00) and dry nitrogen to provide an internal pressure that is slightly above atmospheric pressure. For airplanes without Kit 129-5020 installed, the oil must be drained from the new compressor and replaced with compressor oil (70, Table Chapter 91-00-00) and recharged with refrigerant (71, Table 1, Chapter 91-00-00). For airplanes with Kit 129-5020 installed, the new compressors do not require exchanging the oil and refrigerant. NOTE: If the compressor is replaced, drain the old compressor and measure the amount of oil collected. Add one ounce to this measurement to determine the amount of new oil to add to the new compressor after the oil is drained from the new compressor. This method maintains the same amount of oil in the system. Additional oil may be needed to replace oil lost through evacuation, leakage or component replacement.

J. Depressurizing the Vapor Cycle System (1) Connect a servicing unit that recovers and recycles the refrigerant to the service valves and open the low pressure valve (Ref. Figure 201). (2) When the pressure is depleted, open the high pressure valve and operate the vacuum pump to completely remove the refrigerant from the system. (3) After the system is depressurized, check the oil level in the compressor. Refer to COMPRESSOR OIL CHECK procedure.

K. Evacuating the Vapor Cycle System A service unit should be equipped with a vacuum pump capable of obtaining an absolute pressure of 125 microns or less and a pressure gage capable of indicating an absolute pressure of 125 microns or less. Evacuate the system as follows: (1) Depressurize the system with a recycle/recovery servicing unit. (2) Attach the suction service hose to the suction service valve. The service valves are located in a service box in the underside of the fuselage (Ref. Figure 201). (3) Start the vacuum pump, gradually opening the vacuum valve until it is fully open. (4) Loosen the ballast valve on the vacuum pump slightly until the pressure is approximately 200 microns, then close the valve. (5) The system should be evacuated to 125 microns or less. This pump-down procedure should require no more than four hours if there are no leaks in the system.

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L. Cleaning the Vapor Cycle System (1) Remove and discard the receiver-dryer (Ref. 21-52-04, RECEIVER-DRYER REMOVAL). (2) Remove the compressor (Ref. 21-52-02, COMPRESSOR REMOVAL). NOTE: If the compressor is damaged, discard it. If the compressor is being reused, set it aside to drain the oil while the plumbing and coils are being flushed. If the system has been contaminated by a damaged compressor or receiver-dryer, the coils may need to be removed and flushed individually. CAUTION: Do not use solvent. It will cause contamination of the refrigerant oil. (3) Remove the two expansion valves. The hot gas bypass valves must be removed or electrically held open during flushing. (4) Flush the system using air conditioning system cleaner (77, Table 1, Chapter 91-00-00) and dry nitrogen on airplanes without Kit 129-5020 installed. Use AC Flush Fluid (118, Table 1, Chapter 91-00-00) and dry nitrogen to clean the refrigerant system on airplanes with Kit 129-5020 installed. (5) Restore both expansion valves and install or close the hot gas bypass valves. (6) Install the compressor (Ref. 21-52-02, COMPRESSOR INSTALLATION). If the existing compressor is being installed, insert 4 ounces of the appropriate oil for your system. New compressors already contain 4 ounces of oil in them. (7) Install a new receiver-dryer (Ref. 21-52-04, RECEIVER-DRYER INSTALLATION). (8) Insert 27 ounces of oil into the high pressure plumbing, preferably at the nacelle firewall. (9) Perform the EVACUATING THE VAPOR CYCLE SYSTEM procedure. (10) Perform the CHARGING THE VAPOR CYCLE SYSTEM procedure.

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Figure 201 Refrigerant Service Valves and Reset Switches

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M. Charging the Vapor Cycle System (Airplanes without Kit 129-5020) CAUTION: Use only R-12 refrigerant (71, Table 1, Chapter 91-00-00) to charge the air-conditioning system. The system should be charged when: (1) There are bubbles in the refrigerant as seen through the sight glass. If the OAT is above 75°F, the presence of bubbles in the refrigerant indicates that the refrigerant quantity is low. NOTE: The presence of bubbles if the OAT is below 75° is not an indication that the refrigerant quantity is low. Refer to the pressure graph as an indication of proper charge (Ref. Figure 202). (2) Leaks have been detected in the system. (3) Air has entered the system. (4) Components that carry refrigerant have been replaced. The vapor cycle system should be serviced by a qualified air conditioning service mechanic. Use only R-12 air conditioning refrigerant (71, Table 1, Chapter 91-00-00). Refrigerant service valves are located in the underside of the fuselage in a service box that is in line with the flaps. It is recommended that the service unit used for charging be equipped with a supply cylinder heated to maintain sufficient pressure to force refrigerant into the system without operating the compressor. If a heated cylinder is not available, the system may be charged without operating the compressor by allowing refrigerant to flow into the system until equilibrium pressure is reached. After equilibrium pressure is reached, allow the compressor to engage with the temperature mode switch set to MAN and the RH engine operating at approximately 65 percent N1. Add refrigerant vapor to the suction side of the system. Refrigerant should be added in vapor form to prevent compressor damage. Operate the compressor until no bubbles are seen in the sight glass then add approximately 8 ounces of additional refrigerant. Add the refrigerant to the suction port of the system until the pressures approach the pressures shown on the pressure graph (Ref. Figure 202). The total capacity of the vapor cycle system is 85 ounces of R-12 refrigerant. The system requires a total oil charge of 31 ounces (70, Table 1, Chapter 91-00-00).

N. Charging the Vapor Cycle System (Airplanes with Kit 129-5020) CAUTION: Use only R-134a refrigerant (116, Table 1, Chapter 91-00-00) to charge the air-conditioning system. The system should be charged when: (1) A reduction in cooling system ability and a decline in system pressures indicates that the refrigerant quantity is low. (2) Leaks have been detected in the system. (3) Air has entered the system. (4) Components that carry refrigerant have been replaced.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL The vapor cycle system should be serviced by a qualified air conditioning service mechanic. Refrigerant service valves are located in the underside of the fuselage in a service box that is in line with the flaps. It is recommended that the service unit used for charging be equipped with a supply cylinder heated to maintain sufficient pressure to force refrigerant into the system without operating the compressor. If a heated cylinder is not available, the system may be charged without operating the compressor by allowing refrigerant to flow into the system until equilibrium pressure is reached. After equilibrium pressure is reached, allow the compressor to engage with the temperature mode switch set to MAN and the RH engine operating at approximately 65% N1. Add refrigerant vapor to the suction side of the system. Refrigerant should be added in vapor form to prevent compressor damage. Add the refrigerant to the suction port of the system until the pressures approach the pressures shown on the pressure graph (Ref. Figure 203). The total capacity of the vapor cycle system is 75 ounces of refrigerant R-134a. The system requires a total oil charge of 31 ounces (117, Table 1, Chapter 91-00-00).

O. Vapor Cycle System Operational Check NOTE: Operating the air-conditioning system is not possible when the outside air temperature is below 50°F. The system may be serviced when the outside air temperature is below 50°F, but must be disabled by pulling the compressor clutch circuit breaker (CB206), placing a placard near the cabin temperature controls, and making an appropriate logbook entry. Once the outside air temperature is above 50°F, the operational check may be performed and the system returned to service. This procedure can be used to evaluate the operational efficiency of the vapor cycle system. This check should be performed only when the outside air temperature is above 50°F to insure that the 40°F OAT limit switch is closed. (1) Place thermometers, known to be accurate, in a forward cool air outlet and in an aft cool air outlet. (2) Start the right engine and increase the engine rpm to 65% N1. Do not turn the environmental bleed air on. (3) Place the cabin temperature mode switch in MAN and hold the manual temperature control switch in DECR for 60 seconds. Place the blower switch in HI. (4) Take a cabin air temperature reading near the floor of the cabin, and after the temperature of the air coming out of the cool air outlets has stabilized (5 to 10 minutes should be sufficient), record the temperatures of the cool air outlet air. (5) Subtract each of the temperature readings at the cool air outlets from the cabin air temperature recorded earlier. A differential of at least 20°F at both cool air outlets would be indicative of a properly operating vapor cycle system.

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Figure 202 R-12 Vapor Cycle System Suction and Discharge Pressures Versus Ambient Temperatures Page 210 Nov 1/13

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Figure 203 R-134a Vapor Cycle System Suction and Discharge Pressures Versus Ambient Temperatures

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21-52-01 200200

1. PROCEDURES A. Filter Replacement The filters on the evaporators should be replaced at the interval specified in Chapter 5-20-00 or more frequently if operating in heavy smoke or dusty conditions. Two springs installed on the aft end of each evaporator secure the filters (Ref. Figure 201). The evaporators are installed under the center aisle floorboards in line with the first and sixth cabin windows.

2. FORWARD EVAPORATOR A. Removal (1) Open the access door to the service box (located in the fuselage underside) and discharge the vapor cycle system slowly with a recycle/recovery unit until all pressure is bled off (Ref. 21-52-00). (2) Remove the carpet and floorboards in the center aisle in line with the first cabin window. (3) Remove the attaching parts securing the clamps on the hot gas bypass line and the inlet line to the evaporator assembly (Ref. Figure 201). (4) Disconnect the inlet line, outlet line, and bypass line from the evaporator fittings. (5) Remove insulation and tape wrapping around thermal bulb (of expansion valve) clamped on outlet line (suction line). Remove clamp and the thermal bulb from outlet line (suction line). (6) Remove the attaching parts securing the aft end of the evaporator to the mounting clips. (7) Loosen the clamp securing the evaporator to the blower. (8) Remove the evaporator assembly.

B. Installation (1) Install the evaporator and secure the aft end with the attaching parts. Ensure that the drain valve aligns with the gasket on the fuselage skin (Ref. Figure 201). (2) Tighten the clamp to secure the evaporator to the blower. (3) Connect the inlet line, expansion valve, the outlet line and the bypass line to the evaporator fittings. (4) Install the attaching parts to secure the clamp on the bypass line to the evaporator plenum assembly. (5) The expansion valve has a capillary tube with one end coiled into a bulb. This bulb must be attached to the outlet line (suction line) that connects to the evaporator outlet fitting. The capillary tube (including the coiled tube bulb) may be 19 to 32 inches in length. The extra length of the capillary tubing should be coiled using large bend radius to prevent crimping. Use ties or cushioned clamps to support capillary tubing. The bulb must be in direct contact with the surface of

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL the outlet line (suction line). Clean the outlet line (suction line) outer surface before clamping the bulb in place. The bulb may be installed on top of the outlet line (suction line) or side mounted (preferably at the 3 o’clock position). Never mount a bulb on the bottom of the outlet line (suction line) because a mixture of refrigerant and oil may be present at that point. Clamp bulb on outlet line (suction line). Wrap insulation and tape around the bulb and outlet line (suction line). (6) Install the floorboards and the carpet in the center aisle. (7) Recharge the system (Ref. 21-52-00, CHARGING THE VAPOR CYCLE SYSTEM). (8) Install the access door on the service box (located in the fuselage underside).

3. AFT EVAPORATOR A. Removal (1) Open the access door to the service box (located in the fuselage underside) and discharge the vapor cycle system slowly with a recycle/recovery unit until all pressure is bled off (Ref. 21-52-00). (2) Remove the carpet and floorboards in the center aisle in line with the sixth cabin window. (3) Remove the clamp securing the hot gas bypass line to the evaporator assembly (Ref. Figure 201). (4) Disconnect the inlet line, outlet line and bypass line from the evaporator fittings. (5) Remove insulation and tape wrapping around thermal bulb (of expansion valve) clamped on outlet line (suction line). Remove clamp and the thermal bulb from the outlet line (suction line). (6) Remove the attaching parts securing the aft end of the evaporator to the mounting clips. (7) Loosen the clamps securing the evaporator to the blower and remove the evaporator assembly.

B. Installation (1) Install the evaporator and secure the aft end with the attaching parts. Ensure that the drain valve aligns with the gasket on the fuselage skin (Ref. Figure 201). (2) Secure the evaporator to the blower by tightening the clamp on the evaporator. (3) Connect the inlet line, outlet line and bypass line to the evaporator fittings. (4) Secure the clamp on the bypass line to the evaporator plenum assembly. (5) The expansion valve has a capillary tube with one end coiled into a bulb. This bulb must be attached to the outlet line (suction line) that connects to the evaporator outlet fitting. The capillary tube (including the coiled tube bulb) may be 19 to 32 inches in length. The extra length of the capillary tubing should be coiled using large bend radius to prevent crimping. Use ties or cushioned clamps to support capillary tubing. The bulb must be in direct contact with the surface of the outlet line (suction line). Clean the outlet line (suction line) outer surface before clamping the bulb in place. The bulb may be installed on top of the outlet line (suction line) or side mounted (preferably at the 3 o’clock position). Never mount a bulb on the bottom of the outlet line (suction line) because a mixture of refrigerant and oil may be present at that point. Clamp bulb on outlet line (suction line). Wrap insulation and tape around the bulb and outlet line (suction line).

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(6) Install the floorboards and the carpet in the center aisle. (7) Recharge the system (Ref. 21-52-00, CHARGING THE VAPOR CYCLE SYSTEM). (8) Install the access door to the service box (located in the fuselage underside).

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Figure 201 Evaporator Installation

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

21-52-02 200200

ENVIRONMENTAL SYSTEMS COMPRESSOR (UA-1 AND AFTER; UB-1 AND AFTER; UC-1 THRU UC-100 NOT MODIFIED BY SERVICE BULLETIN NO. 2345) MAINTENANCE PRACTICES 1. COMPRESSOR A. Removal NOTE: The receiver-dryer should be replaced if the vapor cycle system has been opened. The receiver-dryer should be connected last to ensure maximum protection of the vapor cycle system against moisture (Ref. 21-52-00). (1) Perform the DEPRESSURIZING THE VAPOR CYCLE SYSTEM procedure (Ref. 21-52-00). (2) Perform the EVACUATING THE VAPOR CYCLE SYSTEM procedure (Ref. 21-52-00). (3) Remove the right engine cowling panels as required (Ref. Chapter 71-10-00, COWLING REMOVAL). (4) Perform REMOVING GROUND POWER procedures (Ref. Chapter 24-40-00). (5) Perform BATTERY DISCONNECTION procedure (Ref. Chapter 24-31-00). (6) Attach a red tag to the battery switch and external power switch with the words “Do Not Operate, Maintenance In Progress”. NOTE: The power wire (66) is hard wired to the compressor clutch and is spliced (67) to the airplane wiring (Ref. Figure 201). (7) Disconnect the electrical power wire (66) at the splice (67). (8) Remove screw (69) and clip (70) and disconnect the ground wire (68) from the compressor (43). (9) Disconnect and cap the discharge hose (47) and the suction hose (52). Cap fittings (72 and 73) on the compressor (43). (10) Cut the safety wire and loosen the compressor pivot bolts (25 and 36) and the belt tension turnbuckle (57) lower attachment bolt (55) and nut (37). (11) Cut the safety wire and loosen both jamnuts (56). Adjust the tension turnbuckle (57) as required to relieve belt tension. (12) Remove the lower attachment bolt (55), washers (38 and 54) and nut (37) from the tension turnbuckle (57) lower rod end (53). (13) Remove the drive belt (19) from the compressor (43). NOTE: Note the position and number of shims (28 and 31) to aid later installation. (14) Remove the compressor pivot bolts (25 and 36), washers (26 and 35) and laminated shims (28 and 31). Remove the compressor (43).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (15) Remove bolts (49 and 50), washers (34, 42, 48 and 51) and nuts (33 and 41) from the forward mount plate. Remove the compressor forward mount plate (32). (16) Remove the bolts (45), washers (40 and 44) and nuts (39). Remove the compressor aft mount plate (27).

B. Installation NOTE: The receiver-dryer should be replaced if the vapor cycle system has been opened. The receiver-dryer should be connected last to ensure maximum protection of the vapor cycle system against moisture (Ref. 21-52-00). (1) Install the compressor aft mount plate (27) by installing bolts (45), washers (40 and 44) and nuts (39) and torque 280 to 300 inch-pound (Ref. Figure 201). (2) Install the compressor forward mount plate (32) by installing bolts (49 and 50), washers (34, 42, 48 and 51) and nuts (33 and 41). Install with bolt heads facing forward and torque bolts (49 and 50) 280 to 300 inch-pounds. (3) Install the compressor by installing pivot bolts (25 and 36), washers (26 and 35) and laminated shims (28 and 31). (a) Install shim (27) and remove laminated layers as required to achieve 1.480 ± 0.030 inches from the forward edge of the clutch pulley (20) to the aft edge of the compressor mount assembly (22) (Ref. Figure 202). (b) Install shim (23) and remove laminated layers as required to fill in the gap. (c) Torque bolts (24) from 180 to 200 inch-pounds, then verify the 1.480 ± 0.030 inches dimension, repeat Step (3) if necessary. (4) Perform the COMPRESSOR BELT INSTALLATION procedure and install the drive belt (19) on the compressor (43) (Ref. Figure 201). (5) Attach the tension turnbuckle (57) lower rod end (53) by installing the lower attachment bolt (55), washers (38 and 54) and nut (37). (6) Perform the COMPRESSOR BELT TENSION procedure. (7) Ensure jamnuts (56) are safety wired. (8) Torque attachment bolt (55) and nut (37) 280 to 300 inch-pounds. (9) Torque the pivot bolts (25 and 36) 180 to 200 inch-pounds. (10) Ensure all electrical power to the airplane is removed, the airplane battery is disconnected and a red tag is attached to the battery switch and external power switch with the words “Do Not Operate, Maintenance In Progress”. NOTE: The power wire (66) is hard wired to the compressor clutch (71) and is spliced (67) to the airplane wiring. (11) Splice the electrical power wire (66) to the airplane wiring. (12) Connect the ground wire (68) to the compressor (43) by installing clip (70) and screw (69).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL CAUTION: Exercise care to ensure proper clearance between suction and discharge hoses at the engine truss area. (13) Remove caps from fittings (72 and 73) on the compressor (43). Remove plugs and connect the discharge hose (47) and the suction hose (52). (14) Perform BATTERY CONNECTION procedure (Ref. Chapter 24-31-00) and remove red tags from the battery switch and external power switch. (15) Replace the receiver dryer after the system has been opened for component replacement (Ref. 21-52-04). (16) Perform the EVACUATING THE VAPOR CYCLE SYSTEM procedure (Ref. 21-52-00). (17) Perform the CHARGING THE VAPOR CYCLE SYSTEM procedure (Ref. 21-52-00). (18) Install cowling panels (Ref. Chapter 71-10-00, COWLING INSTALLATION). (19) Perform the VAPOR CYCLE SYSTEM OPERATIONAL CHECK procedure (Ref. 21-52-00).

2. COMPRESSOR MOUNT A. Removal (1) Remove the right engine cowling panels as required (Ref. Chapter 71-10-00, COWLING REMOVAL). (2) Perform the STARTER-GENERATOR REMOVAL procedure (Ref. Chapter 24-30-01). (3) Perform the COMPRESSOR BELT REMOVAL procedure. (4) Perform the COMPRESSOR REMOVAL procedure. (5) Remove the bearings (17 and 20) and drive pulley (18) assembly (Ref. Figure 201). (6) Remove clamp (5) and disconnect drain tube (6). NOTE: Inspect bolt (21) for wear, replace if required. (7) Remove cotter key (61), nut (60), bolt (21) and washers (22 and 59), discard cotter key (61). Remove the belt tension turnbuckle (57) assembly from mount (30). (8) Cut the safety wire and remove five bolts and washers (9) attaching the compressor mount (30) to the accessory gearbox pads. (9) Cut the safety wire and remove bolts (8 and 23) and washers (7 and 24) that attach the compressor mount (30) to the compressor mount support (65). NOTE: Be careful not to lose the spring (1) when removing mount (30) and the quill shaft (2). (10) Remove the compressor mount (30) and discard the packings (3). (11) Remove the quill shaft (2) and spring (1).

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B. Installation (1) Ensure the support mounting surface (4) is flush with the accessory gearbox mounting pad (1) within 0.002 inch. If adjustment is required perform the COMPRESSOR MOUNT SUPPORT REMOVAL and COMPRESSOR MOUNT SUPPORT INSTALLATION procedures (Ref. Figure 202). (2) Install the spring (1) into the engine accessory gearbox gear shaft (Ref. Figure 201). (3) Install the quill shaft (2) into the compressor mount (30). (4) Install compressor mount (30) with two new packings (3), onto the engine accessory gear box and the compressor mount bracket (65). (5) Install five bolts and washers (9) attaching the compressor mount (30) to the accessory gearbox pads. (6) Install the two bolts (8 and 23) and washers (7 and 24) that attach the compressor mount (30) to the compressor mount support (65). (7) Torque bolts (8, 9 and 23) 40 to 50 inch-pounds. (8) Install MS20995C32 safety wire on all bolts. (9) Connect drain tube (6) and install clamp (5). NOTE: Inspect bolt (21) for wear, replace if required. (10) Install the belt tension turnbuckle (57) assembly on mount (30) by installing bolt (21), washers (22 and 59), nut (60) and new cotter key (61). (11) If bearings (17 and 20) were removed from the drive pulley (18) shaft, apply a thin coat of retaining compound (194, Table 1, 91-00-00) and press the bearings onto the shaft. (12) Install the bearings (17 and 20) and drive pulley (18) assembly. (13) Perform the STARTER-GENERATOR INSTALLATION procedure (Ref. Chapter 24-30-01). (14) Perform the COMPRESSOR BELT INSTALLATION procedure. (15) Perform the COMPRESSOR INSTALLATION procedure. (16) Install cowling panels as required (Ref. Chapter 71-10-00, COWLING INSTALLATION). (17) Perform the VAPOR CYCLE SYSTEM OPERATIONAL CHECK procedure (Ref. 21-52-00).

3. COMPRESSOR MOUNT SUPPORT A. Removal (1) Remove the right engine cowling panels as required (Ref. Chapter 71-10-00, COWLING REMOVAL). (2) Perform the COMPRESSOR BELT REMOVAL procedure.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (3) Perform the COMPRESSOR REMOVAL procedure. (4) Perform the COMPRESSOR MOUNT REMOVAL procedure. NOTE: Note the position and thickness of laminated shims (64) during removal (Ref. Figure 201). (5) Remove nuts (62) and washers (63), discard nuts. (6) Remove the compressor mount support (65) and shims (64), retain the shims.

B. Installation NOTE: Use shims (64) to ensure the support (65) mounting surface is flush with the accessory gearbox mounting pad within 0.002 inch when nuts (62) are torqued (Ref. Figure 201). (1) Install shims (64) and compressor mount support (65). (2) Install washers (63) and new nuts (62). (3) Torque nuts (62) 32 to 36 inch-pounds reference, Pratt and Whitney Maintenance Manual P/N 3041195 for nut torque value. (4) Check that support (2) mounting surface (4) is flush with the accessory gearbox mounting pad (1) within 0.002 inch (Ref. Figure 202). If no adjustment is required proceed to Step (5). If adjustment is required, perform the following Steps: (a) Remove nuts (62) and washers (63) (Ref. Figure 201). (b) Remove the compressor mount support (65) and add or remove shims (64) or laminations to obtain the 0.0 to 0.002 inch dimension. (c) Install the compressor mount support (65), washers (63) and nuts (62). Torque nuts (62) 32 to 36 inch-pounds. (5) Perform the COMPRESSOR MOUNT INSTALLATION procedure. (6) Perform the COMPRESSOR BELT INSTALLATION procedure. (7) Perform the COMPRESSOR INSTALLATION procedure. (8) Install cowling panels (Ref. Chapter 71-10-00, COWLING INSTALLATION). (9) Perform the VAPOR CYCLE SYSTEM OPERATIONAL CHECK procedure (Ref. 21-52-00).

4. COMPRESSOR BELT A. Removal (1) Remove the right engine cowling panels as required (Ref. Chapter 71-10-00, COWLING REMOVAL). (2) Cut safety wire and loosen the compressor pivot bolts (25 and 36) and the belt tension turnbuckle (57) lower attachment bolt (55) and nut (37) (Ref. Figure 201).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (3) Cut the safety wire and loosen both jamnuts (56). Adjust the tension turnbuckle (57) as required to relieve belt tension. (4) Remove the lower attachment bolt (55), washers (38 and 54) and nut (37) from the tension turnbuckle (57) lower rod end (53). (5) Cut the safety wire and remove three socket head cap screws (13) and washers (12) from the belt housing (11). (6) Remove the belt housing (11), laminated shim(s) (16), spacer (15) and snap ring (14). CAUTION: Do not let the compressor (43) drop after the belt (19) is removed. (7) Remove the drive belt (19).

B. Installation (1) Install the drive belt (19) on drive pulley (18) (Ref. Figure 201). (2) Install the drive belt housing (11), using three socket head cap screws (13) and washers (12). Torque the three socket head cap screws 25 to 30 inch-pounds and install MS20995C32 safety wire. (3) Insert the laminated shim(s) (16) into the belt housing (11). (4) Insert the spacer (15) into the belt housing (11). (5) Install the snap ring (14) and check for a 0.010 ± 0.002 gap between the snap ring (14) and the spacer (15). If gap is not correct, remove the snap ring (14) and the spacer (15) and add or remove laminated shim(s) (16) or laminations, install the snap ring (14) and the spacer (15), check gap and repeat until the correct gap is obtained. (6) If the compressor is not installed, perform the COMPRESSOR INSTALLATION procedure. (7) Install the drive belt (19) on the clutch pulley (71). (8) Perform the COMPRESSOR BELT TENSION procedure. (9) Install cowling panels (Ref. Chapter 71-10-00, COWLING INSTALLATION). (10) Perform the VAPOR CYCLE SYSTEM OPERATIONAL CHECK procedure (Ref. 21-52-00).

C. Tension NOTE: The tension of a new compressor belt should be checked after 50 hours of operation. (1) Remove the right engine cowling panels as required (Ref. Chapter 71-10-00, COWLING REMOVAL). (2) Using a push-pull gauge (6, Chart 1, Chapter 27-00-00) apply a force of 4.7 to 5.9 pounds at the midspan of the belt. The belt should deflect 0.15-inch. (3) If no adjustment is required proceed to Step (4). If adjustment is required, perform the following Steps:

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL NOTE: The maximum length of rod end (28) that is allowed to extend out of the tension turnbuckle (30) is 1.50 inches at each rod end (28). When these maximum dimensions are reached, the belt should be replaced (Ref. Figure 202). (a) Loosen the belt tension turnbuckle (57) lower attachment bolt (55) and nut (37) (Ref. Figure 201). (b) Cut safety wire and loosen the compressor pivot bolts (25 and 36). (c) Cut the safety wire and loosen both jamnuts (56). Adjust the tension turnbuckle (57) as required to obtain the belt tension required in Step (2). (d) Tighten both jamnuts (56) and install MS20995C32 safety wire. (e) Torque the belt tension turnbuckle (57) lower attachment bolt (55) and nut (37) 280 to 300 inch-pounds. (f) Torque the compressor pivot bolts (25 and 36) 180 to 200 inch-pounds and install MS20995C32 safety wire. (4) Install cowling panels as required (Ref. Chapter 71-10-00, COWLING INSTALLATION). (5) Perform the VAPOR CYCLE SYSTEM OPERATIONAL CHECK procedure (Ref. 21-52-00).

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1 2 5

3 1. SPRING 2. QUILL SHAFT 3. PACKING (2 PLACES) 4. TUBE END 5. CLAMP 6. DRAIN TUBE 7. WASHER 8. BOLT 9. BOLT, WASHER (5 PLACES) 10. PIN 11. BELT HOUSING 12. WASHER (3 PLACES) 13. SOCKET HEAD CAP SCREWS (3 PLACES) 14. SNAP RING 15. SPACER 16. LAMINATED SHIM 17. BEARING 18. DRIVE PULLEY 19. BELT 20. BEARING 21. BOLT 22. WASHER 23. BOLT 24. WASHER 25. BOLT 26. WASHER 27. AFT MOUNT PLATE 28. LAMINATED SHIM 29. INSERT 30. COMPRESSOR MOUNT 31. LAMINATED SHIM 32. FORWARD MOUNT PLATE 33. NUT 34. WASHER 35. WASHER 36. BOLT 37. NUT

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38. WASHER 39. NUT 40. WASHER 41. NUT 42. WASHER 43. COMPRESSOR 44. WASHER 45. BOLT 46. PRESSURE RELIEF VALVE 47. HIGH PRESSURE DISCHARGE HOSE ASSEMBLY 48. WASHER 49. BOLT 50. BOLT 51. WASHER 52. SUCTION HOSE ASSEMBLY 53. ROD END 54. WASHER 55. BOLT 56. JAMNUTS 57. TURNBUCKLE 58. ROD END 59. WASHER 60. NUT 61. COTTER PIN 62. NUTS 63. WASHERS 64. LAMINATED SHIMS 65. COMPRESSOR MOUNT SUPPORT 66. POWER WIRE 67. SPLICE 68. GROUND WIRE 69. SCREW 70. CLIP 71. CLUTCH PULLEY 72. SUCTION FITTING 73. DISCHARGE FITTING

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UA21B 052943AA.AI

Figure 201 Vapor Cycle Compressor Installation (UA-1 and After; UB-1 and After; UC-1 thru UC-100 not Modified by Service Bulletin NO. 2345)

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1. ACCESSORY GEARBOX MOUNTING PAD 2. COMPRESSOR MOUNT SUPPORT 3. SCALE 4. COMPRESSOR MOUNT SUPPORT MOUNTING SURFACES 5. BELT 6. SUPPORT (REF) 7. RETAINING RING 8. SPACER 9. LAMINATED SHIM 10. DRIVE PULLEY 11. ALIGNMENT PIN 12. MOUNT ASSEMBLY 13. BEARING (TYPICAL 2 PLACES) 14. QUILL SHAFT 15. PACKING 16. SPRING 17. GEAR SHAFT 18. CLAMP 19. DRAIN TUBE 20. CLUTCH PULLEY 21. SUPPORT 22. MOUNT ASSEMBLY 23. LAMINATED SHIM 24. BOLT, WASHER (TYPICAL 2 PLACES) 25. PLATE 26. BOLT, WASHER (2 EACH REQUIRED) (TYPICAL 2 PLACES) 27. LAMINATED SHIM 28. ROD END 29. JAMNUT 30. TURNBUCKLE

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UA-1 AND AFTER; UB-1 AND AFTER; UC-1 AND AFTER NOT MODIFIED BY SERVICE BULLETIN NO. 2345

UA21B 052910AA.AI

Figure 202 Vapor Cycle Compressor Alignment (UA-1 and After; UB-1 and After; UC-1 thru UC-100 not Modified by Service Bulletin NO. 2345)

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

ENVIRONMENTAL SYSTEMS CONDENSER AND BLOWER MAINTENANCE PRACTICES

21-52-03 200200

1. PROCEDURES A. Condenser Removal (1) Perform the DEPRESSURIZING THE VAPOR CYCLE SYSTEM procedure (Ref. 21-52-00). (2) Perform the EVACUATING THE VAPOR CYCLE SYSTEM procedure (Ref. 21-52-00). (3) Remove the access panel on top surface of the right-hand inboard wing to gain access to the condenser. (4) Remove the inlet line and the outlet line from the condenser fittings (Ref. Figure 201). (5) Remove the attaching parts and the condenser from the compartment.

B. Condenser Installation (1) Install the condenser and secure it with the attaching parts (Ref. Figure 201). (2) Connect the inlet line and the outlet line to the condenser. (3) Install the access panel on the top surface of the right-hand inboard wing. (4) Perform the EVACUATING THE VAPOR CYCLE SYSTEM procedure (Ref. 21-52-00). (5) Recharge the system (Ref. 21-52-00, CHARGING THE VAPOR CYCLE SYSTEM). (6) Install the access door on the service box, located in the underside of the fuselage. (7) Perform the VAPOR CYCLE SYSTEM OPERATIONAL CHECK procedure (Ref. 21-52-00).

C. Condenser Blower Removal (1) Remove the access panel on the top surface of the right-hand inboard wing. (2) Remove the electrical leads from the terminals on the blower (Ref. Figure 201). (3) Remove the attaching parts securing the blower to the mounting bracket and remove the blower.

D. Condenser Blower Installation (1) Install the blower on the mounting bracket and secure with the attaching parts (Ref. Figure 201). (2) Connect the electrical leads to the terminals on the blower. (3) Install the access panel on the top surface of the right-hand inboard wing. (4) Perform the VAPOR CYCLE SYSTEM OPERATIONAL CHECK procedure (Ref. 21-52-00).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 201 Condenser and Blower Installation

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

ENVIRONMENTAL SYSTEMS RECEIVER - DRYER MAINTENANCE PRACTICES

21-52-04 200200

1. PROCEDURES A. Removal (1) Perform the DEPRESSURIZING THE VAPOR CYCLE SYSTEM procedure (Ref. 21-52-00). (2) Perform the EVACUATING THE VAPOR CYCLE SYSTEM procedure (Ref. 21-52-00). (3) Remove the carpet and floorboards in the center aisle in line with the third cabin window. (4) Disconnect the inlet line and the outlet line from the receiver-dryer (Ref. Figure 201). (5) Remove the attaching parts on the clamps securing the receiver-dryer to its mounting bracket and remove the receiver-dryer.

B. Installation (1) Install the receiver-dryer and secure it with the attaching parts (Ref. Figure 201). CAUTION: Be certain that the receiver-dryer being installed is compatible with the refrigerant that is being used. (2) Connect the inlet line and the outlet line to the receiver-dryer. (3) Perform the EVACUATING THE VAPOR CYCLE SYSTEM procedure (Ref. 21-52-00). (4) Recharge the system (Ref. 21-52-00, CHARGING THE VAPOR CYCLE SYSTEM). (5) Install the floorboards and carpet in the center aisle. (6) Install the access door on the service box (located on the underside of the fuselage). (7) Perform the VAPOR CYCLE SYSTEM OPERATIONAL CHECK procedure (Ref. 21-52-00).

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Figure 201 Receiver-Dryer Installation

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

ENVIRONMENTAL SYSTEMS COMPRESSOR (UA-1 AND AFTER; UB-1 AND AFTER; UC-1 AND AFTER MODIFIED BY SERVICE BULLETIN NO. 2345) MAINTENANCE PRACTICES

21-52-05 200200

1. COMPRESSOR A. Removal NOTE: The receiver-dryer should be replaced if the vapor cycle system has been opened. The receiver-dryer should be connected last to ensure maximum protection of the vapor cycle system against moisture (Ref. 21-52-00). (1) Perform the DEPRESSURIZING THE VAPOR CYCLE SYSTEM procedure (Ref. 21-52-00). (2) Perform the EVACUATING THE VAPOR CYCLE SYSTEM procedure (Ref. 21-52-00). (3) Remove the right engine cowling panels as required (Ref. Chapter 71-10-00, COWLING REMOVAL). (4) Perform REMOVING GROUND POWER procedures (Ref. Chapter 24-40-00). (5) Perform BATTERY DISCONNECTION procedure (Ref. Chapter 24-31-00). (6) Attach a red tag to the battery switch and external power switch with the words “Do Not Operate, Maintenance In Progress”. NOTE: The power wire (57) is hard wired to the compressor clutch and is spliced (58) to the airplane wiring (Ref. Figure 201). . (7) Disconnect the electrical power wire (57) at the splice (58). (8) Remove screw (60) and clip (61) and disconnect the ground wire (59) from the compressor clutch. (9) Disconnect and plug the discharge hose (55) and the suction hose (56). Cap fittings (62 and 63) on the compressor (64). (10) Loosen the compressor pivot bolts (29 and 30) and the belt tension turnbuckle (43) lower attachment bolt (44) and nut (3). (11) Cut the safety wire and loosen both jamnuts (42). Adjust the tension turnbuckle (43) as required to relieve belt tension. (12) Remove the lower attachment bolt (44), washers (45 and 46) and nut (3) from the tension turnbuckle (43) lower rod end (47). (13) Remove the drive belt (26) from the compressor (64). NOTE: Laminated shims (5), if installed, should be located between the compressor mount plates (8 and 9) and the pivot plate (4) (Ref. Figure 202).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (14) Remove the compressor pivot bolts, washers, nuts (29 and 30) and shims (31), if installed. Remove the compressor (64) (Ref. Figure 201). (15) Cut and remove the safety wire. Remove bolts, nuts and washers (52 and 53) from the forward mount plate. Remove the compressor forward mount plate (51). (16) Cut and remove the safety wire. Remove the bolts, nuts and washers (48 and 50). Remove the compressor aft mount plate (49).

B. Installation NOTE: The receiver-dryer should be replaced if the vapor cycle system has been opened. The receiver-dryer should be connected last to ensure maximum protection of the vapor cycle system against moisture (Ref. 21-52-00). When installing the aft compressor mount plate (49) use one AN960-616 washer or one AN960-616L washer or any combination to obtain clearance between the bolts (48 and 50) and the clutch pulley (5) (Ref. Figure 201). (1) Install the compressor aft mount plate (49) by installing bolts, nuts and washers (48 and 50). Install MS20995C32 safety wire on bolts. (2) Install the compressor forward mount plate (51) by installing bolts, nut and washers (52 and 53). Install with bolt heads facing aft and install MS20995C32 safety wire on bolts. (3) Install the compressor (64) by installing pivot bolts, washers, nuts (29 and 30) and previously installed shims (31). Ensure the shims (31) are between the forward (51) and aft (49) mount plates and the pivot plate (32). (4) Perform the COMPRESSOR BELT INSTALLATION procedure in this section and install the drive belt (26) on the compressor (64). (5) Attach the tension turnbuckle (43) lower rod end (47) by installing the lower attachment bolt (44), washers (45 and 46) and nut (3). (6) Perform the COMPRESSOR BELT ALIGNMENT procedure in this section. (7) Perform the COMPRESSOR BELT TENSION procedure in this section. (8) Ensure jamnuts (42) are safety wired. (9) Torque attachment bolt (44) and nut (3) 60 to 85 inch-pounds (Ref. Figure 201). (10) Ensure the pivot bolts (29 and 30) have been tightened. (11) Ensure all electrical power to the airplane is removed, the airplane battery is disconnected and a red tag is attached to the battery switch and external power switch with the words “Do Not Operate, Maintenance In Progress”. NOTE: The power wire (57) is hard wired to the compressor clutch and is spliced (58) to the airplane wiring. (12) Splice the electrical power wire (57) to the airplane wiring. (13) Connect the ground wire (59) to the compressor (64) by installing clip (61) and screw (60).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL CAUTION: Exercise care to ensure proper clearance between suction and discharge hoses at the engine truss area. (14) Remove caps from fittings (62 and 63) on the compressor (64). Remove caps and connect the discharge hose (55) and the suction hose (56). (15) Perform BATTERY CONNECTION procedure (Ref. Chapter 24-31-00) and remove red tags from the battery switch and external power switch. (16) Replace the receiver dryer after the system has been opened for component replacement (Ref. 21-52-04). (17) Perform the EVACUATING THE VAPOR CYCLE SYSTEM procedure (Ref. 21-52-00). (18) Perform the CHARGING THE VAPOR CYCLE SYSTEM procedure (Ref. 21-52-00). (19) Install cowling panels (Ref. Chapter 71-10-00, COWLING INSTALLATION). (20) Perform the VAPOR CYCLE SYSTEM OPERATIONAL CHECK procedure (Ref. 21-52-00).

2. COMPRESSOR MOUNT A. Removal (1) Remove the right engine cowling panels as required (Ref. Chapter 71-10-00, COWLING REMOVAL). (2) Perform the STARTER-GENERATOR REMOVAL procedure (Ref. Chapter 24-30-01). (3) Perform the COMPRESSOR BELT REMOVAL procedure in this section. (4) Perform the COMPRESSOR REMOVAL procedure in this section. (5) Remove the bearings (17 and 19) and drive pulley (18) assembly (Ref. Figure 201). (6) Remove the quill shaft (2) and spring (1). (7) Remove clamp (34) and disconnect drain tube (33). (8) Remove bolt (16), washers (15 and 11), nut (10), cotter key (9) and remove the belt tension turnbuckle (43) assembly from mount (12). (9) Cut the safety wire and remove six bolts and washers (38) attaching the compressor mount (12) to the accessory gearbox pads and two bolts and washers (37) that attach the compressor mount (12) to the compressor mount support (40). (10) Remove the compressor mount (12) and discard the packings (8).

B. Installation NOTE: Ensure the support mounting surface (15) is flush with the accessory gearbox mounting pad (13) within 0.002 inch. If adjustment is required perform the COMPRESSOR MOUNT SUPPORT REMOVAL and COMPRESSOR MOUNT SUPPORT INSTALLATION procedures in this section (Ref. Figure 202).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (1) Install compressor mount (12) with two new packings (8) (Ref. Figure 201). (2) Install six bolts and washers (38) attaching the compressor mount (12) to the accessory gearbox pads. (3) Install the two bolts and washers (37) that attach the compressor mount (12) to the compressor mount support (40) and torque 40 to 50 inch-pounds. (4) Install MS20995C32 safety wire on all bolts. (5) Connect drain tube (33) and install clamp (34). NOTE: Inspect bolt (16) for wear, replace if required. (6) Install the belt tension turnbuckle (43) assembly on mount (12) by installing bolt (16), washers (15 and 11), nut (10), cotter pin (9). (7) Install the spring (1) and quill shaft (2). (8) If bearings (17 and 19) were removed from the drive pulley (18) shaft, apply a thin coat of retaining compound (194, Table 1, 91-00-00) and press the bearings onto the shaft. (9) Install the bearings (17 and 19) and drive pulley (18) assembly. (10) Perform the COMPRESSOR BELT INSTALLATION procedure in this section. (11) Perform the COMPRESSOR INSTALLATION procedure in this section. (12) Perform the STARTER-GENERATOR INSTALLATION procedure (Ref. Chapter 24-30-01). (13) Install cowling panels as required (Ref. Chapter 71-10-00, COWLING INSTALLATION). (14) Perform the VAPOR CYCLE SYSTEM OPERATIONAL CHECK procedure (Ref. 21-52-00).

3. COMPRESSOR MOUNT SUPPORT A. Removal (1) Remove the right engine cowling panels as required (Ref. Chapter 71-10-00, COWLING REMOVAL). (2) Perform the COMPRESSOR BELT REMOVAL procedure in this section. (3) Perform the COMPRESSOR REMOVAL procedure in this section. (4) Perform the COMPRESSOR MOUNT REMOVAL procedure in this section. NOTE: Note the position and thickness of laminated shims (41) during removal (Ref. Figure 201). (5) Remove nuts (7) and washers (6), discard nuts. (6) Remove the compressor mount support (40) and shims (41), retain the shims.

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B. Installation NOTE: Use shims (41) to ensure the support (40) mounting surface is flush with the accessory gearbox mounting pad within 0.002 inch when nuts (7) are torqued (Ref. Figure 201). (1) Install shims (41) and compressor mount support (40). (2) Install washers (6) and new nuts (7). (3) Torque the nuts (7) 32 to 36 inch-pounds, see Pratt and Whitney Maintenance Manual P/N 3041195 for nut torque value. (4) Check that support mounting surface (15) is flush with the accessory gearbox mounting pad (13) within 0.002 inch (Ref. Figure 202). (5) If no adjustment is required proceed to Step (6). If adjustment is required, perform the following Steps: (a) Remove nuts (7) and washers (6) (Ref. Figure 201). (b) Remove the compressor mount support (40) and add or remove shims or laminations to obtain the 0.0 to 0.002 inch dimension. (c) Install the compressor mount support (40), washers (6) and nuts (7). Torque nuts (7) 32 to 36 inch-pounds. (6) Perform the COMPRESSOR MOUNT INSTALLATION procedure in this section. (7) Perform the COMPRESSOR BELT INSTALLATION procedure in this section. (8) Perform the COMPRESSOR INSTALLATION procedure in this section. (9) Install cowling panels (Ref. Chapter 71-10-00, COWLING INSTALLATION). (10) Perform the VAPOR CYCLE SYSTEM OPERATIONAL CHECK procedure (Ref. 21-52-00).

4. COMPRESSOR BELT A. Removal (1) Remove the right engine cowling panels as required (Ref. Chapter 71-10-00, COWLING REMOVAL). (2) Loosen the compressor pivot bolts (29 and 30) and the belt tension turnbuckle (43) lower attachment bolt (44) and nut (3) (Ref. Figure 201). (3) Cut the safety wire and loosen both jamnuts (42). Adjust the tension turnbuckle (43) as required to relieve belt tension. (4) Remove the lower attachment bolt (44), washers (45 and 46) and nut (3) from the tension turnbuckle (43) lower rod end (47). (5) Cut the safety wire and remove three socket head cap screws (25) and washers (24) from the belt housing (23).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (6) Remove the belt housing (23), laminated shim(s) (20), spacer (21) and snap ring (22). (7) Remove the drive belt (26).

B. Installation (1) Install the drive belt (26) on drive pulley (18) (Ref. Figure 201). (2) Install the drive belt housing (23), using three socket head cap screws (25) and washers (24). Torque the three socket head cap screws 25 to 30 inch-pounds and install MS20995C32 safety wire. (3) Insert the laminated shim(s) (20) into the belt housing (23). (4) Insert the spacer (21) into the belt housing (23). (5) Install the snap ring (22) and check for a 0.010 ± 0.002 gap between the snap ring (22) and the spacer (21). If gap is not correct, remove the snap ring (22) and the spacer (21) and add or remove laminated shim(s) (20) or laminations, install the snap ring (22) and the spacer (21), check gap and repeat until the correct gap is obtained (Ref. Figure 202). (6) Install the drive belt (26) on the clutch pulley (5). If the compressor is not installed, perform the COMPRESSOR INSTALLATION procedure in this section (Ref. Figure 201). (7) Perform the COMPRESSOR BELT ALIGNMENT procedure in this section. (8) Perform the COMPRESSOR BELT TENSION procedure in this section. (9) Install cowling panels (Ref. Chapter 71-10-00, COWLING INSTALLATION). (10) Perform the VAPOR CYCLE SYSTEM OPERATIONAL CHECK procedure (Ref. 21-52-00).

C. Alignment (1) Remove the right engine cowling panels as required (Ref. Chapter 71-10-00, COWLING REMOVAL). (2) The drive pulley groove and the compressor clutch pulley forward groove must be parallel and aligned with each other (Ref. Figure 203). Check the alignment of the pulleys by performing the following Steps: (a) Using a scale (12) measure the distance from the compressor mount (1) to the center of the drive belt (27). The measurement should be 0.490 ± 0.025 inch and the drive belt (27) should sit evenly in the groove of the clutch pulley (10) with no twisting or high sides (Ref. Figure 202). (b) If no adjustment is required go to Step (3). If adjustment is required, perform the following Steps: 1 Loosen the compressor pivot bolts (29 and 30) and the belt tension turnbuckle (43) lower attachment bolt (44) and nut (3) (Ref. Figure 201). 2 Cut the safety wire and loosen both jamnuts (42). Adjust the tension turnbuckle (43) as required to relieve belt tension.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL 3 If the compressor clutch pulley is at an angle (not parallel) with the drive pulley, loosen the pivot plate lock bolt (2) and the pivot plate pivot bolt (3) and rotate the pivot plate adjustment bolt (11) as required to place the pulleys parallel with each other. Tighten the pivot plate lock bolt (2) and the pivot plate pivot bolt (3) (Ref. Figure 202). 4 If the pulleys are parallel but not aligned, add or remove laminated shims (5) as required to obtain the 0.490 ± 0.025 inch required in Step (2)(a). 5 Perform the COMPRESSOR BELT TENSION procedure in this section. 6 Tighten the compressor pivot bolts (29 and 30) and the belt tension turnbuckle (43) lower attachment bolt (44) and nut (3) and repeat Step (2) (Ref. Figure 201). (3) Ensure all hardware has been tightened and safety wired, as required. (4) Install cowling panels (Ref. Chapter 71-10-00, COWLING INSTALLATION). (5) Perform the VAPOR CYCLE SYSTEM OPERATIONAL CHECK procedure (Ref. 21-52-00).

D. Tension NOTE: The tension of a new compressor belt should be checked after 50 hours of operation. (1) Remove the right engine cowling panels as required (Ref. Chapter 71-10-00, COWLING REMOVAL). (2) Using a push-pull gauge (6, Chart 1, Chapter 27-00-00) apply a force of 4.7 to 5.9 pounds at the midspan of the belt. The belt should deflect 0.15 inch. (3) If no adjustment is required proceed to Step (4). If adjustment is required, perform the following Steps: NOTE: The maximum length of rod end that is allowed to extend out of the tension turnbuckle (32) is 0.90 inch at each rod end (30). When these maximum dimensions are reached, the belt should be replaced (Ref. Figure 202). (a) Loosen the belt tension turnbuckle (43) lower attachment bolt (44) and nut (3) (Ref. Figure 201). (b) Loosen the compressor pivot bolts (29 and 30). (c) Cut the safety wire and loosen both jamnuts (42). Adjust the tension turnbuckle (43) as required to obtain the belt tension required in Step (2). (d) Tighten both jamnuts (42) and install MS20995C32 safety wire. (e) Torque the belt tension turnbuckle (43) lower attachment bolt (44) and nut (3) 60 to 85 inch-pounds. (f) Tighten the compressor pivot bolts (29 and 30). (4) Install cowling panels as required (Ref. Chapter 71-10-00, COWLING INSTALLATION). (5) Perform the VAPOR CYCLE SYSTEM OPERATIONAL CHECK procedure (Ref. 21-52-00).

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1. SPRING 2. QUILL SHAFT 3. NUT 4. PIVOT PLATE AJUSTMENT BOLT 5. CLUTCH PULLEY 6. WASHER 7. NUT 8. PACKINGS (2) 9. COTTER PIN 10. NUT 11. WASHER 12. COMPRESSOR MOUNT 13. SELF LOCKING INSERT 14. ALIGNMENT PIN 15. WASHERS 16. BOLT 17. BEARING 18. DRIVE PULLEY 19. BEARING 20. LAMINATED SHIM 21. SPACER 22. SNAP RING 23. BELT HOUSING 24. WASHERS (3) 25. SOCKET HEAD CAP SCREWS (3) 26. BELT 27. BOLT, WASHER AND NUT 28. BOLT, WASHER AND NUT 29. PIVOT BOLT, WASHER AND NUT 30. PIVOT BOLT, WASHER AND NUT 31. SHIM 32. PIVOT PLATE 33. DRAIN-TUBE 34. CLAMP 35. PIN-DOWEL 36. PIVOT-SLIDER 37. BOLT AND WASHERS (2) 38. BOLT AND WASHERS (6) 39. ROD END 40. COMPRESSOR MOUNT SUPPORT 41. SHIM 42. JAMNUTS 43. TURNBUCKLE 44. BOLT 45. COUNTERSUNK WASHER 46. WASHER 47. ROD END 48. BOLT, WASHER(S) AND NUT 49. AFT MOUNT PLATE 50. BOLT, WASHER(S) AND NUT 51. FORWARD MOUNT PLATE

52. BOLT, WASHER AND NUT 53. BOLT, WASHER AND NUT 54. PRESSURE RELIEF VALVE 55. DISCHARGE HOSE 56. SUCTION HOSE 57. POWER WIRE 58. SPLICE 59. GROUND WIRE 60. SCREW 61. CLIP 62. DISCHARGE FITTING 63. SUCTION FITTING 64. COMPRESSOR

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Figure 201 Vapor Cycle Compressor Installation (UA-1 and After; UB-1 and After; UC-1 and After Modified by Service Bulletin No. 2345)

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1. COMPRESSOR MOUNT 2. PIVOT PLATE LOCK BOLT 3. PIVOT PLATE PIVOT BOLT 4. PIVOT PLATE 5. LAMINATED SHIM 6. WASHER 7. PIVOT BOLT 8. FORWARD MOUNT PLATE 9. AFT MOUNT PLATE 10. CLUTCH PULLEY 11. PIVOT PLATE ADJUSTMENT BOLT 12. SCALE 13. ACCESSORY GEARBOX MOUNTING PAD 14. SCALE 15. COMPRESSOR MOUNT SUPPORT MOUNTING SURFACES 16. SPRING 17. PACKING 18. QUILL SHAFT 19. BEARINGS (2) 20. PULLEY 21. RETAINING RING 22. SPACER 23. LAMINATED SHIM 24. DRAIN TUBE 25. CLAMP 26. GEAR SHAFT 27. BELT 28. COMPRESSOR 29. COMPRESSOR MOUNT SUPPORT 30. ROD END 31. JAMNUT 32. TURNBUCKLE

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Figure 202 Vapor Cycle Compressor Alignment (UA-1 and After; UB-1 and After; UC-1 and After Modified by Service Bulletin No. 2345)

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NOT PARALLEL OR ALIGNED

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Figure 203 Pulley Alignment (UA-1 and After; UB-1 and After; UC-1 and After Modified by Service Bulletin No. 2345)

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ENVIRONMENTAL SYSTEMS TEMPERATURE CONTROL DESCRIPTION AND OPERATION

21-60-00 00

1. GENERAL The temperature of the cabin can be automatically controlled by presetting the temperature selector potentiometer to the desired temperature and setting the mode control switch to AUTO. The cabin temperature controller then sends the appropriate “heat” or “cool” command to the air cycle machine bypass valve and the ejector bypass valve. A dual-element temperature sensor installed in the conditioned bleed air main duct completes the Wheatstone Bridge resistance circuitry of the cabin temperature controller. When the cabin temperature falls, the Wheatstone Bridge becomes unbalanced and current flows from the heat command output of the cabin temperature controller. The ACM bypass valve and the ejector bypass valve receive the command and open to allow more bleed air to enter the conditioned bleed air ducts (Ref. Figure 1). The bypass valves modulate the amount of bleed air flow in automatic response to the requirements indicated by the changing resistance of the duct temperature sensors and the temperature sensor located in the cabin temperature controller. The duct temperature sensors are installed in the conditioned bleed air main duct in line with the third cabin window. The solid-state temperature controller is installed under the headliner in line with the fifth cabin window. The temperature of environmental air can be controlled manually by setting the mode control switch to MAN and holding the manual temperature control switch to the INCR or the DECR positions. When the manual control switch is set to INCR or DECR, power is supplied through the CABIN TEMP CONTROL circuit breaker, the mode switch and the manual switch to the appropriate bypass valve. When the switch is set to INCR, the ACM bypass valve opens, allowing warm bleed air to flow through the floor outlets. The ejector bypass valve may open if the ACM bypass valve has opened completely and the manual temperature control switch is held in the INCR position. The bypass valves close, as required, to allow the air cycle system to cool the bleed air when the manual switch is set to DECR. If the manual temperature control switch is held in the DECR position, the ejector bypass valve will close completely. When the ejector bypass valve is closed, the ACM bypass valve will begin to close to direct air through the air cycle machine for cooling of the cabin. All environmental temperature control switches and the cabin air temperature indicator are located on the copilot's inboard subpanel.

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Figure 1 Temperature Control Schematic

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ENVIRONMENTAL SYSTEMS TEMPERATURE CONTROL TROUBLESHOOTING

100100

1. PROCEDURES Troubleshooting of the environmental temperature controls is outlined in Figure 101.

Figure 101 Troubleshooting - Cabin Temperature Control

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ENVIRONMENTAL SYSTEMS AIR DUCT TEMPERATURE SENSOR MAINTENANCE PRACTICES

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1. PROCEDURES A. Resistance Check If the temperature control of the heating and cooling systems is responding to manual control but is not responding to automatic control, the fault may be found in the temperature sensing elements or in the cabin temperature controller. It is suggested that the sensing elements be tested before testing the controller because the controller is more reliable. Two sensing elements are installed downstream of the manual valve in the conditioned bleed air main duct. The elements are in line with the third cabin window. One of the elements has been thermally insulated to provide a response that lags behind the response of the non-insulated element. The lag in response prevents uncomfortably wide temperature ranges that would be caused by the overshoot of cabin heating. The temperature sensing elements can be checked as follows: NOTE: Refer to Figure 201 for a graph showing the resistance values of the duct temperature sensors. In order for the resistance measurements of the elements to comply with the readings shown on the graph, it is important that the temperature at the location of the elements be determined accurately. (1) Remove the headliner panel in line with the fifth cabin window to gain access to the electrical connector on the temperature controller. (2) Disconnect the electrical connector from the cabin temperature controller. (3) Using the graph, determine the correct resistance corresponding to the temperature measurement taken at the element location. Remove the RH floorboard in line with the third cabin window and disconnect the duct adjacent to the temperature sensors to take temperature readings. (4) Using a high precision ohmmeter, measure the resistance between pins 8 and 11 on the connector of the cabin temperature controller. This resistance should correspond to the resistance indicated on the line AB on the graph. The resistance measured between pins 11 and 12 on the controller should correspond to the line CD on the graph. Line AB represents the thermally lagged element. (5) If the measured resistance from a sensing element varies from the resistance shown on the graph by more than 1%, the elements should be replaced. (6) If the operation of the heating and cooling systems is abnormal or intermittent when the temperature mode switch is set to AUTO, the internal connections of the sensing elements may be loose. An intermittent fault in the elements can be detected as follows: (a) Operate the LH engine of the airplane. Place the temperature mode switch in the MAN position. Hold the manual temperature control switch in the INCR position for 60 seconds. (b) Monitor the resistive response of the element between pins 11 and 12 of the controller while the temperature is increasing. If the resistance fluctuates or increases above 200 ohms, then the element may have an intermittent fault.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (c) Monitor the resistive response of the element between pins 8 and 11. Hold the manual temperature control switch in the DECR position for 60 seconds. Any fluctuations in the resistance readings may indicate a fault in the sensing elements or in the element connections. (7) If the elements show no indication of incorrect resistance and the heating and cooling systems do not respond to automatic control, the cabin temperature controller should be tested to determine if the controller is defective.

B. Removal (1) Remove the RH carpet and floorboards in line with the third cabin window to gain access to the sensors installed on the conditioned bleed air main duct. (2) Disconnect the electrical connector from the duct temperature sensor. Remove the attaching screws and the cabin temperature sensor from the main duct. (3) Remove the attaching screws and the cabin temperature sensor from the main duct.

C. Installation (1) Install the temperature sensor and secure with the attaching screws. (2) Install the electrical connector on the sensor. (3) Install the floorboards and carpet.

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Figure 201 Resistance Values of Duct Temperature Sensors

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ENVIRONMENTAL SYSTEMS CABIN TEMPERATURE CONTROLLER MAINTENANCE PRACTICES

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1. PROCEDURES A. Removal (1) Remove the headliner panel in line with the fifth cabin window. (2) Remove the attaching screws from the controller and remove the controller sufficiently to gain access to the electrical connector on the controller (Ref. Figure 201). (3) Remove the electrical connector from the controller.

B. Installation (1) Install the electrical connector on the controller receptacle (Ref. Figure 201). (2) Install the controller on the mounting bracket and secure with the attaching screws. (3) Install the headliner panel to cover the controller.

C. Functional Test If a fault in the operation of the automatic mode of the heating or cooling systems cannot be found in the duct sensing elements, the cabin temperature controller should be tested. The temperature at the inlet of the controller should be between 65°F and 85°F to perform the test procedures. Refer to Figure 202 for a test circuit that can be used to test the controller.

D. Test Equipment (1) A 28 vdc, 2-amp power supply that is accurate within 1%. (2) Two 22-ohm, 50-watt resistors (R1 and R2). (3) A 100-ohm potentiometer with a direct reading dial (TR1). (4) Two 100-ohm (± 1%) potentiometers wired in series with two 100-ohm step decades (TR2 and TR3). (5) Two 28 vdc, 40-ma, MS25231-1819 lamps or equivalent (DS1 and DS2). (6) An air temperature indicator (M1) equivalent to the airplane indicator. (7) Connectors that mate with the cabin temperature controller and the air temperature gage. Refer to the Wiring Diagram Manual for the appropriate part numbers.

E. Test Procedure (1) Remove the headliner panel immediately aft of the fifth cabin window to gain access to the controller.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (2) Connect a power supply and test circuit to the cabin temperature controller. (3) Set TR2 and TR3 (potentiometers) to 118 ohms and turn on the power supply to apply 28 vdc to the controller. Adjust TR1 until DS1 and DS2 are not illuminated. The temperature gage should indicate ambient temperature. (4) Decrease the resistance on TR1 until DS2 begins to illuminate in pulses, then decrease the resistance until DS2 illuminates continuously. (5) Increase the resistance on TR1 until DS1 begins to pulse, then increase the resistance until the lamp is illuminated continuously. Adjust the resistance for TR1 until DS1 and DS2 are not illuminated. (6) Set TR1 to 0 ohms and TR3 to 155 ohms; DS2 should illuminate continuously. (7) Increase TR2 until DS1 pulses; TR2 should be 176 ohms ±2. Continue to decrease the resistance at TR2; DS1 should illuminate continuously. (8) Set TR1 to 2.5 ohms, TR3 to 85 ohms, and TR2 to 100; DS1 should illuminate continuously. Decrease TR2 until DS2 pulses; the resistance at TR2 should be 86.9 ohms ± 7. DS2 should illuminate continuously when continuing to decrease the resistance at TR2. (9) Shut off the power supply and disconnect the test circuit from the temperature controller. (10) Ensure that the air flow across the controller is moving towards the fan (Ref. Figure 202). If the controller does not perform as described above, the controller may be assumed to be defective and must be replaced.

F. Indicator Calibration (1) Connect a 28 vdc auxiliary power unit to the airplane. (2) Turn on the external power switch and place the cabin environmental mode selector switch in AUTO. (3) Remove the cover from the controller in the overhead panel. (4) Hold a thermometer, calibrated from 25°F to 125°F, close to the aft end of the controller for two minutes. (5) The temperature indicator located on the instrument panel should read within 2° of the thermometer reading at the controller. (6) Should the indicator reading be outside acceptable limits, remove the grey sealer from the small round adjusting potentiometer. (7) Using a small blade screwdriver, adjust the potentiometer slowly until the temperature indicator agrees with the thermometer. (8) Reseal the potentiometer with sealer (123, Table 1, Chapter 91-00-00). (9) Install the cover on the controller and return the airplane to service.

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Figure 201 Cabin Temperature Controller Test Circuit

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Figure 202 Cabin Temperature Controller

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CHAPTER 22 - AUTO FLIGHT TABLE OF CONTENTS SUBJECT

PAGE

AUTOPILOT 22-10-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Aileron Servo . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Rudder Servo . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Elevator Servo . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Control Cable Rigging . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Autopilot Control System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Rigging . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Aileron Servo Cable . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Rigging . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202 Elevator Servo Cable . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202 Rigging . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 203 Tensioning . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 204 Rudder Servo Cable . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 206 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 206 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 206 Rigging . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 211 Tensioning . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 211

22-CONTENTS

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List of Effective Pages CH-SE-SU

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22-LOEP

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22-CONTENTS

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1 201 thru 217

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22-LOEP

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

AUTO FLIGHT AUTOPILOT DESCRIPTION AND OPERATION

22-10-00 00

1. GENERAL The Sperry SPZ-2000 autopilot is certified for use in the Model 1900 Series Airliner. This is an integrated flight control system consisting of a full three-axis autopilot and flight director system. A combination of sensors, electrical servos, guidance displays, mode selectors, and computers perform the necessary flight computations. The system provides either full autopilot control of the airplane with simultaneous flight director monitoring, or manual control response to flight director display steering commands. The yaw axis may be engaged independently of roll and pitch for use as a yaw damper. The Sperry SPZ-2000 uses the Beech trim and servo system. NOTE: Improper servo cable tension may cause unstable or erratic response to autopilot signals.

A. Aileron Servo The aileron servo is mounted beneath the center floorboard forward of the rear spar. The cables from the servo capstan are connected to the aileron quadrant. When the autopilot is engaged, the servo rotates causing aileron deflection dependent on the direction of servo rotation.

B. Rudder Servo The rudder servo is located in the aft section of the airplane. One cable is wrapped around the capstan and connected to the main rudder cable. The other cable is connected to the autopilot servo quadrant. When the autopilot is engaged, the servo rotates in response to autopilot signals, causing rudder deflection dependent on the direction of servo rotation.

C. Elevator Servo The elevator servo is located next to the rudder servo in the aft section of the airplane. The cables wrapped around the servo capstan are connected to the main elevator control cables. When the autopilot is engaged, the servo rotates, causing elevator deflection dependent on direction of servo rotation.

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AUTO FLIGHT AUTOPILOT MAINTENANCE PRACTICES

200200

1. PROCEDURES A. Control Cable Rigging NOTE: Improper servo cable tension may cause unstable or erratic response to the autopilot signals. Set the servo cable tension after the primary cables have been rigged and tensioned properly in accordance with the respective control system rigging in Chapter 27. The tension is to be set with the applicable control surface in the neutral position. After the tension is set, work the applicable control surface, then check the tension and reset if necessary. Set the tension of the servo cables to the values specified in AUTOPILOT CONTROL SYSTEM RIGGING in this section.

2. AUTOPILOT CONTROL SYSTEM A. Rigging NOTE: Before the autopilot control cables can be rigged, the respective flight control cables must be rigged as specified in Chapter 27. The autopilot servo cable tension is to be set with the applicable control in the neutral position. After the tension has been set, work the applicable control surface and verify the tension and adjust if necessary. The servo cable tension can then be set as specified in the AILERON AND ELEVATOR SERVO CABLE TENSION GRAPH (Ref. Figure 201) or the RUDDER SERVO CABLE TENSION GRAPH (Ref. Figure 205).

3. AILERON SERVO CABLE A. Removal (1) Remove the aileron servo cables from the aileron quadrant (Ref. Figure 202). (2) Remove the aileron servo from the mounting bracket. (3) Before removing the capstan guard, mark the location of the guard for installation. (4) Remove the capstan guard by cutting the safety wire and removing the screws securing the guard in place. (5) Unwind the cables from the capstan.

B. Installation (1) Insert the cables in the recess provided in the capstan, and install the safety screw. (Install the cable with the turnbuckle near the servo motor) (Ref. Figure 202). (2) Wrap each cable approximately an equal number of turns in opposite directions through the full length of the capstan. (3) Install the capstan guard and safety all screws. Mount the servo to the bracket. (4) Install the servo cables to the aileron quadrant.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (5) Using the turnbuckle set the servo cable tension to approximately 50 pounds and work the control surface. (6) Refer to the AILERON AND ELEVATOR SERVO CABLE TENSION GRAPH and read the pounds of tension required for the measured temperature. Use the graph on Figure 201, Sheet 1 for the 3/32 inch diameter cable for the Sperry installation or use the graph on Figure 201, Sheet 2 for the 1/16 inch diameter cable for the Collins installation. Using the turnbuckle, adjust the servo cable tension to the tension shown in the graph.

C. Rigging (1) Place the aileron quadrant in the neutral position and install the rig pin in the quadrant (7, Table 1, Chapter 27-00-00). (2) Refer to the AILERON AND ELEVATOR SERVO CABLE TENSION GRAPH and read the pounds of tension required for the measured temperature. Use the graph on Figure 201, Sheet 1 for the 3/32 inch diameter cable for the Sperry installation or use the graph on Figure 201, Sheet 2 for the 1/16 inch diameter cable for the Collins installation. Using the turnbuckle, adjust the servo cable tension to the tension shown in the graph. (3) Remove the rig pin. (4) Work the controls and verify the cable clearance and that no binding exists. Run the cables through full travel. Assure a minimum of one-half wrap of cable on the capstan at full travel.

4. ELEVATOR SERVO CABLE A. Removal (1) Disconnect the servo cables from the main elevator control cables by removing the cable clamps (Ref. Figure 203). (2) Before removing the capstan guard mark the location of the guard for installation. (3) Remove the capstan guard by cutting the safety wire and removing the screws securing the guard in place. (4) Unwind and disconnect the cables from the capstan by removing the safety screws.

B. Installation (1) Insert the servo cable in the recess provided in the capstan and install the safety screws. (Insert the cable with the turnbuckle near the servo motor) (Ref. Figure 203). (2) Wrap each cable approximately an equal number of turns in opposite directions through the full length of the capstan. (3) Install the capstan guard and safety all screws. (4) With the elevator in the neutral position, connect the servo cable with the turnbuckle to the elevator UP cable using the attaching bolts and clamp. Torque to 55 ± 5 inch-pounds and check for a minimum gap of 0.005 inch remaining between the clamp halves.

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22-10-00

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (5) With the elevator in the neutral position, connect the other servo cable to the main elevator DOWN cable using the attaching bolt and clamp. Torque to 55 ± 5 inch-pounds and check for a minimum gap of 0.005 inch remaining between the clamp halves. (6) Using the turnbuckle, set the servo cable tension to approximately 50 pounds and work the control surface. (7) Refer to the AILERON AND ELEVATOR SERVO CABLE TENSION GRAPH and read the pounds of tension required for the measured temperature. Use the graph on Figure 201, Sheet 1 for the 3/32 inch diameter cable for the Sperry installation or use the graph on Figure 201, Sheet 2 for the 1/16 inch diameter cable for the Collins installation. Using the turnbuckle, adjust the servo cable tension to the tension shown in the graph.

C. Rigging (1) Install one rig pin (7, Table 1, Chapter 27-00-00) in the forward bellcrank and another in the elevator aft bellcrank . Access to the elevator aft bellcrank is through an opening in the skin on the left side of the vertical stabilizer just below the horizontal stabilizer (Ref. Chapter 6-50-00, STABILIZER ACCESS PANELS). (2) Attach the servo bridle cables to the capstan. (3) Install the cable retaining screws and safety. (4) Wrap the bridle cables on the capstan as follows: NOTE: For an initial starting position, ensure the cable clamps are located at the dimensions shown in Figure 203. (a) Wrap the bridle cable with the turnbuckle clockwise to the center of the capstan; route the cable through the idler pulley (if installed) and attach it to the main elevator up cable with the cable clamp. Torque to 55 ± 5 inch-pounds and check for a minimum gap of 0.005 inch remaining between the clamp halves. (b) Wrap the servo bridle cable without the turnbuckle counterclockwise to the center of the capstan; route the cable through the idler pulley (if installed) and attach it to the elevator down cable with the cable clamp. Torque to 55 ± 5 inch-pounds and check for a minimum gap of 0.005 inch remaining between the clamp halves. (5) Install the cable guard pins. (6) Refer to the AILERON AND ELEVATOR SERVO CABLE TENSION GRAPH and read the pounds of tension required for the measured temperature. Use the graph on Figure 201, Sheet 1 for the 3/32 inch diameter cable for the Sperry installation or use the graph on Figure 201, Sheet 2 for the 1/16 inch diameter cable for the Collins installation. Using the turnbuckle, temporarily adjust the servo cable tension to approximately twice the tension shown in the graph. (7) Remove the rig pins. (8) With the elevator in both the elevator up and elevator down positions, ensure there is a minimum of 0.50 inch clearance between the cable clamps and the surrounding aircraft structure. (9) Move the control surface through full travel several times to verify cable/clamp clearance with no binding.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (10) Install the rig pins. (11) Refer to the AILERON AND ELEVATOR SERVO CABLE TENSION GRAPH and read the pounds of tension required for the measured temperature. Use the graph on Figure 201, Sheet 1 for the 3/32 inch diameter cable for the Sperry installation or use the graph on Figure 201, Sheet 2 for the 1/16 inch diameter cable for the Collins installation. Using the turnbuckle, adjust the servo cable tension to the tension shown in the graph. (12) Remove the rig pins. (13) Perform a ground test of the autopilot/flight director system in accordance with the Pilot's Operating Handbook. (14) Install the access plates on the aft fuselage just aft and below the stabilon.

D. Tensioning NOTE: Before the autopilot elevator servo cables can be tensioned, the elevator flight control primary cables must be rigged. (1) Remove fuselage access panel 7 (Ref. Chapter 6-50-00, FUSELAGE ACCESS PANELS). (2) Remove the lower aft rig pin hole access panels 34 on the left and right side of the upper vertical stabilizer (Ref. Chapter 6-50-00, STABILIZER ACCESS PANELS). (3) Move the elevator surface (1) and install a rig pin (3) (7, Table 1, Chapter 27-00-00) in the elevator aft bellcrank through the vertical stabilizer (4). Using minimum force try to manually move the elevators up and down to verify that the rig pin is correctly installed. Ensure that the elevators do not move. Look at the opposite side of the vertical stabilizer to verify that the rig pin has protruded (Ref. Figure 206, Sheet 1). NOTE: Do not measure the tension when the temperature is varying rapidly or with the airplane located in direct sunlight. (4) Allow the temperature of the airframe to stabilize before measuring and adjusting the cable tension. (5) Measure the temperature in the compartment next to the elevator servo cables. (6) Refer to the AILERON AND ELEVATOR SERVO CABLE TENSION GRAPH and read the pounds of tension required for the measured temperature. Use the graph on Figure 201, Sheet 1 for the 3/32 inch diameter cable for the Sperry installation or use the graph on Figure 201, Sheet 2 for the 1/16 inch diameter cable for the Collins installation. (7) Position a cable tensiometer (4, Table 1, Chapter 27-00-00) on the elevator servo cable and measure the cable tension. Cable diameter is noted in Figure 201. NOTE: Cable tension tolerance is ± 2 pounds of the tension found in Figure 204. (8) Using the turnbuckle (Ref. Figure 202), temporarily set the elevator servo cable tension to approximately twice the required setting indicated by the tension graph in Figure 204. (9) Remove rig pin (3) from the elevator aft bellcrank (Ref. Figure 206, Sheet 1).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (10) Move the control surface through full travel three times to verify cable/clamp clearance and no binding. If binding is detected, perform the ELEVATOR SERVO CABLE RIGGING procedures in this section. (11) Move the elevator surface (1) and install a rig pin (3) (7, Table 1, Chapter 27-00-00) in the elevator aft bellcrank through the vertical stabilizer (4). Using minimum force try to manually move the elevators up and down to verify that the rig pin is correctly installed. Ensure that the elevators do not move. Look at the opposite side of the vertical stabilizer to verify that the rig pin has protruded (Ref. Figure 206, Sheet 1). (12) Using the turnbuckle (Ref. Figure 203), adjust the elevator servo cable to the tension shown in the graph in Figure 201 and install safety clips. (13) Remove rig pin (3) from the elevator aft bellcrank (Ref. Figure 206, Sheet 1). (14) Install the lower aft rig pin hole access panels 34 on the left and right side of the upper vertical stabilizer (Ref. Chapter 6-50-00, STABILIZER ACCESS PANELS). (15) Perform a ground test of the autopilot/flight director system in accordance with the Pilot’s Operating Handbook. (16) Install fuselage access panel 7 (Ref. Chapter 6-50-00, FUSELAGE ACCESS PANELS).

3/32" AILERON AND ELEVATOR SERVO CABLE TENSION GRAPH 50

POUNDS OF TENSION

40

30

20

10

0

NOTE:

SERVO CABLE TENSION TOLERANCE IS 2 POUNDS.

UC22B 062153AA.AI

Figure 201 (Sheet 1 of 2) Aileron and Elevator Servo Cable Tension Graph (Sperry Installation)

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1/16" ELEVATOR SERVO CABLE TENSION GRAPH 50

POUNDS OF TENSION

40

30

20

10

0

TOLERANCE:

UC22B 070568AB.AI

Figure 201 (Sheet 2 of 2) Aileron and Elevator Servo Cable Tension Graph (Collins Installation)

5. RUDDER SERVO CABLE A. Removal (1) Disconnect the servo cables from the main rudder control cables (Ref. Figure 204). (2) Before removing the capstan guards, mark their location for installation. (3) Remove the capstan guards by cutting the safety wire and removing the screws securing the guard in place. (4) Unwind and remove the cables from the capstan.

B. Installation (1) Insert the servo cables in the recess provided in the capstan and install the safety screws. (Install the cable with the turnbuckle at the far side of the servo motor) (Ref. Figure 204). (2) Wrap each servo cable approximately an equal number of turns in opposite directions through the full length of the capstan. (3) Attach the capstan guard and safety all screws. (4) With the rudder in the neutral position, connect the servo cable with the turnbuckle to the main rudder RIGHT control cable, using the attaching bolts and clamps. Torque to 55 ± 5 inch-pounds and check for a minimum gap of 0.005 inch remaining between the clamp halves.

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22-10-00

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (5) With the rudder in the neutral position, connect the other servo cable to the rudder servo quadrant. (6) Refer to the RUDDER SERVO CABLE TENSION GRAPH and read the pounds of tension required for the measured temperature. Use the graph on Figure 205, Sheet 1 for the 3/32 inch diameter cable for the Sperry installation or use the graph on Figure 205, Sheet 2 for the 1/16 inch diameter cable for the Collins installation. Using the turnbuckle, adjust the servo cable tension to the tension shown in the graph.

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Figure 202 Aileron Servo Installation

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Figure 203 Elevator Servo Installation

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Figure 204 Rudder Servo Installation

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

C. Rigging (1) Install rig pin (2) (9, Table 1, Chapter 27-00-00) in the rudder forward bellcrank (1) (Ref. Figure 208). (2) Install rig pin (2) (7, Table 1, Chapter 27-00-00) in the hole in the rudder aft torque tube (1) assembly. Using minimum force, try to manually move the rudder to verify proper rig pin installation (Ref. Figure 207). (3) Using the turnbuckle, adjust the servo cable as indicated in the RUDDER SERVO CABLE TENSION GRAPH. Use the graph on Figure 205, Sheet 1 for the 3/32 inch diameter cable for the Sperry installation or use the graph on Figure 205, Sheet 2 for the 1/16 inch diameter cable for the Collins installation. (4) Remove the rig pins. (5) Work the controls through full travel and verify the cable/clamp clearance and that no binding exists. Assure a minimum of one-half of a wrap of cable around the capstan at full travel.

D. Tensioning NOTE: Before the autopilot rudder servo cables can be tensioned, the rudder flight control primary cables must be rigged. (1) Remove fuselage access panel 8 (Ref. Chapter 6-50-00, FUSELAGE ACCESS PANELS). (2) Install rig pin (2) (7, Table 1, Chapter 27-00-00) in the hole in the rudder aft torque tube (1) assembly. Using minimum force, try to manually move the rudder to verify proper rig pin installation (Ref. Figure 207). NOTE: Do not measure the tension when the temperature is varying rapidly or with the airplane located in direct sunlight. (3) Allow the temperature of the airframe to stabilize before measuring and adjusting the cable tension. (4) Measure the temperature in the compartment next to the rudder servo cables. (5) Refer to Rudder Servo Cable Tension Graph and read the pounds of tension required for the measured temperature. Use the graph on Figure 205, Sheet 1 for the 3/32 inch diameter cable for the Sperry installation or use the graph on Figure 205, Sheet 2 for the 1/16 inch diameter cable for the Collins installation. (6) Position a cable tensiometer (4, Table 1, Chapter 27-00-00) on the rudder servo cable and measure the cable tension. Cable diameter is noted in Figure 205. NOTE: Cable tension tolerance is ± 3 pounds of the tension found in Figure 205. (7) Using the turnbuckle (Ref. Figure 204), temporarily set the rudder servo cable tension to approximately twice the required setting indicated by the tension graph in Figure 205, Sheet 1 or Figure 205, Sheet 2. (8) Remove rig pin (2) from the rudder aft torque tube (1) (Ref. Figure 207).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (9) Move the control surface through full travel three times to verify cable/clamp clearance and no binding. (10) Install rig pin (2) (7, Table 1, Chapter 27-00-00) in the rudder aft torque tube (1) (Ref. Figure 207). (11) Using the turnbuckle, adjust the rudder servo cable tension. Use the graph on Figure 205, Sheet 1 for the 3/32 inch diameter cable for the Sperry installation or use the graph on Figure 205, Sheet 2 for the 1/16 inch diameter cable for the Collins installation. Install the safety clips (Ref. Figure 204). (12) Remove rig pin (2) from the rudder aft torque tube (1) (Ref. Figure 207). (13) Perform the YAW DAMP/RUDDER BOOST CHECK or the AUTOPILOT CHECK procedure in the Airplane Flight Manual. (14) Install fuselage access panel 8 (Ref. Chapter 6-50-00, FUSELAGE ACCESS PANELS).

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Figure 205 (Sheet 1 of 2) Rudder Servo Cable Tension Graph (Sperry Installation) 1/16" AILERON AND RUDDER SERVO CABLE TENSION GRAPH 50

POUNDS OF TENSION

40

30

20

10

0

UC22B 070560AB.AI

Figure 205 (Sheet 2 of 2) Rudder Servo Cable Tension Graph (Collins Installation)

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1. ELEVATOR 2. RUDDER 3. ELEVATOR AFT BELLCRANK RIG PIN 4. VERTICAL STABILIZER 5. HORIZONTAL STABILIZER

A 5

1

B

4

2

3

VIEW LOOKING UP LEFT HAND SIDE DETAIL

A

Figure 206 (Sheet 1 of 2) Elevator Aft Bellcrank Rig Pin Installation

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UE22B 061915AB.AI

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

6. ELEVATOR AFT BELLCRANK 7. ELEVATOR AFT RIG PIN HOLE

6

FWD 7

VERTICAL STABILIZER (REF)

DETAIL

B UC22B 070104AA.AI

Figure 206 (Sheet 2 of 2) Elevator Aft Bellcrank Rig Pin Installation

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A

1. RUDDER AFT TORQUE TUBE 2. RIG PIN

1 2

VIEW

NOTE: EARLIER VERSIONS OF THE TORQUE TUBE SECTOR MAY NOT HAVE LIGHTENING HOLES INSTALLED.

A

UE22B 061914AA.AI

Figure 207 Rudder Aft Rig Pin Installation

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1. FORWARD RUDDER BELLCRANK 2. RIG PIN

A

1

2

DETAIL

A

UC22B 061980AA.AI

Figure 208 Rudder Forward Bellcrank Rig Pin Installation

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CHAPTER 23 - COMMUNICATIONS TABLE OF CONTENTS SUBJECT

PAGE

SPEECH COMMUNICATIONS 23-10-00 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Antenna . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Bonding Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 201 201 201

STATIC DISCHARGING 23-60-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Static Dischargers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Troubleshooting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 101 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 101 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Static Discharger . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Base Mount Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202 Base Mount Installation (On Aluminium Surfaces) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202 Stabilon and Taillet Static Discharger Mount Replacement . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 203 Base Mount Installation (On Composite Surfaces) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 203 Static Discharger Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 204 Static Discharger Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 204

23-CONTENTS

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List of Effective Pages CH-SE-SU

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DATE

23-LOEP

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201 and 202

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23-60-00

1 101 201 thru 204

Nov 1/09 Nov 1/09 Nov 1/09

23-LOEP

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

COMMUNICATIONS SPEECH COMMUNICATIONS MAINTENANCE PRACTICES

23-10-00 200200

1. ANTENNA A. Removal (1) Remove all electrical power from the airplane and disconnect the battery. (2) Remove the attaching hardware from around the antenna base (Ref. Figure 201). (3) Carefully break the seal between the antenna and the airplane skin surface. (4) Disconnect the coaxial cable from the antenna and secure the cable so it will not fall into the opening in the airplane skin. (5) Remove the antenna from the airplane.

B. Installation (1) Ensure that all electrical power is off and the battery is disconnected. (2) Perform antenna bonding procedures as described in this section. (3) Connect the coaxial connector to the antenna base and position antenna for installation (Ref. Figure 201). (4) Secure the antenna to the airplane skin with the attaching hardware. (5) Restore electrical power to the airplane. (6) Perform a ramp check of the system associated with the component installed.

C. Bonding Procedures NOTE: The pattern and efficiency of an antenna is dependent upon a low impedance (Z) path to the ground plane. It is imperative that a uniform resistance and capacitance be maintained between the contacting surface of the antenna and the surface on which it is mounted. The mating surface between the antenna and the base must be smooth and contoured to match, keeping the mating surfaces in actual contact. (1) Determine the contact area of the antenna being mounted. (2) Using trichloroethane (21, Table 1, Chapter 91-00-00) clean all grease, oil and other nonconductive films from the contact surfaces of the antenna, spacers and the mounting surface of the fuselage. WARNING: Use trichloroethane solution only in an adequately ventilated area. (3) Clean nonsoluble films from the contacting surfaces by sanding and polishing with fine garnet paper or silicon carbide, taking care not to remove excessive metal. Small areas on a aluminum surface may be cleaned with a stainless steel wire brush. The area should then be brushed clean and wiped with trichloroethane (21, Table 1, Chapter 91-00-00) and a clean cloth.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL CAUTION: Do not use emery or iron oxide paper/cloth in removing nonsoluble films from the contact area. (4) Apply Alodine 1200 (88, Table 1, Chapter 91-00-00) to the area with a clean pad, sponge, or equivalent applicator. Keep the area wet for 5 minutes or until film develops. WARNING: Use rubber gloves when handling Alodine. (5) Using low pressure water, gently wash the area to neutralize and remove the Alodine coating. (6) Allow the area to thoroughly dry for a maximum of one hour prior to installation of the mating parts. If the drying time exceeds one hour, retouch the area with Alodine and neutralize with water. (7) Attach the coaxial cable to the antenna. Any excess coaxial cable should be stored inside the airplane. Apply sealer (166, Table 1, Chapter 91-00-00) around the coaxial cable where it passes through the skin. CAUTION: Do not alter the length of any coaxial cable. A change in length could affect system operation. (8) Install the antenna with the attaching screws. All screws should be uniformly torqued. NOTE: If the screws are not uniformly torqued, the radiation pattern of the antenna may be affected. (9) Seal around the entire periphery of the antenna with sealer (166, Table 1, Chapter 91-00-00). (10) Touch up the painted area around the antenna as required.

VOR/LOC ANTENNA COM ANTENNA

COM ANTENNA

UC27B 042941AA.AI

Figure 201 Antenna Installation

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

23-60-00 00

COMMUNICATIONS STATIC DISCHARGING DESCRIPTION AND OPERATION 1. STATIC DISCHARGERS A static charge may build up on the surface of the airplane while in flight. This electrical charge, if retained, can cause interference in radio and avionics equipment operation. It is also dangerous to personnel disembarking after landing and to personnel servicing the airplane. Static dischargers keep the buildup of static charge at a minimum by continuously releasing it back into the atmosphere. Consequently, static dischargers are installed on the trailing edge of the flight surfaces and the wing tip (one each on the wing tips, the stabilons and tail lets, three on each aileron, four on each elevator, three on the rudder, two on the rudder tab, two on the ventral fin, one on each horizontal stabilizer tip and one on each of the forward and aft vertical stabilizer upper fairings) to aid in dissipating the electrical charge (Ref. Figure 1). NOTE: Due to the weight difference of available static dischargers, it is recommended that all static dischargers on a control surface be of the same type and brand to avoid balancing the control surface for each discharger replacement.

Figure 1 Static Wicks

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COMMUNICATIONS STATIC DISCHARGING TROUBLESHOOTING

100100

1. PROCEDURES When there is a noticeable or a suspected buildup of static effects due to precipitation or if a lightning strike has occurred, an inspection must be accomplished. Perform the inspection procedure for the static dischargers and control surface bonding jumpers (Ref. Model 1900 Airliner Series Corrosion Control Manual, P/N 114-590021-197, Chapters 23-60-01 and 23-60-02).

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COMMUNICATIONS STATIC DISCHARGING MAINTENANCE PRACTICES

200200

1. STATIC DISCHARGER A. Inspection The static dischargers should be inspected at the proper interval (Ref. Chapters 5-20-02 and 5-20-07). To functionally check, measure the resistance of the static discharger with a megohmmeter (28, Table 7, Chapter 91-00-00) at a minimum of 500 volts, a lesser test voltage will not generate a leakage rate sufficient to provide an accurate reading. When there is a noticeable or a suspected buildup of static effects due to precipitation or if a lightning strike has occurred perform the inspection procedure for the static dischargers and control surface bonding jumpers (Ref. Model 1900 Airliner Series Corrosion Control Manual, P/N 114-590021-197, Chapters 23-60-01 and 23-60-02). NOTE: To ensure the best possible functioning of the static dischargers, the resistance between the mounting base of the static discharger and the central grounding point must be 0.1 ohm or less. The resistance between the mounting base of static dischargers mounted on composite surfaces such as the tail lets and stabilons to the central grounding point should be 20 ohms or less. If resistances in excess of this amount are encountered, the mounting base should be removed and replaced. The resistance between the control surface on which the static discharger is attached and the central ground point must be less than 0.01 ohm. In order to make these small resistance measurements use a milliohmmeter (31, Table 7, Chapter 91-00-00). Check the Static Dischargers using the following procedure: WARNING: Electrical shock can result if the megohmmeter is improperly used. Refer to the applicable Manufacturer’s operating handbook for proper operating procedures. (1) Connect one megohmmeter test lead to the metal tip on the static discharger (on dischargers that have bristles instead of the metal tip use an alligator clip on the meter lead and clamp as many of the bristles as possible). (2) Connect the other megohmmeter test lead to the static discharger base mount. (3) Place the megohmmeter power switch to ON and observe meter indication. For trailing edge mounted dischargers the acceptable resistance range is 6 to 200 megohms. For tip mounted dischargers the acceptable range is 6 to 120 megohms. If a given discharger reads outside of the appropriate range, it must be replaced (Ref. Figure 201). Perform the STATIC DISCHARGER REMOVAL and the STATIC DISCHARGER INSTALLATION procedures in this section. (4) Check the resistance for the new static discharger. If the resistance is still too high, the threaded portion of the base mount is corroded. Base mount removal and installation procedures are different depending on if the base mount is located on an aluminium surface, on a stabilon, or on a composite surface. Perform the appropriate removal and installation procedure in this chapter. (5) Place the megohmmeter power switch to off and disconnect the test leads.

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B. Base Mount Removal NOTE: The static discharger base mounts need not be removed unless the resistance requirements under the TO CHECK RESISTANCE BETWEEN DISCHARGER BASE MOUNT AND AIRFRAME section of the corrosion control manual are not met (Ref. Model 1900 Airliner Series Corrosion Control Manual, P/N 114-590021-197, Chapters 23-60-01 and 23-60-02). Static dischargers on the stabilons are installed in tubes which are bonded into the stabilon tips. Special procedures are required for replacement of these mounts which are covered separately under STABILON STATIC DISCHARGER BASE MOUNT REPLACEMENT procedures in this section. (1) Unscrew the discharger from the base mount. (2) Remove the screws and lock washers securing the base mount to the surface. (3) Remove the base mount from the surface.

C. Base Mount Installation (On Aluminium Surfaces) (1) Clean an area slightly larger than the area to be bonded (minimum 0.25 inch) by removing all non-conductive material such as films, grease, oil, paints, metal finishes or other high resistance elements with 3M No. 600 grit sandpaper or equivalent, and solvent (2, Table 1, Chapter 91-00-00). The mating surface must be smooth and contoured so that maximum surface area is in contact. NOTE: An acceptable substitute for the preceding may be used in accordance with MIL-B-5087B (Bonding, Electrical and Lightning Protection for Aerospace Systems). The alodine solution should have an amber color. If the solution is coffee colored it has been contaminated. Repeat the cleaning procedure if contaminated alodine has been applied to the bonding surface. (2) Shake the alodine solution (88, Table 1, Chapter 91-00-00) vigorously just prior to application, then apply to the bonding surface with a clean Scotch Brite pad, sponge, brush or cloth. (3) Keep the treated area wet with alodine for approximately 3 to 5 minutes until a yellow color develops. Should the alodine not change color, it is an indication that the surface was not properly cleaned. NOTE: The bonding surfaces must be assembled within one hour of alodine treatment. Once dried, alodine must be softened before it can be effectively used in bonding. If more than an hour has passed, soften the alodine by applying wet alodine to the dried surface. (4) After the alodine has changed color rinse the area with clean, deionized water and gently wipe dry. Touch-up any areas where the alodine does not cover the bonding surface. Care must be taken not to damage the alodine coating as it is still soft when bonding. (5) Install the static discharger base mount, using the attaching screws and washers, onto the aluminum surface. Perform the Steps for the TO CHECK RESISTANCE BETWEEN DISCHARGER BASE MOUNT AND AIRFRAME section of the corrosion control manual (Ref. Model 1900 Airliner Series Corrosion Control Manual, P/N 114-590021-197, Chapters 23-60-01 and 23-60-02). (6) Perform the STATIC DISCHARGER INSTALLATION procedure in this section. Page 202 Nov 1/09

23-60-00

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (7) Perform STATIC DISCHARGER INSPECTION procedure in this section. (8) Finish the surface area around the base mount attachment point with the original finish.

D. Stabilon and Taillet Static Discharger Mount Replacement (1) Carefully remove the static discharger attach tube from the stabilon or taillet tip and clean out the hole to 0.16-inch diameter. (2) Fill the hole with P/N 610-1016 conductive adhesive (188, Table 1, Chapter 91-00-00) using a probe to help spread the adhesive evenly inside of the hole. (3) Apply conductive adhesive liberally to the outside diameter of the tube and insert the tube into the hole until the end is flush with the trailing edge of the stabilon. NOTE: Assure that the threads in the tube are protected from the adhesive as static discharger installation could be hampered. (4) Allow the adhesive to dry according the adhesive manufacturer's instructions. (5) After the adhesive is cured perform the Steps for the TO CHECK RESISTANCE BETWEEN DISCHARGER BASE MOUNT AND AIRFRAME section of the corrosion control manual (Ref. Model 1900 Airliner Series Corrosion Control Manual, P/N 114-590021-197, Chapter 23-60-02). (6) Perform the STATIC DISCHARGER INSTALLATION procedure in this section. (7) Perform STATIC DISCHARGER INSPECTION procedure in this section.

E. Base Mount Installation (On Composite Surfaces) NOTE: Do not file, grind, sand or otherwise remove material from the surface. Clean with alcohol only. (1) Abrade with 180 grit sandpaper or finer to roughen the surface resin uniformly without damaging the underlying fiber. (2) Remove dust by vacuum and wipe the surface using a lint-free cleaning cloth (189, Table 1, Chapter 91-00-00) dampened with isopropyl alcohol. Wipe in one direction until non-composite contaminants are gone from surface. Wipe surface dry without letting alcohol evaporate. (3) Install the static discharger base mount onto the composite surface. Perform the Steps for the TO CHECK RESISTANCE BETWEEN DISCHARGER BASE MOUNT AND AIRFRAME section of the corrosion control manual (Ref. Model 1900 Airliner Series Corrosion Control Manual, P/N 114-590021-197, Chapters 23-60-01 and 23-60-02). (4) Perform the STATIC DISCHARGER INSTALLATION procedure in this section. (5) Perform STATIC DISCHARGER INSPECTION procedure in this section. (6) Finish the surface area around the base mount with the original finish.

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F. Static Discharger Removal (1) Static dischargers may be removed by turning the threaded discharger assemblies counterclockwise until they are disconnected from the base. The static discharger base mounts need not be removed unless the resistance requirements under the TO CHECK RESISTANCE BETWEEN DISCHARGER BASE MOUNT AND AIRFRAME section of the corrosion control manual are not met (Ref. Model 1900 Airliner Series Corrosion Control Manual, P/N 114-590021-197, Chapters 23-60-01 and 23-60-02). (2) If the static discharger is to be reused, do not discard the lock washer installed on each static discharger.

G. Static Discharger Installation NOTE: When installing new static discharger assemblies, discard the lock nut which is installed on each new static discharger assembly. This lock nut is used merely as a retainer nut to ensure against loss of the lock washer. (1) To ensure proper function of the static dischargers, all threaded connections should be cleaned with Toluol (Toluene) (18, Table 1, Chapter 91-00-00) prior to installation of the static dischargers. (2) Apply one drop of Locktite 222 thread locker to the threads of the discharger and insert the threaded connection, with the lock washer installed, into the base mount. Screw the discharger in by hand until the lock washer is completely compressed.

Figure 201 Static Discharger Installation

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CHAPTER 24 - ELECTRICAL POWER TABLE OF CONTENTS SUBJECT

PAGE

ELECTRICAL SYSTEM 24-00-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . AC Power . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . DC Power . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Battery Power . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . External Power . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Electrical Load Distribution . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Starter and Ignition . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

1 1 1 1 2 2 4 5

AC POWER AND CONTROL 24-20-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Troubleshooting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 101 Inverter . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 101 Power Select Relay Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 101 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Inverter Blower . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Fan Operational Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201

DC GENERATION AND CONTROL 24-30-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Power Panel . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Generator Control Panels . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Voltage Regulation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Differential Voltage Control . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Reverse Current Protection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2 Overvoltage Protection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3 Overexcitation Protection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3 Field Flash Circuit . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3 Remote Trip Function . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3 Cross-Start Overload Protection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3 Bus Tie System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4 Troubleshooting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 101 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 101 Voltage Checks . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 101 Overvoltage Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 101 Generator Control Panel Test Unit . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 106

24-CONTENTS

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CHAPTER 24 - ELECTRICAL POWER TABLE OF CONTENTS (CONTINUED) SUBJECT

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STARTER-GENERATOR 24-30-01 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal (UA-1 and After, UB-1 and After, UC-1 thru UC-142) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Shaft Spline Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation (UA-1 and After, UB-1 and After, UC-1 thru UC-142). . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal (UC-143 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation (UC-143 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Brush Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 201 201 204 207 207 210

GENERATOR CONTROL PANEL 24-30-02 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Voltage Adjustment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Field Sense Relay Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 201 201 201 202

BATTERY POWER AND CONTROL 24-31-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Troubleshooting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 101 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 101 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Maintenance Log . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Charging . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Disconnection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 203 Connection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 203 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 203 Pre-Installation Instructions for Nickel-Cadmium Batteries . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 204 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 204 Battery Cleaning and Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 205

BATTERY MONITOR 24-32-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Functional Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201

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CHAPTER 24 - ELECTRICAL POWER TABLE OF CONTENTS (CONTINUED) SUBJECT

PAGE

EXTERNAL POWER AND CONTROL 24-40-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Troubleshooting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 101 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 101 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Connecting the Ground Power Unit . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Applying Ground Power . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202 Removing Ground Power . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202 Disconnecting the Ground Power Unit . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202

ELECTRICAL LOAD DISTRIBUTION 24-50-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Bus Conformity Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Triple Fed Bus Diodes, Operational Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 203

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List of Effective Pages CH-SE-SU

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DATE

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1 thru 3

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1 thru 5

Nov 1/13

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1 101 thru 103 201

Nov 1/09 Nov 1/09 Nov 1/09

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Nov 1/09 Nov 1/09

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1 101 and 102 201 and 202

Nov 1/09 Nov 1/09 Nov 1/09

24-50-00

1 thru 9 201 thru 203

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24-LOEP

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

ELECTRICAL POWER ELECTRICAL SYSTEM DESCRIPTION AND OPERATION

24-00-00 00

1. GENERAL Direct current for the electrical systems of the airplane is supplied by two 30-volt, 300-ampere starter-generators and either a 24-volt, 34-ampere hour battery or a 24-volt, 23-ampere hour battery. Alternating current for certain engine instruments and for avionics is supplied by either one of two inverters. Electrical system repair methods used on Beechcraft Corporation airplanes must be made in accordance with the most current revision of the Federal Aviation Administration's “Aircraft Inspection and Repair” manual AC43.13-1 and with the “Aircraft Alterations” manual AC43.13-2. Any components replaced and any wire, cable, or terminals used in the maintenance of the electrical system must be of airplane quality. Any solder connections must be made in an approved manner. Any solderless terminals or splices used must be applied with tooling specified by the supplier.

A. AC Power The AC power for the avionic equipment and the AC powered engine instruments is supplied by either one of two inverters. The inverters are installed in the aft portion of each nacelle. The inverter operation is controlled through an inverter select switch on the left outboard subpanel. Selection of either inverter energizes the inverter power relay installed near that inverter to supply the operating DC power. Operation and control of the inverters is discussed in greater detail in 24-20-00. An AC volt/frequency meter has been provided for the purpose of monitoring inverter output and performance. The meter is located in the overhead meter panel. (Ref. Figure 2). All AC operated components are designed to operate within the range of voltage and frequency shown on the volt/frequency meter.

B. DC Power DC power from the generators and the battery is distributed to the various airplane systems by way of three primary buses, two generator buses and the center/battery bus. The center/battery bus has a RH side and a LH side located in each nacelle. Refer to the Wiring Diagram Manual to locate components on this bus. Each generator is connected to its respective bus by a line contactor. Under normal operation, all buses are tied together by bus tie relays. The bus tie relays are controlled by unidirectional high current sensors; therefore, any bus can be isolated from the other buses should a fault condition exist which would cause current flow through the bus to exceed a safe limit. An additional bus, triple fed by all three power sources through blocking diodes, has been installed for the purpose of supplying power to certain selected equipment. The hot battery bus is continuously powered to provide current to certain lights and other equipment without the necessity of turning on the battery switch. Two loadmeters and a voltmeter have been installed in the overhead meter panel (Ref. Figure 2) for monitoring generator output. The voltmeter and select switch makes it possible to monitor voltages on all the buses including external power at the plug. Proper operation of the two generators in concert with each other demands that the generators are properly controlled. Proper control of the generators is the responsibility of the generator control panels. The control panels provide not only the necessary control functions such as voltage regulation, equalization and load sharing, but certain protective functions as well. Reverse current protection, overvoltage protection, overexcitation protection and cross-start overload protection are all provided for by the control panels and will be discussed in greater detail in 24-30-00.

24-00-00

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

C. Battery Power Battery power is controlled through the battery switch located under the master switches gang bar with the generator switches. When the battery switch is turned on, the battery relay closes and the battery is connected to the triple fed bus through a 60-ampere feeder limiter and to the RH pitot heat through a 35-ampere feeder limiter. Primary control of the battery bus tie relay also occurs when the battery switch is turned on and the bus tie relay closes connecting the battery to the center/battery bus. The battery compartment is located in the top of the right inboard wing leading edge (Ref. Figure 3). The hot battery bus is always connected to the battery when the battery is installed and cannot be isolated from the battery. All equipment connected to the hot battery bus is fed through individual fuses located in the battery compartment. A thorough discussion and maintenance procedures of the battery system and its associated controls will be found in 24-31-00.

D. External Power The external power receptacle is located behind an access door on the lower surface of the left inboard wing section and is used to connect external power to the airplane electrical system. When the external power switch is turned on, the external power relay closes and external power is applied to both generator buses and the center/battery bus. The battery and hot bus will not be connected until the battery switch is turned on. An overvoltage sensor prevents the external power relay from closing if a voltage in excess of 32vdc is connected to the external power receptacle. More information on this system is detailed in 24-40-00.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 1 Electrical System Schematic

24-00-00

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 2 Overhead Meter Panel

E. Electrical Load Distribution Load distribution is the manner in which the various primary buses in the electrical system are loaded. Each bus is broken down as to exactly what equipment is installed on the bus and the equipment power requirements. AC load distribution is covered in this chapter as well as DC load distribution. This information has been detailed in 24-50-00.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 3 Battery Compartment

F. Starter and Ignition The starter and ignition circuits provide power to motor the engine and for ignition during engine starts. By selection of the starter only mode, the starter can be used for motoring the engine to clear the engine of excess fuel; however, under no circumstances should the intermittent duty cycle for the starter be exceeded. An annunciator is illuminated to indicate when power is applied to the igniters. More detailed information on the ignition circuits will be found in Chapter 74-00-00.

24-00-00

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

ELECTRICAL POWER AC POWER AND CONTROL DESCRIPTION AND OPERATION

24-20-00 00

1. GENERAL During normal operation, an inverter power select relay is energized and power is supplied from the starter-generator bus. Should a fault occur that would interrupt power to that bus, the power select relay would de-energize and operating power would be taken from the center bus of the airplane, precluding the possibility of loss of an inverter due to failure of power sources. On airplanes equipped with inverter cooling blowers, the blowers are wired in parallel with the inverter power relays. As a result, any time an inverter power relay is energized, its respective cooling blower will be operating. A voltage-frequency meter is installed to monitor the inverter output. The meter is located in the overhead meter panel. The meter monitors frequency, except when the button in the lower left corner of the meter is pressed to monitor voltage. The inverter select relay is energized when the number one inverter is selected and de-energized when the number two inverter is selected. The inverter warning relay is energized open when either inverter is selected to prevent the illumination of the inverter warning annunciator when the selected inverter is supplying 115 VAC. The inverter select relay and the inverter warning relay are located on the forward right equipment panel. Pictorial coverage of the equipment panels is provided in Chapter 39. The inverter system described here is the standard installation. The circuit diagram in the wiring diagram manual provides the circuit routing of the DC and AC power for the standard airplane instrumentation.

24-20-00

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

ELECTRICAL POWER AC POWER AND CONTROL TROUBLESHOOTING

100100

1. INVERTER When troubleshooting the inverter system, all electrically operated systems and components should be turned off except the battery which must be turned on in order to power the inverters and associated controls. When checking for continuity and resistance, all electrical power must be remove from the airplane. Failure to do so may result in erroneous measurements and damage to test equipment (Ref. Figures 101 and 102).

A. Power Select Relay Check A failure in one of the inverter power select relay circuits could go undetected for some time; therefore, it is recommended that the following check be performed at regular intervals. Any results contrary to the expected results of the check will indicate that a fault in the inverter power select relay or its associated circuitry is present. (1) Start both engines according to the Pilot’s Operating Handbook and FAA Approved Airplane Flight Manual, and bring both starter-generators on the line. (2) Place the BUS SENSE switch in the TEST position; center bus voltage should read zero; starter-generator bus voltages should read 28 vdc. (3) Set the INV SELECT switch to NO. 1; the FREQ/AC VOLTS meter should read 115 ± 3.5 volts at 400 ± 4 hz. (4) Open the GEN 1 INV PWR SEL circuit breaker; the FREQ/AC VOLTS meter should read zero. (5) Place the BUS SENSE switch in the RESET position closing the bus ties. The FREQ/AC VOLTS meter should read 115 ± 3.5 volts at 400 ± 4 hz. (6) Place the BUS SENSE switch in the TEST position opening all bus ties. (7) Close the GEN 1 INV PWR SEL circuit breaker; the FREQ/AC VOLTS meter should read 115 ± 3.5 volts at 400 ± 4 hz. (8) Set the INV SELECT switch to NO. 2; the FREQ/AC VOLTS meter should read 115 ± 3.5 volts at 400 ± 4 hz. (9) Open the GEN 2 INV PWR SEL circuit breaker; the FREQ/AC VOLTS meter should read zero. (10) Close the bus ties again by placing the BUS SENSE switch in the RESET position. The FREQ/AC VOLTS meter should read 115 ± 3.5 volts at 400 ± 4 hz. (11) Place the BUS SENSE switch in the TEST position. (12) Close the GEN 2 INV PWR SEL circuit breaker; the FREQ/AC VOLTS meter should read 115 ± 3.5 volts at 400 ± 4 hz. (13) Shut down both engines according to the Pilot’s Operating Handbook and FAA Approved Airplane Flight Manual.

24-20-00

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 101 AC Power Schematic

Page 102 Nov 1/09

24-20-00

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 102 AC Power and Control (Inverter Inoperative)

24-20-00

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ELECTRICAL POWER AC POWER AND CONTROL MAINTENANCE PRACTICES

200200

1. INVERTER BLOWER A. Fan Operational Check (1) Remove access panels 60 and 61 (UA-1 and After, UB-1 and After) or left and right access panels 4 (UC-1 and After) (Ref. Chapter 6-50-00, Wing Access Panels). (2) Apply external electrical power (Ref. Chapter 24-40-00). (3) Set the INV SELECT switch to NO. 1. (a) Check inverter blower fan for operation and correct direction of flow (toward inverter). (4) Set the INV SELECT switch to NO. 2. (a) Check inverter blower fan for operation and correct direction of flow (toward inverter). (5) Remove external electrical power. (6) Install access panels 60 and 61 (UA-1 and After, UB-1 and After) or left and right access panels 4 (UC-1 and After) (Ref. Chapter 6-50-00, Wing Access Panels).

24-20-00

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

ELECTRICAL POWER DC GENERATION AND CONTROL DESCRIPTION AND OPERATION

24-30-00 00

1. GENERAL The generator phase of operation is controlled through the generator control switches located under the master switches gang bar with the battery switch. If the generator is taken off the line while the generator is running, the generator control switch must be placed in reset in order to bring the generator back on the line. The DC generation system makes use of DC starter-generators, power panels, generator control panels, battery bus tie relay and current sensor, control switches, a voltmeter, load meters and annunciator lights and the bus tie printed circuit board. Proper operation of each of these components is essential to obtain satisfactory generator operation. Many different component failures can result in similar generator system response, making it extremely difficult to isolate a faulty component merely by analyzing the system responses. The generator provides electrical power to recharge the airplane battery and operate the airplane electrical loads. These loads require a constant voltage source for proper operation. The generator output is maintained at a constant level by controlling the shunt field excitation. The interpole and compensating windings of the generator are in series with the armature and provide a voltage proportional to the generator current.

A. Power Panel The DC power panels, mounted on the LH and RH nacelle electrical equipment panels, contain in a single package, the line contactor, bus tie relay, start relay, and the unidirectional high current sensor. The loadmeter shunt and the bus tie limiter are mounted externally on the power panel. Each relay in the power panels is equipped with auxiliary contacts which are used for various control and logic functions.

B. Generator Control Panels All phases of generator operation are controlled by the generator control panels which are mounted in the center aisle subfloor just aft of the main spar. The generator control panels provide voltage regulation, generator load sharing, differential voltage and reverse current sensing and control, overvoltage and over excitation protection, the field flash circuit and cross-start overload protection (Ref. Figure 1).

C. Voltage Regulation The generator output voltage is sensed at the generator side of the generator line contactor and is used to power the voltage regulator circuit of the control panel at pin J, as well as to sense input to the control panel at pin B. The regulator circuit of the control panel supplies the generator field excitation current required to supply current for the electrical loads and to maintain a bus voltage of 28.25 ± 0.25 vdc.

D. Differential Voltage Control When the generator control switch is turned off, the power to the line contactor coil control power is interrupted at pin H of the control panel, thus removing the generator from the bus. At the same time, a ground signal is applied to the remote trip circuit at pin N of the control panel to trip the control panel's internal field relay, which de-excites the generator.

24-30-00

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Equalization of the generators is accomplished by utilizing the voltage developed across the generator interpole and compensating windings and is sensed by the control panel at pin D. The paralleling circuit includes the LH and RH generator control panels. An interconnect between pins E on the control panels completes the paralleling circuit. Equal load sharing is dependent upon equal resistance of the generator winding and of the external circuitry. The generator control panels are designed to control the generators and the load shared within 10% of the rating of one generator for total loads greater than 25%.

Figure 1 Generator Control Panels

E. Reverse Current Protection The reverse current protection and differential voltage control are provided by the generator control panels. Bus voltage is sensed at the bus side of the line contactor by control panel pin A. The generator output voltage is sensed at the generator side of the line contactor by control panel pin B. Whenever the generator is operating and the control switch is placed in the reset position, the generator output voltage will rise to the regulated voltage. When the generator switch is placed in the on position, a voltage output from pin H of the control panel will close the line contactor to connect the generator to the bus, if the generator voltage is greater than bus voltage or not more than 0.5 vdc below bus voltage. When the generator field becomes underexcited for any reason, or when the generator slows down to where it can no longer maintain a positive load, the generator will begin to draw current from the airplane bus. This current, termed REVERSE CURRENT, passes through the interpole and compensating windings of the generator. The voltage developed by this current is sensed at pin D of the generator control panel. When this voltage exceeds 0.5 volts, the generator control panel removes the voltage from the coil of the line contactor, permitting the contactor to open, removing the generator from the airplane bus.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

F. Overvoltage Protection Overvoltage protection is provided by the generator control panel. In the event the output voltage reaches 32.5 vdc, the overvoltage protection portion of the generator control panel will trip an internal field relay and open the coil circuit of the line contactor. The overvoltage generator is thus de-excited and isolated from the airplane bus. Should a circuit fault occur which would supply generator output or bus voltage to the generator field, or should the voltage regulation of one of the control panels fail, the affected generator would attempt to assume the full electrical load. If the bus voltage were to rise above 28.25 ± 25 vdc reverse current would begin to flow in the normally regulated generator and it would be removed, temporarily, from the bus. Should the resultant voltage increase above 32.5 vdc in the affected generator, it will be removed from the bus also and the non-affected generator can be reconnected to the bus. The resultant bus voltage would depend upon the generator speed, the electrical load and the nature of the fault. A positive test of this overvoltage function of the control panel is accomplished by applying generator output voltage to pin P of the control panel through the overvoltage test switch. This activates the overvoltage circuit which trips the internal field relay of the control panel and de-excites the generator. At no time during this test will the generator actually go into overvoltage.

G. Overexcitation Protection Overexcitation protection is provided by the generator control panel. This portion of the control panel will activate in the event the generator load and speed conditions are such, or the nature of the fault is such, that the generator voltage starts to increase without control, but does not go into overvoltage. Should the generator load unbalance reach the designed limitation value, the circuitry will be activated to remove the affected generator from the bus.

H. Field Flash Circuit When the generator switch is placed in reset, the generator residual voltage from terminal B+ of the starter-generator is applied to the generator field at terminal A+ through a low resistance circuit, bypassing the regulator until the generator voltage builds up high enough for the voltage regulator to effectively control the generator. Any time the generator control panel has been tripped for overvoltage or the generator has a low residual voltage, reset must be used in order to bring the generator on the line.

I. Remote Trip Function When the generator control switch is turned off, a ground signal is applied to pin N of the control panel, tripping the control panel internal field relay which opens the generator field and de-excites the generator.

J. Cross-Start Overload Protection The generator control panel has been designed to protect certain electrical system components from damage in the event cross-starts are attempted. This is accomplished by way of the current limiting feature of the control panel. When the start switch is placed in start, the start signal is applied to pin S of the control panel and disables the voltage regulator circuit, preventing the generator from being excited. Simultaneously the start signal is applied to pin R of the opposite control panel and activates the current limiting circuit of the opposite control panel, which protects the opposite generator from overload.

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K. Bus Tie System A system of current sensors, bus tie relays and the bus tie printed circuit board is utilized to provide protection in the event of a ground fault condition on one of the buses. Three sensors, one on the center/battery bus and one in each power panel, monitor current flow between the buses. Any time excess current is detected (approximately 275 amps) flowing through the current sensor in the reverse direction back toward one of the buses, the sensor will open the coil of bus tie relay for that bus and the faulted bus will be isolated from the rest of the system. The generator bus tie printed circuit board initiates the closed mode activity of the bus tie system by supplying energizing current to the coils of the generator bus tie relays, located in the power panels, when the generator line contactors close. The generator bus tie control circuits of the bus tie pcb will also be energized when an external power supply is connected and the external power switch and battery switch are on. The generator bus tie relays can also be closed manually through the bus tie switch. The battery bus tie relay closes automatically when the battery switch is turned on. When a sensor detects high current on the bus it is controlling, it supplies a ground signal to its respective bus tie deactivate circuit of the bus tie pcb and opens the coil circuit of its respective bus tie relay. The bus tie relay will remain open until closed by placing the BUS SENSE switch in the RESET position. During engine starts, the bus tie pcb disarms the high current sensors to prevent tripping of the bus tie relays. The bus tie system can be functionally checked by placing the test switch, located on the left outboard subpanel, in the test mode which provides a 28 vdc signal to the test circuit of the sensors and simulates a high current condition. The bus tie relays can be reset by placing the BUS SENSE switch in the reset mode. The bus tie switch, located on the left outboard subpanel, makes it possible to manually open the generator bus ties when they are in the closed mode by opening the grounding circuits of the bus tie relays. The circuits are restored to their closed state when the switch is placed in the NORM position. Annunciators for L GEN TIE OPEN, R GEN TIE OPEN and BATT TIE OPEN are activated through the annunciation circuits of the bus tie pcb. Additionally, the MAN TIES CLOSE annunciator is activated when the GEN TIES switch is placed in the CLOSED position.

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ELECTRICAL POWER DC GENERATION AND CONTROL TROUBLESHOOTING

100100

1. PROCEDURES The starter-generator and controls troubleshooting is best accomplished by monitoring each of the control panel inputs for proper response during various modes of operation and comparing these responses to those of a normal system. Table 101 contains detailed information on continuity and voltage measurements that may be expected at the generator control panel of a normal system. The continuity checks should be performed with the control panel disconnected, the battery off, external power off and the engines shut down. The resistance values shown in the Table 101 are approximate values. It is not expected that these exact values will be obtained during checkout; they should be used rather as a guide to locate open and short circuits in the airplane wiring. Usually, it should not be necessary to check each of the inputs given in Table 101 to isolate a defective component. Checking the ground wires and one or two specific circuits is normally sufficient. The troubleshooting charts in this chapter list the observed system fault and probable cause and identify the pins of the generator control panel which should be monitored as well as the expected voltages. Many times the description of the system malfunction is sufficient to indicate which circuits to check (Ref. Figures 103 thru 107).

A. Voltage Checks A test jack on the RH inboard subpanel provides for sampling DC voltage during routine checks. If voltage adjustments are to be made, voltage measurements should be taken at the jack provided on the top of each generator control panel.

B. Overvoltage Check An overvoltage protect-switch located on the generator control panel shelf of each control panel should be used to test the overvoltage function of each control panel. Placing the test switch in the TEST position applies bus voltage to pin P of the control panel. A good test of the overvoltage circuits of the control panels results in the generator under test being removed from the line and de-excited; this response is confirmed by residual voltage at pin B of the control panel. Any other response is indicative of a fault in the control panel.

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Figure 101 DC Generation and Control Schematic

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Figure 102 Generator Control Panel Test Unit

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C. Generator Control Panel Test Unit A “Generator Control Panel Test Unit Assembly” (P/N 1999/935) should be used to gain access to the individual inputs and outputs of the generator control panel. A wiring schematic provides adequate information for building the test unit (Ref. Figure 102). CAUTION: Never connect or disconnect the generator control panel wiring harness plug when the engine is running as this may damage the control panel. The wiring harness plug is disconnected from the generator control panel and connected to the receptacle of the test unit. The plug from the test unit is inserted into the receptacle of the generator control panel, thereby completing the series attachment of the test unit to the control panel. Do not connect the control panel when checking continuities. A sensitive multimeter or volt/ohm meter, digital being preferred, capable of measuring voltages accurate to within one percent, is connected to the test unit by way of the banana jacks on the face of the test unit; proper polarity of these connections must be closely observed as both positive and negative values will be measured. The alligator clip from pin G should be positively grounded to the airplane structure. Individual inputs or outputs are selected by rotating the knob on the rotary switch to the desired switch position. A test switch in the unit applies a ground signal to pin E to check the operation of the equalizer circuit. A push-to-test light provides information concerning the operation of the generator field sense relay during engine start. During start, a dimly glowing test light is a normal indication; however, should the light flash brightly, it may be assumed that the field sense relay is not closing and that the control panel may have been damaged by the transient voltages generated during engine start. The test light may only flash once and then burn out. Refer to the procedure to check the field sense relays (Ref. 24-30-02, FIELD SENSE RELAY CHECK). CAUTION: Never replace a damaged control panel until the proper operation of the field sense relay has been confirmed. All voltage measurements are made with one generator on and a 50 percent load. Since maximum generator output is 300 amperes, enough electrical equipment should be turned on to establish approximately 150 amperes of load on the electrical system. Refer to 24-50-00 for loads utilized by the various electrical systems and components. When checking the generator equalizer circuit for operation, a voltage drop of two or three volts at pin B is an indication that the generator equalizer circuit is operating properly.

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Table 101 TROUBLESHOOTING DC GENERATION AND CONTROL VOLTAGE CONTINUITY VALUES AT CONTROL PANEL CONNECTOR Pin

Function

Continuity

Voltage OFF

RESET

ON

A

Bus Voltage Signal Input

∞

26

26

28

B

Gen Voltage Signal Input

1-5 ohms

Residual

28

28

C

Line Contactor Control Power Input

∞

0

0

28

D

Interpole Voltage Signal Input

Continuity

0

0

-0.7

E

Equalizer

∞

0

0

-0.7

F

Battery Input (Reset)

∞

0

26

0

G

Ground

Continuity

0

0

0

H

Line Contactor Control Output

40 ohms**

0

0

26

J

Regulator Power Input

1-5 ohms

Residual

28

28

K

Gen Start-up Power Input (reset) ∞

0

28

0

L

Power Ground

Continuity

0

0

0

M

Regular Power Output

2-5 ohms

0

4

6

N

Remote Trip

Continuity

0

28

28

P

Overvoltage

∞

1

11

11

R

Current Limiter input

300 ohms

0

0

0

S

Start Signal Input

300 ohms

26*

0

0

NOTE: All voltage measurements are made with one engine running at 70% N1 and a 50% load. The continuity checks should be performed with the control panel disconnected, the battery off, external power off and the engines shut down. * This value will be seen only during engine start and will be battery voltage. ** Some Power Panels may read high resistance due to a diode in series with the coil of the relay.

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Figure 103 Troubleshooting DC Generation and Control Generator Does Not Reset

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Figure 104 Troubleshooting DC Generation and Control Generator Does Not Come On Line

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Figure 105 Troubleshooting DC Generation and Control Generators Do Not Share Loads

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Figure 106 Troubleshooting DC Generation and Control Bus Ties Not Closing When Generator is Brought On Line

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Figure 107 (Sheet 1 of 2) Troubleshooting DC Generation and Control Bus Ties Not Opening During Test

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Figure 107 (Sheet 2 of 2) Troubleshooting DC Generation and Control Bus Ties Not Opening During Test

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ELECTRICAL POWER STARTER-GENERATOR MAINTENANCE PRACTICES

24-30-01

200200

1. PROCEDURES A. Removal (UA-1 and After, UB-1 and After, UC-1 thru UC-142) (1) Remove the engine cowling (Ref. Chapter 71-10-00). (2) Perform REMOVING GROUND POWER procedure (Ref. 24-40-00). (3) Tag and disconnect the electrical leads from the starter-generator (Ref. Figure 201). (4) Cut the safety wire, loosen the clamp and remove the air inlet cap from the aft end of the starter-generator. CAUTION: It is mandatory that the starter-generator be fully supported from the time the retaining clamp is loosened until the unit is removed from the engine. The starter-generator must never be allowed to support its own weight through the spline shaft engagement or damage to the shaft shear section will result. (5) Remove safety wire from the quick disconnect clamp. (6) Loosen the T-bolt on the quick disconnect clamp which secures the starter-generator on the quick disconnect mounting adapter. (7) Open the clamp and remove the starter-generator from the mounting adapter. (8) Remove and discard the packing from the splined drive shaft. (9) Perform the STARTER-GENERATOR SHAFT SPLINE INSPECTION.

B. Shaft Spline Inspection Whenever the starter-generator has been removed for any reason, the starter-generator shaft spline (male) must be cleaned and inspected as follows: CAUTION: It is extremely important not to allow contact between the oven cleaner and aluminum parts. Oven cleaner is detrimental to aluminum. (1) Apply a coating of oven cleaner (141, Table 1, Chapter 91-00-00) on a tooth brush, then utilizing the tooth brush, apply the oven cleaner on the surface of the starter-generator shaft spline (male). (2) Leave the cleaner on the surface for 15 to 30 minutes. (3) Using a tooth brush, remove the debris from the male spline area. (4) Use a wet (with water) cloth to remove any remaining debris and cleaning agent from the male spline. (5) Using a 10X power magnifying glass, inspect the starter-generator shaft spline (male) for damage that may be the result of Electrical Discharge Damage (EDD) (Ref. Figure 202).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL NOTE: Electrical pitting or arcing damage is recognized by its characteristic crater like indentations in the metal surfaces. The craters often have shiny floors (formerly molten metal) and may be surrounded by expelled material and heat discoloration. (6) If a black or reflective foreign material is observed, clean the area again as described in Steps (1) thru (4), and reinspect to confirm that the pits are not mechanical damage or wear caused by debris lodged between the male (starter-generator) spline and female (engine) spline. (7) If any EDD pitting is identified on the male spline, do an oil patch test per Pratt and Whitney Canada Service Bulletin No. 14318 and contact Hawker Beechcraft Corporation Customer Support. If the starter-generator is to be reinstalled, record the pitting observed for future reference.

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Figure 201 Starter - Generator Installation

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (8) If no electrical pitting has been found on the male spline, the starter-generator may be installed as described in STARTER-GENERATOR INSTALLATION. NOTE: Make sure the starter-generator male spline and surrounding areas are free of all cleaning agents before installing the starter-generator. The cleaning agent is corrosive to aluminum and may be removed with water.

Figure 202 View of Starter-Generator Shaft Electrical Discharge Damage (EDD)

C. Installation (UA-1 and After, UB-1 and After, UC-1 thru UC-142) (1) Install a new packing on the starter-generator splined drive shaft (Ref. Figure 201). CAUTION: It is mandatory that the starter-generator be fully supported from the time the unit is placed in position, until the clamp is installed and properly torqued. The starter-generator must never be allowed to support its own weight through the spline shaft engagement or damage to the shaft shear section will result. (2) Align the starter-generator with the mounting adapter and secure it in place with the quick-disconnect clamp. Using an inspection mirror, ensure that the clamp groove fully captures both mating flanges.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (3) When the unit is properly positioned, torque the T-bolt retaining nut to 50 inch-pounds. (4) Connect electrical wiring to the starter-generator according to tags or the appropriate Model 1900 or 1900C Airliner Wiring Diagram Manual. (5) Torque #10 terminal stud nuts (2) 20 to 25 in/lbs (Ref. Figure 203). (6) Torque the 3/8 inch terminal stud nuts (1) 220 to 235 in/lbs (Ref. Figure 203). (7) Install the cooling cap on the aft end of the starter-generator, then install the clamp and secure with safety wire. (8) After the installation is complete, run the engine at idle speed for at least two minutes. Shut down the engine and recheck the quick-disconnect clamp retaining nut for proper torque. NOTE: If torque has fallen below 25 inch-pounds, loosen the quick-disconnect clamp, check the starter-generator for proper alignment, and follow Steps (3) and (6) again to ensure proper installation of the unit. (9) Safety wire the T-bolt after the proper torque has been verified. (10) Install the engine cowling (Ref. Chapter 71-10-00).

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1. 3/8 TERMINAL STUD NUTS 2. #10 TERMINAL STUD NUTS

A 1 2

2

B

DETAIL

B

DETAIL

A UC24B 100069AA.AI

Figure 203 Starter-Generator Terminal Studs (UA-1 and After, UB-1 and After, UC-1 thru UC-142)

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D. Removal (UC-143 and After) (1) Remove the engine cowling (Ref. Chapter 71-10-00). (2) Perform REMOVING GROUND POWER procedure (Ref. 24-40-00). (3) Remove protective nipples from the terminal adapter studs. (4) Tag and disconnect the electrical leads from the starter-generator (Ref. Figure 201). (5) Cut the safety wire, loosen the clamp and remove the air inlet cap from the aft end of the starter-generator. CAUTION: It is mandatory that the starter-generator be fully supported from the time the retaining clamp is loosened until the unit is removed from the engine. The starter-generator must never be allowed to support its own weight through the spline shaft engagement or damage to the shaft shear section will result. (6) Remove safety wire from the quick disconnect clamp. (7) Loosen the T-bolt on the quick disconnect clamp which secures the starter-generator on the quick disconnect mounting adapter. (8) Open the clamp and remove the starter-generator from the mounting adapter. (9) Remove and discard the packing from the splined drive shaft. (10) Remove the three terminal adapters (4) for use on the new starter-generator (Ref. Figure 204, Detail C). (11) Perform the STARTER-GENERATOR SHAFT SPLINE INSPECTION.

E. Installation (UC-143 and After) (1) Install a new packing on the starter-generator splined drive shaft (Ref. Figure 201). CAUTION: It is mandatory that the starter-generator be fully supported from the time the unit is placed in position, until the clamp is installed and properly torqued. The starter-generator must never be allowed to support its own weight through the spline shaft engagement or damage to the shaft shear section will result. (2) Align the starter-generator with the mounting adapter and secure it in place with the quick-disconnect clamp. Using an inspection mirror, ensure that the clamp groove fully captures both mating flanges. (3) When the unit is properly positioned, torque the T-bolt retaining nut to 50 inch-pounds. (4) Install the three terminal adapters (4) (Ref. Figure 204, Detail C). (5) Torque the Starter/Generator terminal adapter stud nuts (2) 220 to 235 inch-pounds (Ref. Figure 204). (6) Connect electrical wiring to the starter-generator according to tags or the Model 1900/1900C Airliner Wiring Diagram Manual.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (7) Torque the 3/8 terminal stud nuts (1) 95 to 110 inch-pounds (Ref. Figure 204). (8) Install the protective nipples onto the three terminal adapters studs. (9) Torque #10 terminal stud nuts (3) 20 to 25 inch-pounds. (10) Install the cooling cap on the aft end of the starter-generator, then install the clamp and secure with safety wire. (11) After the installation is complete, run the engine at idle speed for at least two minutes. Shut down the engine and check the quick-disconnect clamp retaining nut for proper torque. NOTE: If torque has fallen below 25 inch-pounds, loosen the quick-disconnect clamp, check the starter-generator for proper alignment, and follow Steps (3) and (7) again to ensure proper installation of the unit. (12) Safety wire the T-bolt after the proper torque has been verified. (13) Install the engine cowling (Ref. Chapter 71-10-00).

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1. 3/8 TERMINAL STUD NUTS 2. TERMINAL ADAPTER STUD NUTS 3. #10 TERMINAL STUD NUTS 4. TERMINAL ADAPTER

A 1 1 3

2

2

1 3

1 4

2

DETAIL

DETAIL

C

B

B C

DETAIL

A UC24B 100111AA.AI

Figure 204 Starter-Generator Terminal Studs and Terminal Adapters (UC-143 and After)

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F. Brush Inspection The starter-generator brushes are provided with a diagonal groove from one side of the contact surface to a point on the opposite side which corresponds to the point of maximum permissible wear (Ref. Figure 205). Brush wear can be estimated by the position of the groove on the contact surface. If the wear grooves indicate any one is at, or near, 1/4 of its remaining life or if inspection reveals the need for any other maintenance, the unit should be removed from the airplane and maintenance performed in accordance with the applicable Lucas Aerospace (Lear Siegler) Component Maintenance Manual with Illustrated Parts List for the particular model number unit. NOTE: It should also be noted that, although cored brushes may cause the commutator to appear grooved in the area of the brush cores, this unusual appearance of the commutator should not be interpreted as cause for rejection unless any groove exceeds 0.020 inch. When brushes are replaced, full brush seating is comprised of both coarse brush pre-seating by sanding combined with final brush run-in by running the unit on a drive stand or operating the unit as a motor (Refer to the applicable Lucas Aerospace (Lear Siegler) Maintenance Manual. The frequency of brush inspection should be determined by the operator based upon experience with individual units. The brushes must be replaced if the wear limit will be reached before the next scheduled inspection.

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Figure 205 Starter-Generator Brush Wear Groove (View from Drive End)

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ELECTRICAL POWER GENERATOR CONTROL PANEL MAINTENANCE PRACTICES

24-30-02 200200

1. PROCEDURES A. Removal (1) Remove all electrical power from the airplane. (2) Locate the generator control panel shelf in the center aisle subfloor immediately aft of the main spar. (3) Remove the electrical connector from the control panel. (4) Remove the four retaining screws from the base of the control panel and lift the control panel out of the subfloor.

B. Installation (1) Position the control panel on the control panel shelf with the electrical receptacle pointing aft. (2) Install the four retaining screws in the flange of the control panel base. (3) Connect the electrical connector and restore power to the airplane. (4) Check the generator output voltage using the control panel voltage jacks. (5) If the generator voltage at the control panel is not 28.25 ± 0.25 vdc, refer to VOLTAGE ADJUSTMENT in this section. (6) Replace the floor panel after assuring proper generator output voltage.

C. Voltage Adjustment NOTE: A test jack under the RH inboard subpanel provides for sampling DC voltage during routine checks. This procedure requires that the voltage measurements be taken at the jack provided on top of each generator control panel. Anytime a voltage adjustment is required on one generator control panel, both units must be adjusted to ensure proper parallel operation. (1) Remove the cabin floorboard to access the generator control panels at FS 295.00. (2) Remove the screw, washer and voltage adjustment access hole cover on both generator control panels to expose each voltage adjustment screw. (3) Connect a voltmeter to the voltmeter jacks located on the front of the RH generator control panel. (4) A multimeter or volt/ohmmeter accurate within 1% should be used for electrical measurements. A digital readout is preferred. (5) Start the engines with procedures outlined in the Model 1900C Airliner Airplane Flight Manual. (6) Set both engines N1 speed to 71 ± 1%.

24-30-02

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (7) Bring the RH starter-generator on-line by placing the RH generator control switch in the RESET position; then release to ON. (8) Place the generator bus tie switch in the OPEN position to open the generator bus tie relays. (9) Insert a small screwdriver into the voltage adjustment screw access hole of the RH generator control panel. (10) Adjust the screw clockwise to increase, or counterclockwise to decrease output until a reading of 28.25 ± 0.25 vdc is obtained. (11) Place the RH generator control switch in the OFF position. (12) Bring the LH starter-generator on-line. (13) Repeat the adjustment procedure for the LH generator control panel. (14) Shut down the engines with procedures outlined in the Model 1900C Airliner Airplane Flight Manual. (15) Disconnect the voltmeter. (16) Replace both voltage adjustment access hole covers and secure to the generator control panels with the cover screws and washers. (17) Install the floorboard panel.

D. Field Sense Relay Check (1) Connect the generator control panel test unit between the control panel and the control panel plug. (2) While monitoring the test lamp on the control panel test unit, turn on the battery and place the start control switch in START. (3) If the test lamp on the test unit flashes brightly then goes out, the field sense relay is not operating properly. A continuously glowing test lamp during start is a normal indication.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

ELECTRICAL POWER BATTERY POWER AND CONTROL DESCRIPTION AND OPERATION

24-31-00 00

1. GENERAL The airplane is equipped with a 20-cell, nickel-cadmium battery rated at 24 volts and 34 ampere-hours (one-hour rate) or a 23-ampere-hour battery. The battery is installed in the RH inboard wing and is ventilated. Vents located in the top surface of the wing and under the battery permit cooling of the battery during operation and allow the escape of gases produced if an overcharge occurs. The nickel-cadmium battery consists of a steel case containing a number of identical and individual cells connected to each other in series and fitted side-by-side in the battery case. The cells are interconnected by link bars. These bars are held in place on each cell by nuts on threaded terminals. The end cells are connected by a solid or a flexible link to a terminal on one face of the battery case or to a terminal that extends through the battery cover. The design of nickel-cadmium batteries allows for replacement of the individual cells if one cell becomes damaged. Service facilities for nickel-cadmium batteries must be separate from lead-acid battery facilities. The electrolyte contained in nickel-cadmium batteries is a highly alkaline solution of potassium hydroxide and water. This solution is a chemical “opposite” to the sulfuric acid contained in lead-acid batteries. Anything associated with lead-acid batteries, including acid fumes, should never come in contact with a nickel-cadmium battery or its electrolyte. If traces of sulfuric acid enter a nickel-cadmium battery, it can become damaged permanently. If the electrolyte becomes contaminated with tap water, acids or other non-compatible substances, poor performance or complete failure of the battery will result. If a battery operates with damaged, missing or loose vent caps, the result will be a low battery capacity caused by the loss of electrolyte.

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ELECTRICAL POWER BATTERY POWER AND CONTROL TROUBLESHOOTING

100100

1. PROCEDURES Battery power is available to the hot battery bus whenever the battery is connected to its power cables (Ref. Figure 101). When the battery switch is placed in the ON position, the battery relay and the battery bus tie relay close, making battery power available to the triple fed bus, the RH pitot heat feeder, and the center battery bus. The battery circuit breaker, located adjacent to the battery, protects the control power circuit to the coil of the battery relay. The BAT position of the voltmeter select switch can be selected to monitor battery voltage at the battery relay. For battery troubleshooting data (Ref. Figures 102, 103 and 104).

Figure 101 Battery Power Schematic

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Figure 102 Troubleshooting Battery Power and Control No Power on Busses

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Figure 103 Troubleshooting Battery Apparent Loss of Capacity or Frequent Addition of Electrolyte

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Figure 104 Troubleshooting Battery Excessive Spewage or Distorted Cell Cases

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ELECTRICAL POWER BATTERY MAINTENANCE PRACTICES

200200

1. PROCEDURES CAUTION: Service facilities for nickel-cadmium batteries must be entirely separate from those for lead-acid batteries. Fumes from lead-acid batteries or small traces of sulfuric acid entering a nickel-cadmium battery can damage it permanently. Use not only separate filler equipment, but separate supplies of distilled water to avoid contamination of the nickel-cadmium cells. Brushes, scrapers, cloths, tools or other implements used to maintain lead-acid batteries must never be used on nickel-cadmium batteries. Maintenance of the battery should be performed at regular intervals to obtain maximum service from the battery. The battery must be removed from the airplane to perform the necessary servicing. A new battery should be serviced at the first 100 flight-hours, and thereafter at the battery manufacturer’s recommendations. However, if battery is subjected to exceptionally heavy use such as frequent engine starting using the battery, or if the battery is operated at temperatures higher than what the manufacturer recommends, the service interval should be reduced. Battery condition should be monitored by using a battery maintenance log (Ref. Model 1900 Airliner Series Component Maintenance Manual, P/N 114-590021-11). For best operation and maximum life, nickel-cadmium batteries should be completely disassembled and all components thoroughly inspected and cleaned at least once a year as outlined in the Model 1900 Airliner Series Component Maintenance Manual, P/N 114-590021-11. For most applications, this maintenance can be scheduled to coincide with a major inspection of the airplane itself. Since complete battery servicing requires two days, an additional battery may be required to allow use of the airplane. WARNING: The electrolyte is caustic and can cause serious burns if it comes into contact with the skin. If electrolyte does contact the skin, the area should be flushed immediately with large amounts of water, and neutralized with a 3% solution of acetic acid, vinegar, lemon juice, or a 10% solution of boric acid. For treatment of electrolyte in the eyes, flush with large amounts of water and contact a physician immediately.

A. Maintenance Log Because of the importance of keeping track of the liquid level as well as the general state of charge and condition of the battery, it is strongly advised that a maintenance log be kept of all service and maintenance. Not only are careful records helpful in correcting battery malfunctions in normal servicing, but they are vital to the substantiation of battery warranty claims. Refer to Table 201 for a sample format of a service log.

B. Charging All charging maintenance shall be performed according to the manufacturer’s instructions outlined in the Model 1900 Airliner Series Component Maintenance Manual, P/N 114-590021-11.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Table 201 Battery Maintenance Log Catalog No.

Serial No.

Date Installed:

Installed On:

Removal Date and Reason

General Condition

End of CC Charge Voltage Range Electrical Level Electrical Leak Check Capacity

M - Maint. F - Failure (Indicate Type)

Page 202 Nov 1/09

Case and cover

Hardwa re and liners

Cells and vents

24-31-00

Lowest reading Min. 1.50

Highest reading Max. 1.70

Remarks Indicate: Average water added, hardware or cell replacements, burns or discolorations, etc

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

C. Disconnection WARNING: Remove all watches, rings, and metal jewelry before attempting maintenance on the battery. If metal articles contact the intercell connectors of opposite polarity, the objects will fuse themselves to the connectors and result in severe burns to the wearer. (1) Remove battery access panel located on the top surface of the RH wing center section (Ref. Chapter 6-50-00, WING ACCESS PANELS). (2) Check for installation and condition of the vent gasket on the battery access panel. Replace as necessary. (3) Cut the safety wire and remove the battery connector (1) from the battery (2). Install placard on the connector stating “DO NOT CONNECT TO BATTERY” (Ref. Figure 201). (4) Position the cable so it will not accidently make contact with the battery terminal.

D. Connection WARNING: Remove all watches, rings, and metal jewelry before attempting maintenance on the battery. If metal articles contact the intercell connectors of opposite polarity, the objects will fuse themselves to the connectors and result in severe burns to the wearer. (1) If necessary remove battery access panel located on the top surface of the RH wing center section (Ref. Chapter 6-50-00, WING ACCESS PANELS). (2) Connect the battery connector (1) to the battery (2) and safety wire (Ref. Figure 201). (3) Install battery access panel located on the top surface of the RH wing center section (Ref. Chapter 6-50-00, WING ACCESS PANELS).

E. Removal WARNING: Remove all watches, rings, and metal jewelry before attempting maintenance on the battery. If metal articles contact the intercell connectors of opposite polarity, the objects will fuse themselves to the connectors and result in severe burns to the wearer. (1) Remove battery access panel located on the top surface of the RH wing center section (Ref. Chapter 6-50-00, WING ACCESS PANELS). (2) Check for installation and condition of the vent gasket on the battery access panel. Replace as necessary. (3) Cut the safety wire and remove the battery connector (1) from the battery (2) (Ref. Figure 201). (4) Cut the safety wire and remove the hold down wing nuts (4), keeper (7) and washer (8) then push the clevis (9) aside. (5) Lift the battery (2) out of the battery box (3).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (6) Check the gasket in the bottom of the battery box (3) for installation and condition, replace as necessary. (7) Check the Vent at the bottom of the battery box (3) for corrosion or obstruction, clean as necessary.

F. Pre-Installation Instructions for Nickel-Cadmium Batteries NOTE: Unless otherwise indicated by a red warning tag, nickel-cadmium batteries are shipped in a completely discharged state, but with the proper amount of electrolyte in the cells. Observe the following precautions to ensure maximum performance and protect the battery warranty. (1) Do not remove the shorting clip until just before the battery is to be charged. Batteries that have had the shorting clip removed (even for a short period of time) must be considered to have an unknown charge and require a complete discharge prior to charging and installation procedures. (2) Inspect batteries shipped from the factory for shipping plugs in the vent holes of each battery cell. The blunt aluminum screws that serve as shipping plugs must be removed prior to operation of the battery. The bunson valves, included with the battery in a separate plastic bag, should then be screwed into the vent cap assembly in place of the screw plugs. The bunson valves will release excessive pressure to prevent cell rupture caused by gas accumulation. NOTE: On batteries not equipped with the screw-type plugs and bunson valves, remove the shipping plugs and clean the filler cap vent plugs as noted under battery cleaning procedures. Tighten the cell vents with the vent plug wrench included with the battery. (3) Check the terminal screws securing the cell links for tightness. Refer to the service sheet furnished with the battery or to the manufacturer’s maintenance manual for the proper torque value. (4) Before charging, determine that all cells are properly installed by making a cumulative voltage check. (5) After determining the battery is in good physical condition and is properly assembled, the electrolyte level should be adjusted and the battery charged as outlined by battery charging procedures.

G. Installation WARNING: Remove all watches, rings, and metal jewelry before attempting maintenance on the battery. If metal articles contact the intercell connectors of opposite polarity, the objects will fuse themselves to the connectors and result in severe burns to the wearer. (1) If necessary remove battery access panel located on the top surface of the RH wing center section (Ref. Chapter 6-50-00, WING ACCESS PANELS). (2) Position the battery (2) in the battery box (3) (Ref. Figure 201). (3) Secure hold down bar (5) with the keepers (7), washers (8) and wing nuts (4). Tighten and secure the wing nuts (4) with safety wire to the keepers (7). (4) Connect the battery connector (1) to the battery (2) and safety wire.

Page 204 Nov 1/09

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (5) Install battery access panel located on the top surface of the RH wing center section (Ref. Chapter 6-50-00, WING ACCESS PANELS).

H. Battery Cleaning and Inspection Reference the manufacturer’s instructions in the Model 1900 Airliner Series Component Maintenance Manual, P/N 114-590021-11 for when to inspect and how to clean the battery. Always keep a record of maintenance, service and condition in a battery maintenance log. Use the log to monitor battery condition over time, and to decide if the battery needs a shorter maintenance interval than what is called out in the manufacturer’s instructions.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

4 8 7

A

9

1

DETAIL

2

B

3 1. BATTERY CONNECTOR 2. BATTERY 3. BATTERY BOX 4. WING NUT 5. HOLD-DOWN BAR 6. VENT 7. KEEPER 8. WASHER 9. CLEVIS

4

B 5 FWD

6

A

INBD

DETAIL VIEW LOOKING DOWN UC24B 070791AA.AI

Figure 201 Battery Installation

Page 206 Nov 1/09

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

ELECTRICAL POWER BATTERY MONITOR DESCRIPTION AND OPERATION

24-32-00 00

1. GENERAL A battery charge current monitor is installed to provide a visual indication of an abnormal battery charge current. The system annunciator light illuminates if conditions exist that cause the charge current to be higher than normal. The battery may then be disconnected from the charging circuit, if necessary. During normal operation, the idle current of the battery is less than one amp. It increases significantly above the normal level when the battery is charged at an elevated temperature or from a high charge voltage. A high idle current increases water consumption and may destroy the gas barrier (cellophane separator) between the plates. Once a battery has sustained damage to the gas barrier, it will have a high idle current and will be subject to thermal runaway. The battery monitor system provides an indication of the high current resulting from high battery temperature, or high charging voltage, or gas barrier damage. The battery monitor system consists of a 250-ampere shunt in the negative lead of the battery, a battery charge current monitor, and a BATTERY CHARGE light (yellow) in the caution/advisory annunciator panel. The shunt is located adjacent to the battery box (RH inboard wing). The current monitor module is installed beneath the center aisle floor aft of the LH forward partition. For the shunt and the module maintenance practices (Ref. Chapter 39-20-00). Following an engine start, the BATTERY CHARGE light will illuminate for approximately five minutes or until the charge current decreases to the reset level of the current monitor. The light will remain illuminated more than five minutes if the trigger level of the monitor is set too low, or if the battery is in a low state of charge, or has been discharging slowly. This illumination functions as a self-test of the battery charge monitor system. After the BATTERY CHARGE light extinguishes, it should remain off unless a battery condition needs monitoring or unless the battery idle current increases in response to an increase in electrical system voltage. Such a voltage increase normally results from poor generator paralleling or load switching. The light may illuminate for short intervals as the battery recharges until the generator speed is increased above cut-in speed.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

ELECTRICAL POWER BATTERY MONITOR MAINTENANCE PRACTICES

200200

1. PROCEDURES The charge current monitor receives a signal from the battery shunt and provides 28 volts to illuminate the BATTERY CHARGE light when the signal exceeds the trigger level. This trigger level corresponds to 8 ± 1 amp of battery charge current. The circuit will reset when the signal decreases by a set amount below the trigger level. A six-second time delay in the monitor circuit prevents the illumination of the annunciator during momentary recharges of the battery.

A. Functional Test Calibration of the battery monitor system can be checked on the airplane by simulating a charge current sufficient to trigger the current detector circuit. A circuit constructed to accomplish this function is shown in Figure 201. It is recommended that system calibration be checked at regular intervals when the battery is removed for service. NOTE: The signal level is very low and susceptible to resistance imbalance in the shunt signal leads. All connections in these leads must be clean and tight to prevent improper system calibration and erratic operation. Connect a variable 40-volt DC power supply; a variable 0 to 10- amp DC power supply; a ± 10 volt DC voltmeter, two lamps (P/N 327), and a 250-amp 50-millivolt shunt (Ref. Figure 201). Test the battery charge monitor circuit as follows: (1) Apply 28 volts to pin 14. Set the 10-amp power source to 0 amps. The two lights (P/N 327) should not illuminate. (2) Apply 36 volts to pin 14. Apply 10 amps to the shunt. The two lights (P/N 327) should illuminate after approximately six seconds. Connect the voltmeter to pin 13; the reading should be 35 ± 0.5 volts. (3) Set the 10-amp power source to 0 amps. The two lights (P/N 327) should extinguish and the voltage reading at pin 13 should be less than one volt. (4) Apply 28 volts to pin 14. Connect the voltmeter to pin 5; the reading should be + 8 ± 2 volts. (5) Increase the amperage at the 10-amp power source until the voltmeter reading at pin 5 switches to -8 volts. The amperage reading should be 8 ± 1 amp. The two lights (P/N 327) should illuminate approximately 6 seconds after the voltage at pin 5 switches from + 8 volts to -8 volts. Decrease the amperage 0.5 amp. The voltage reading for pin 5 should switch to + 8 volts and the two lights (P/N 327) should extinguish. Connect the voltmeter to pin 13; the reading should be less than one volt.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 201 Battery Monitor Functional Test Circuit

Page 202 Nov 1/09

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

ELECTRICAL POWER EXTERNAL POWER AND CONTROL DESCRIPTION AND OPERATION

24-40-00 00

1. GENERAL The external power receptacle is located immediately aft of the LH main gear door. The receptacle is designed for use with an auxiliary ground power unit equipped with a standard AN plug. An overvoltage sensor module circuit protects the airplane electrical system from an APU with reversed polarity or excessively high output voltage. When an APU is connected to the receptacle, the sensor module utilizes voltage from the hot battery bus to allow voltage to be delivered between the positive terminal and the small, polarizing terminal of the APU plug to illuminate the EXTERNAL POWER annunciator light (green). The annunciator light will be illuminated when the APU is turned ON or OFF if the APU is plugged into the receptacle. The external power relay and the overvoltage sensor PCB are located adjacent to the inverter in the LH nacelle. The circuit breaker is located adjacent to the external power receptacle. Refer to Chapter 39-20-00 for further information on these components.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

ELECTRICAL POWER EXTERNAL POWER AND CONTROL TROUBLESHOOTING

100100

1. PROCEDURES If APU polarity and voltage are correct, sensor output voltage is routed to the external power switch. When the switch is placed in the ON position, the external power relay closes, supplying power to the center battery bus (Ref. Figure 101). Control voltage from the small pin of the external power receptacle is routed through a circuit breaker and the external power switch to the bus tie PCB to close the generator bus tie relays. Power is then applied through these closed relays to the generator busses and the triple fed bus. In addition, small pin voltage is routed to the select switch on the overhead meter panel to allow monitoring of APU voltage. If the battery is not connected to its power cables, the battery switch must be placed in the ON position to supply external power through the switch to close the battery bus tie relay and the battery relay. When these relays are closed, external power is supplied to the hot battery bus and the RH pitot heat feeder. Refer to Figure 102 for troubleshooting data.

Figure 101 External Power Schematic

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 102 Troubleshooting External Power and Control No Voltage on Busses

Page 102 Nov 1/09

24-40-00

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

ELECTRICAL POWER EXTERNAL POWER AND CONTROL MAINTENANCE PRACTICES

200200

1. PROCEDURES The ground power unit used for engine starts or for ground operation without the engines running must be capable of supplying load requirements without excessive voltage drop. The unit must be capable of delivering up to 1000 amperes for one second and be capable of delivering up to 300 amperes continuously at 24 to 30 volts. Use of an inadequate ground power unit can cause a voltage drop that is below the dropout voltage of the external power relay and the starter relay. This will result in relay chatter or welded contacts. CAUTION: Use only an auxiliary power source that is negatively grounded. If the polarity of the power source is unknown, determine the polarity by using a voltmeter before connecting the power unit to the airplane. The output setting must not exceed 1000 amperes on external power sources with a higher current-carrying capability. Any current in excess of 1000 amperes may overtorque the drive shaft of the starter-generator or produce heat sufficient to shorten the life of the unit. Voltage is required to energize the Avionics Master power relays to remove power from the avionics equipment. Therefore, never apply external power to the airplane without first applying battery voltage. If the battery is removed from the airplane or if the battery switch is to be placed in the OFF position, connect the external battery in parallel to the external power unit prior to energizing the auxiliary power unit. If external power is to be used for ground maintenance, such as landing gear rigging, ensure that every avionics unit is turned off. The battery may be damaged if exposed to voltages higher that 30 volts for extended periods of time. A continuous load in excess of 350 amperes will damage the external power relay and the power cables of the aircraft.

A. Connecting the Ground Power Unit (1) Review the cautions above. (2) Start the ground power unit and adjust the output to 28 ± 0.5 vdc with a maximum current output of 1000 amperes. (3) Shut the ground power unit down. (4) Open the ground power receptacle access panel. (5) Connect the ground power unit electrical connector to the airplane ground power receptacle.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

B. Applying Ground Power (1) If not previously accomplished, perform the CONNECTING THE GROUND POWER UNIT procedure. (2) Start the ground power unit. (3) Rotate the voltmeter select switch to the EXT PWR position. (4) Monitor the voltage of the ground power unit on the voltmeter for correct input voltage. (5) Set the BATT switch to ON. (6) Set the EXT PWR switch on the left outboard subpanel to the EXT PWR position. (7) Perform the required maintenance actions.

C. Removing Ground Power (1) Set the EXT PWR switch to OFF. (2) Set the BATT switch to OFF. (3) If no further use of ground power is required, perform the DISCONNECTING THE GROUND POWER UNIT procedure.

D. Disconnecting the Ground Power Unit (1) Ensure the REMOVING GROUND POWER procedure has been performed. (2) Shut down the ground power unit. (3) Disconnect the ground power unit electrical connector from the aircraft ground power receptacle. (4) Close the ground power receptacle access panel. (5) Position the ground power unit clear of the airplane.

Page 202 Nov 1/09

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

ELECTRICAL POWER ELECTRICAL LOAD DISTRIBUTION DESCRIPTION AND OPERATION

24-50-00 00

1. GENERAL Table 1 provides the electrical load requirements of each item of electrical equipment on the airplane. All resistance loads are calculated at 80% of normal bus voltage. Lamp loads are calculated at 87% of full load current. To determine the total electrical load of the airplane, add the electrical load of the standard equipment to the load of the optional equipment installed in the airplane. The total load shall not exceed 90% (540 amperes) of the total generating capacity of the two 30-volt, 300-ampere starter-generators. When an item of equipment functions in more than one system, the load value per unit listed in the Table represents the highest value required to operate that item. Each generator supplies half of the load of the triple fed bus when both generators are operating with the bus ties closed. Table 1 Electrical Load Distribution Equipment

No. Units Used

Load Ea. Unit (Amps DC)

Cruise Load (Amps DC)

LH Fire Extinguisher

1

3.00

RH Fire Extinguisher

1

3.00

Cabin Fluorescent Lights

2

0.245

Door Lock Illumination Lamp

1

0.17

Door Post Lights

2

0.17

Door Entry Light

1

0.17

RH Firewall Shutoff Valve

1

2.00

LH Firewall Shutoff Valve

1

2.00

Step Illumination Light

1

0.30

Cargo Compartment Lights

4

.30

Over-Aisle Entry Light

1

0.30

0.30

Cockpit Emergency Lights

4

0.17

0.68

Baggage Compartment Light

1

0.67

0.67

Overvoltage Sensor and Advisory Light PCB

1

0.01

0.01

Battery Relay

1

0.35

0.35

External Power Relay

1

0.50

RH Pitot Heat

1

5.40

5.40

Cabin Reading Lights

4

0.30

1.20

Notes

Hot Battery Bus (W214)

Hot Battery Bus Total

0.49

Bus (W309)

9.1

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Page 1 Nov 1/09

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Table 1 Electrical Load Distribution (Continued) Equipment

No. Units Used

Load Ea. Unit (Amps DC)

Cruise Load (Amps DC)

Flap Motor

1

16.00

Flap Position Transmitter

1

0.08

0.08

Flap Position Indicator

1

0.08

0.08

Chip Detector Warning Lights

4

0.04

Flight Instrument Light Control

1

0.05

0.05

Flight Instrument Lights

18

0.024

0.43

RH Pneumatic Bleed Air Shutoff Valve

1

0.80

0.80

RH Bleed Air Control Relay

1

0.30

0.30

Fwd Vent Blower Relay - Low

1

0.35

0.35

LH Wing Fuel Vent Heater

1

2.25

RH Environmental Bleed Air Shutoff Valve

1

1.25

1.25

RH Precooler-Through Valve

1

1.00

1.00

LH Ice Vane Actuator

1

3.00

1.00

RH Ice Vane Actuator

1

3.00

RH Precooler Bypass Valve

1

1.00

LH Bleed Air Control Relay

1

0.30

0.30

Aft Vent Blower Relay - Low

1

0.35

0.35

No. 1 Inverter Power Select Relay

1

0.27

0.27

No. 1 Inverter Power Relay

1

0.27

0.27

Inverter-Select Relay

1

0.09

0.09

No. 1 Inverter

1

6.58

6.58

Pilot's Windshield Anti-ice Controller

1

0.30

0.29

Pilot's Windshield Anti-ice High Heat Relay

1

0.35

Pilot's Windshield Anti-ice Relay

1

42.7

Pwr Steering Pump Motor

1

24.00

Avionics Bus No. 2

1

11.61

7.53

Fwd Vent Blower Motor

1

21.00

21.00

Fwd Vent Blower Relay - High

1

0.35

0.35

Brake Deice Control Relay

1

0.30

Flap Motor Relay

1

0.95

Notes

LH Generator Bus (A184W2)

Page 2 Nov 1/09

24-50-00

1

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Table 1 Electrical Load Distribution (Continued) Equipment

No. Units Used

Load Ea. Unit (Amps DC)

Cruise Load (Amps DC)

Flight Hour Meter

1

0.25

0.25

LH Landing Light

1

8.93

Lower Rotating Beacon

1

3.60

3.60

Upper Rotating Beacon

1

3.60

3.60

Tail Floodlights (OPT)

2

2.68

Tail Floodlights (OPT)

4

2.67

RH Firewall Fuel Shutoff Valve

1

2.00

Pilot's Windshield Anti-ice Relay

1

0.35

Anti-skid Pump

1

15.00

Power Steering Relay

1

0.30

Power Steering Pump Relay

1

0.60

Power Steering Actuator

1

1.00

Power Steering Solenoid Valve

1

1.5

Power Steering Signal Amplifier

1

0.620

Autofeather Control Relay

2

0.30

L Bus Tie Relay

1

0.58

Autofeather Dump Solenoids

2

0.8

Autofeather Control Relays

2

0.30

Pilot's Map Light

1

0.34

0.34

Pilot's Control Wheel Clock

1

0.003

0.003

Copilot's Control Wheel Clock

1

0.003

0.003

Copilot's Map Light

1

0.34

0.34

Engine Instrument Lights Dimmer

1

0.05

0.05

Engine Instrument Lights Control

1

0.05

0.05

Engine Instrument Lights

26

0.04

1.04

LH Generator Bus Total

Notes

0.58

52.23

RH Generator Bus (A185W2) Antiskid Control Unit

1

0.54

Antiskid Pump Relay

1

0.27

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Page 3 Nov 1/09

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Table 1 Electrical Load Distribution (Continued) Equipment

No. Units Used

Load Ea. Unit (Amps DC)

Antiskid Solenoid Valve

1

1.5

Electric Trim Disconnect Relay

1

0.09

0.09

Synchrophaser Control Box

1

1.00

1.00

Overhead Floodlight

1

0.30

0.30

Overhead Floodlight Control

1

0.05

0.05

Subpanel and Pedestal Light Control

1

0.05

0.05

Cabin Fluorescent Lights

9

0.245

2.20

RH Landing Light

1

8.93

LH and RH Recognition Lights

2

2.68

Spar Cover Lights

4

0.14

0.56

Cabin Reading Lights

18

0.30

5.4

No Smoking/Fasten Seat Belt Signs Lights

2

0.1

0.2

Prop Sync Caution Light

4

0.04

0.16

Air Cond. Clutch

1

3.28

3.28

RH Fuel Vent Heat

1

2.25

N1 Speed Sensor PCB

1

0.10

0.10

Fwd and Aft Hot Gas Bypass Valves

2

1.20

2.40

Wing Strobe Lights

2

1.20

2.40

Tail Strobe Light

1

1.50

1.50

Alternate Static Port Heat

2

1.60

3.20

Stall Vane Heater

1

9.40

9.40

No. 2 Inverter

1

6.58

6.58

No. 2 Inverter Power Relay

1

0.27

No. 2 Inverter Power Select Relay

1

0.27

Copilot's Windshield Anti-ice

1

42.70

Copilot's Windshield High Heat Relay

1

0.35

0.35

Copilot's Windshield Anti-ice Control

1

0.30

0.30

Avionics Bus No. 3

1

3.91

3.91

Copilot's Windshield Anti-ice Control Relay

1

0.35

0.35

Aft Vent Blower Relay - High

1

0.35

0.35

Aft Vent Blower Motor

1

21.00

21.00

Page 4 Nov 1/09

24-50-00

Cruise Load (Amps DC)

0.27

Notes

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Table 1 Electrical Load Distribution (Continued) Equipment

No. Units Used

Load Ea. Unit (Amps DC)

Antiskid Relay

1

0.30

Electroluminescent Panels Power Supply

1

1.5

LH Firewall Shutoff Valve

1

2.00

RH Bus Tie Relay

1

0.58

0.58

Subpanel and Pedestal Lights

16

0.04

0.64

Condenser Blower Relay

1

0.60

0.60

RH Loadmeter Light

1

0.05

0.05

LH Loadmeter Light

1

0.05

0.05

Voltmeter Light

1

0.05

0.05

Ammeter Light

1

0.05

0.05

RH Generator Bus Total

Cruise Load (Amps DC)

Notes

1.5

68.92

Triple Fed Bus (W217) RH Start Relay

1

1.85

RH Start Control Relay

1

0.09

RH Field Sense Relay

1

0.09

Cabin Pressure Preset Solenoid Valve

1

0.21

0.21

Aural Annunciator Amplifier

1

1.00

1.00

Cabin Pressure Dump Solenoid Valve

1

0.90

Landing Gear Downlock Solenoid

1

0.25

0.83

Ram Air Door Solenoid Valve

1

0.21

0.21

LH Environmental Bleed Air Shutoff Valve

1

1.25

1.25

LH Precooler-Through Valve

1

1.00

1.00

LH Precooler Bypass Valve

1

1.00

1.00

LH Pneumatic Bleed Air Shutoff Valve

1

0.80

0.80

Bleed Air Control Relay

1

0.30

0.30

Cabin Air Temperature Indicator

1

0.33

0.33

Cabin Temperature Control Box

1

0.10

0.10

Air Cycle Machine Bypass Valve

1

2.00

2.00

Ejector Bypass Valve

1

2.00

2.00

24-50-00

1

Page 5 Nov 1/09

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Table 1 Electrical Load Distribution (Continued) Equipment

No. Units Used

Load Ea. Unit (Amps DC)

Cruise Load (Amps DC)

Air Duct Temperature Sensor

1

0.15

0.15

Landing Gear Handle Light

2

0.04

Avionics Load Bus No. 1

1

8.12

4.04

LH Pitot and Static Heat Element

1

5.40

5.40

Relay Panel No. 8, Landing Gear Control

1

0.30

Landing Gear Control Assembly

1

0.09

Landing Gear Power Pack Selector Valve

1

1.0

LH Fuel Quantity

1

0.03

0.03

LH Low Fuel Level Sensor

1

0.20

0.20

Fuel Cross Transfer Valve

1

1.42

1.42

Pilot's Encoding Altimeter

1

0.10

0.10

Avionics Master Relay

3

0.50

Annunciator Control Card

1

0.06

0.06

LH Fuel Flow Indicator and Transmitter

1

0.43

0.43

LH Oil Temperature and Pressure Indicator

1

0.15

0.15

LH Low Fuel Feed Sensor

1

0.20

0.20

LH Oil Pressure Indicator and Transmitter

1

0.05

0.05

Fire Detector Control

2

0.07

0.14

LH Igniter Exciter

1

1.00

LH Start Relay

1

1.85

LH Field Sense Relay

1

0.09

LH Start Control Relay

1

0.09

Stall Warning Lift Computer

1

2.00

Landing Gear Control Relay

1

0.09

Pilot's Turn and Slip Indicator

1

0.50

0.50

Cabin Fluorescent Lights - Partial

2

0.245

0.49

Prop Deice Timer

1

0.285

0.285

Automatic Prop Deice Control Relay

1

0.35

0.35

Cabin Reading Lights - Partial

4

0.30

1.20

Instrument Indirect Lights

10

0.17

1.70

RH Fuel Flow Indicator and Transmitter

1

0.43

0.43

Page 6 Nov 1/09

24-50-00

Notes

2

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Table 1 Electrical Load Distribution (Continued) Equipment

No. Units Used

Load Ea. Unit (Amps DC)

Cruise Load (Amps DC)

RH Oil Temperature and Pressure Indicator

1

0.05

0.05

RH Propeller Governor Solenoid

1

0.80

LH Propeller Governor Solenoid

1

0.80

Propeller Governor Test Relay

2

0.30

RH Igniter Exciter

1

1.00

RH Low Fuel Feed Sensor

1

0.20

0.20

RH Low Fuel Level Sensor

1

0.20

0.20

RH Fuel Quantity

1

0.03

0.03

Triple Fed Bus Total

Notes

28.84

Center Battery Bus (A184W3 and A185W3) Condenser Blower Motor

1

50.00

50.00

Deice Distribution Valve

1

3.50

0.04

Windshield Wiper Motor

1

6.00

6.00

LH and RH Manual Prop Deice Relay

2

0.27

LH and RH Wing Navigation Lights

2

0.92

1.84

Tail Navigation Light

1

1.02

1.02

Wing Ice Lights

2

1.43

2.86

Taxi Light

1

8.93

Inverter No. 1 and 2 Relay

1

6.58

No. 2 Inverter Power Relay

1

0.27

0.27

RH Prop Deice Heater

1

28.00

28.00

LH Standby Fuel Pump

1

10.00

Voltage Regulator

2

0.60

Generator Reset

1

0.09

Landing Gear Motor

1

180.00

No. 1 Inverter Power Relay

1

0.27

LH Prop Deice Heater

1

28.00

No. 1 Inverter Power-Select Relay

1

0.27

0.27

Inverter-Select Relay

1

0.09

0.09

1.20

0.27

24-50-00

Page 7 Nov 1/09

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Table 1 Electrical Load Distribution (Continued) Equipment RH Standby Fuel Pump

No. Units Used

Load Ea. Unit (Amps DC)

1

10.00

Center Battery Bus Total

Cruise Load (Amps DC)

91.86

Avionics Bus No. 1 No. 1 Comm (Receiver)

1

0.51

0.51

No. 1 Comm (Transmit)

1

5.10

1.02

No. 1 Nav

1

0.80

0.80

No. 1 Glidescope

1

0.51

0.51

Marker Beacon

1

1.20

1.20

Avionics Bus No. 1 Total

4.04

Avionics Bus No. 2 No. 2 Comm (Receiver)

1

0.51

0.51

No. 2 Comm (Transmit)

1

5.10

1.02

No. 1 ADF

1

1.00

1.00

Radar

1

3.50

3.50

No. 2 Transponder

1

1.50

1.50

Avionics Bus No. 2 Total

7.53

Avionics Bus No. 3 No. 2 Nav

1

0.80

0.80

No. 2 Glidescope

1

0.51

0.51

No. 1 DME

1

1.10

1.10

No. 1 Transponder

1

1.50

1.50

Avionics Bus No. 3 Total

Page 8 Nov 1/09

24-50-00

3.91

Notes

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Table 1 Electrical Load Distribution (Continued) Equipment

No. Units Used

Load Ea. Unit (Amps DC)

Cruise Load (Amps DC)

LH Torque Pressure Indicator

1

6.00

6.00

RH Torque Pressure Indicator

1

6.00

6.00

RH Torque Transmitter

1

6.00

6.00

LH Torque Transmitter

1

6.00

6.00

AC Inverter Warning Light Relay

1

8.30

8.30

Autopilot Gyro - Horizontal

1

41.30

41.30

Servo Amplifier

1

2.60

2.60

Notes

AC Equipment

Total AC Load

76.20

NOTE 1. Energized when the bleed air switch is set to OFF. 2. Energized when the avionics master switch is set to OFF.

24-50-00

Page 9 Nov 1/09

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

ELECTRICAL POWER ELECTRICAL LOAD DISTRIBUTION MAINTENANCE PRACTICES

200200

1. PROCEDURES Open limiters and diodes cannot be detected easily during the normal operation of the airplane electrical system. However, the voltmeters can be used for operational checks of the electrical system. The following procedures should be performed after electrical system maintenance has been performed. Ensure that all electrical equipment, including avionics, is turned off.

A. Bus Conformity Check (1) Set the battery switch to the OFF position and read battery voltage by placing the voltmeter select switch to the BAT position. The voltmeter should indicate zero volts for all other positions of the switch. (2) Set the battery switch to the ON position and the GEN TIES switch in the OPEN position. The voltmeter should indicate battery voltage when the select switch is set to the BAT position, or slightly less than battery voltage when the switch is set to the TRIPLE BUS position. Select the CENTER BUS position; the voltmeter should indicate battery voltage if the battery bus tie is closed. Select the L GEN and the R GEN positions; each position should result in a zero voltage indication if the bus ties are open. (3) Set the GEN TIES switch to the CLOSE position. The MAN TIES CLOSE annunciator (green) should illuminate. The R GEN TIE OPEN and the L GEN TIE OPEN annunciator lights (yellow) should extinguish. The voltmeter should indicate battery voltage corresponding to the BAT position, the L GEN position and the R GEN position of the voltmeter select switch. Select the EXT PWR position; the voltmeter should indicate zero volts. Both loadmeters should indicate zero percent. (4) Set the BUS SENSE switch to the TEST position and release. The annunciators labeled L GEN TIE OPEN, R GEN TIE OPEN, and BATT TIE OPEN should illuminate. Set the voltmeter select switch to the BAT position; the voltmeter should indicate battery voltage. Select the TRIPLE BUS position; the voltmeter should indicate slightly less than battery voltage. The voltmeter should indicate zero volts for all other positions of the select switch. (5) Set the BUS SENSE switch to the RESET position and release. The annunciators labeled L GEN TIE OPEN, R GEN TIE OPEN, and BATT TIE OPEN should extinguish. (6) Start the RH engine. Set the RH generator switch to the RESET position, then to the ON position after RH engine power is set to HIGH IDLE. All busses should now be powered. (7) Rotate the voltmeter select switch to L GEN and R GEN. The voltmeter should indicate 28.25 ± 0.25 volts corresponding to each switch position. Select the BAT position; the voltmeter should indicate generator voltage. (8) Start the LH engine. Set the LH generator switch to the RESET position, then to the ON position after setting LH engine power to HIGH IDLE. RH load meter and LH load meter indications should be approximately equal. (9) Turn on electrical equipment sufficient to result in a 50% load indication on the RH load meter.

24-50-00

Page 201 Nov 1/09

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (10) Set the GEN TIES switch to the OPEN position. The annunciator lights labeled L GEN TIE OPEN and R GEN TIE OPEN should illuminate. (11) Set the LH generator switch and the battery switch to OFF. Rotate the voltmeter select switch to the L GEN position. The voltmeter should indicate zero volts. Select the R GEN position; the voltmeter should indicate 28.25 ± 0.25 volts. Select the BAT position; the voltmeter should indicate battery voltage. (12) Set the LH generator switch to RESET and then to ON. Set the RH generator switch to OFF. The voltmeter should indicate battery voltage as in Step (11). Set the select switch to L GEN; the voltmeter should indicate 28.25 ± 0.25 volts. Set the select switch to R GEN; the voltmeter should indicate 0 volts. Set the GEN TIES switch to the NORM position. The bus ties should close automatically and the voltmeter should indicate 28.25 ± 0.25 volts DC. (13) Set the RH generator switch to RESET and then to ON. Set the GEN TIES switch to OPEN and open the circuit breakers labeled NO. 1 INV PWR SEL and NO. 2 INV PWR SEL (located on the RH circuit breaker panel). Set the inverter select switch to the NO. 1 position, the NO. 2 position, and the OFF position. Zero voltage corresponding to each switch position should be indicated on the AC voltage/frequency meter. (14) Set the GEN TIES switch to the NORM position. Ensure that the inverter select switch is in the OFF position. The AC voltage/frequency meter should indicate zero volts and the INSTR INV annunciator (red) should illuminate. (15) Set the inverter select switch to the NO.1 position. The AC voltage/frequency meter should indicate 115 vac and 400 Hz. The meter should also indicate 115 vac and 400 Hz for the No.2 position. Turn the inverter select switch OFF. (16) Reset the NO.1 INV PWR SEL and the NO.2 PWR SEL circuit breakers. Set the inverter select switch to NO. 1 and NO. 2. The AC voltage/frequency meter should indicate 115 vac corresponding to each position of the inverter switch. The torquemeter and other equipment utilizing 26 vac should be operational.

Page 202 Nov 1/09

24-50-00

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

B. Triple Fed Bus Diodes, Operational Check The circuit of each power source that feeds the triple fed bus incorporates a blocking diode mounted on a heat sink assembly. It is important to verify that continuity, in one direction only, exists through each diode to ensure that the triple fed bus can be fed by each power source individually, if necessary. The following operational check should be performed to coincide with periodic detailed inspections of the airplane. (1) Set the battery switch to ON and set the GEN TIES switch to OPEN. Set the voltmeter select switch to TRIPLE BUS. The voltmeter should indicate slightly less than battery voltage. Press the annunciator test switch; the annunciator lights should illuminate. (2) Set the voltmeter select switch to the R GEN and L GEN positions, these buses should read zero volts if the diodes have not shorted. (3) Start the RH engine. Set the RH generator switch to the RESET position and release to the ON position. Set the battery switch to OFF. The triple bus should read slightly less than generator voltage. Turn the battery switch ON for LH engine start. (4) Start the LH engine. Set the LH generator switch to RESET and release to ON. Set the RH generator switch in the OFF position. The triple bus should read slightly less than generator voltage. Press the annunciator test switch; the annunciator lights should illuminate.

24-50-00

Page 203 Nov 1/09

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

CHAPTER 25 - EQUIPMENT/FURNISHINGS TABLE OF CONTENTS SUBJECT

PAGE

FLIGHT COMPARTMENT - SEATS 25-10-00 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Procedure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 201 201

PASSENGER COMPARTMENT - SEATS 25-20-00 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Track Wear Limits . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Corrosion on the Seat Track . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Sidewall Folding Passenger Seat . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Bench Type Passenger Seat . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Passenger Seat Cover . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Passenger Seat Bottom Cover . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 201 201 201 201 204 204 204 204 204 204 206 206 206 208 208 208

PASSENGER COMPARTMENT - CARPET 25-20-01 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Removal and Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201

PASSENGER COMPARTMENT SIDEWALL UPHOLSTERY 25-20-03 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Escape Hatch Upholstery . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .201

25-CONTENTS

Page 1 May 1/11

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

CHAPTER 25 - EQUIPMENT/FURNISHINGS TABLE OF CONTENTS (CONTINUED) SUBJECT

PAGE

EMERGENCY LOCATOR TRANSMITTER (ELT) 25-60-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Maintenance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Narco Battery Replacement . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Dorne and Margolin Battery Replacement . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 203 Testing the Emergency Locator Transmitter . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 204

Page 2 May 1/11

25-CONTENTS

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

List of Effective Pages CH-SE-SU

PAGE

DATE

25-LOEP

1

May 1/11

25-CONTENTS

1 and 2

May 1/11

25-10-00

201

Nov 1/09

25-20-00

201 thru 208

Nov 1/09

25-20-01

201 and 202

Nov 1/09

25-20-03

201 and 202

May 1/11

25-60-00

1 and 2 201 thru 204

Nov 1/09 Nov 1/09

C5

25-LOEP

Page 1 May 1/11

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

EQUIPMENT/FURNISHINGS FLIGHT COMPARTMENT - SEATS MAINTENANCE PRACTICES

25-10-00 200200

1. PROCEDURE A. Removal NOTE: On airplanes with the half length curtains, the flight compartment seats are moved aft for removal. On airplanes with the optional solid partition, the seats must be moved forward for removal. Before removing the pilot's seat, lower the seat to its lowest position and place the left armrest down prior to removal. If the pilot's seat is not lowered, or if the left armrest is up when the seat is moved forward, components of the fuel control panel may be damaged. (1) Remove the fire extinguisher from the floor behind the pilot's seat. (2) Remove the seat stops at the end of each seat track. (3) Move the seat forward (or aft) until it clears the mounting track. NOTE: On airplanes with the optional solid partition, the seats will not clear the compartment door in the upright position. To remove a seat from the compartment, place the armrests up, turn the seat on its side, and work it through the door.

B. Installation (1) Place the seat in position and align the seat guide with the mounting tracks. CAUTION: Lower the seat to its lowest position and place the left armrest down before sliding the seat into position. Take care not to force the seat against the fuel control panel during installation. (2) Release the fore and aft adjustment lock and slide the seat into the desired position. Engage the fore and aft lock making certain it holds the seat firmly in place. (3) Replace the seat stops at the end of each track. (4) Replace the fire extinguisher behind the pilot's seat.

25-10-00

Page 201 Nov 1/09

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

EQUIPMENT/FURNISHINGS PASSENGER COMPARTMENT - SEATS MAINTENANCE PRACTICES

25-20-00 200200

1. PROCEDURES A. Track Wear Limits As the seat tracks wear, their strength is reduced. Refer to Figure 201 for acceptable wear limitations which have been established for the passenger compartment seat tracks. If gouging or chaffing is present and the depth does not exceed the values shown in Figure 201, proceed with the following repair procedure. •

Remove any sharp edges using aluminum oxide sandpaper. Do not sand any deeper than the initial gouge or chafe.

•

Apply Alodine 1200 or 600 on any exposed surface.

If the minimum and maximum limitations shown in Figure 201 can not be met, contact Hawker Beechcraft Corporation Technical Support.

B. Corrosion on the Seat Track If corrosion is found on the seat tracks, refer to Chapter 20-09-00 for information on corrosion removal. Inspect each seat track. Refer to Figure 201 for acceptable corrosion limitations, which have been established for the passenger compartment seat tracks. If the minimum and maximum limitations shown in Figure 201 can not be met, contact Hawker Beechcraft Corporation Technical Support.

C. Removal CAUTION: Care should be taken when removing the passenger compartment seats. Excessive damage to the seat track may lead to a costly seat track repair or replacement. (1) Unlock the seat from the floor and side track at the sleeved retainers built into the chair legs (Ref. Figure 202). (2) Remove the seat from the track.

D. Installation (1) Inspect the seat tracks for wear and corrosion. Refer to the PASSENGER SEAT TRACK WEAR LIMITS and CORROSION ON THE SEAT TRACKS procedures in this section (Ref. Figure 202). CAUTION: Care should be taken when installing the passenger compartment seats. Excessive damage to the seat track may lead to a costly seat track repair or replacement. (2) Locate forward foot of seats at index points painted on the floor and side tracks. (3) Lock the seat into the tracks with the sleeved retainers built into the chair leg.

25-20-00

Page 201 Nov 1/09

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Y

SEAT TRACK LUG REF

.45 OVER LUG REF

Y 0.03 WIDE BY 0.06 DEEP MAX IN CUTOUT AREA NOT OVER LUG

MAX 0.050 WEAR COMBINED ALLOWED ON UPPER AND LOWER SURFACE e.g. AS SHOWN BELOW: 0.02 + 0.03 = 0.050 MAX

0.050 MAX WEAR ALLOWED

0.065 MIN (0.020 IN FROM EDGE) 0.03 REF 0.297 MIN

0.02 REF 0.02 REF (IN FROM EDGE)

0.03 MAX

VIEW

Y-Y

Figure 201 Passenger Compartment Seat Track Wear Limits

Page 202 Nov 1/09

25-20-00

UE53B 990997AA.AI

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 202 Passenger Seat Installation

25-20-00

Page 203 Nov 1/09

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

2. SIDEWALL FOLDING PASSENGER SEAT A. Removal CAUTION: Care should be taken when removing the passenger compartment seats. Excessive damage to the seat track may lead to a costly seat track repair or replacement. (1) Remove the connecting bolts from the side mounting rings on the airplane sidewall. (2) Unlock the seat from the floor track at the sleeved retainer built into the chair leg. (3) Remove the seat from the floor tracks.

B. Installation (1) Inspect the seat tracks for wear and corrosion. Refer to the PASSENGER SEAT TRACK WEAR LIMITS and CORROSION ON THE SEAT TRACKS procedures in this section. CAUTION: Care should be taken when installing the passenger compartment seats. Excessive damage to the seat track may lead to a costly seat track repair or replacement. (2) Locate forward foot of seats at the index points painted on the seat tracks. (3) Install the connecting bolts in the side mounting ring on the airplane sidewall. (4) Lock the seat into the floor track with the sleeved retainer built into the chair leg.

3. BENCH TYPE PASSENGER SEAT A. Removal CAUTION: Care should be taken when installing the passenger compartment seats. Excessive damage to the seat track may lead to a costly seat track repair or replacement. (1) Remove the screws and washers that secure the seat back to the rear bulkhead (Ref. Figure 203). (2) Unlock the seat from the floor and side tracks at the sleeved retainer built into the seat. (3) Remove the seat from the tracks.

B. Installation (1) Inspect the seat tracks for wear and corrosion. Refer to the PASSENGER SEAT TRACK WEAR LIMITS and CORROSION ON THE SEAT TRACKS procedures in this section. CAUTION: Care should be taken when installing the passenger compartment seats. Excessive damage to the seat track may lead to a costly seat track repair or replacement. (2) Locate forward foot of seats at index points painted on the seat tracks (Ref. Figure 203). (3) Install the screws and washers that secure the seat back to the rear bulkhead. (4) Lock the seat into the floor track with the sleeved retainer built into the seat.

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Figure 203 Bench Type Passenger Seat Installation

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4. PASSENGER SEAT COVER A. Removal CAUTION: The seat cover snaps are designed with a tab that only allows them to be released by lifting on the side nearest the center line of the seat. Care should be taken when unsnapping the seat cover. It is possible to tear the snaps loose from the seat cover if the cover is removed improperly. NOTE: There are four (4) snaps and several strips of velcro holding the seat back cover to the seat frame. (1) Reach under the seat from the aft side and separate the outer cover from the velcro that is attached to the seat bottom (Ref. Figure 204). (2) Unsnap the lower two snaps on the back of the seat by putting your finger under the cover and pushing out and away from the center line of the seat. (3) If a fold-down table is installed on the back of the seat: (a) Release the table latch and lower the table. (b) Remove the two screws securing the table latch to the seat and remove the table latch. (4) Separate the cover from the velcro on the back of the seat up to the upper two snaps. (5) Unsnap the upper two snaps on the back of the seat by putting your finger under the cover and pushing out and away from the center line of the seat. (6) Reach under the seat from the aft side and separate the inner cover from the velcro that is attached to the seat bottom. (7) Remove the cover from the seat back.

B. Installation NOTE: The seat cover snaps are designed with a tab that only allows them to be engaged on the side farthest from the center line of the seat then pushed inward toward the center line of the seat for total engagement. (1) Pull the seat cover down over the seat back (Ref. Figure 204). (2) If a fold-down table is installed on the back of the seat: (a) Position the cover so that the holes for the table latch are aligned properly. (b) Install the table latch with the two screws removed during the removal procedure. (3) Engage the two upper snaps on the edge farthest from the center line of the seat and push them inward toward the center line of the seat. (4) Press the cover onto the velcro on the back of the seat down to the two lower snaps.

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(5) Raise the fold-down table and latch it in the upright position. (6) Engage the two lower snaps on the edge farthest from the center line of the seat and push them inward toward the center line of the seat. (7) Reach under the seat from the aft side and press the inner cover, then the outer cover onto the velcro attached to the seat bottom.

Figure 204 Seat Cover Snap Release and Engagement Procedures

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5. PASSENGER SEAT BOTTOM COVER A. Removal (1) Pull the seat bottom loose from the velcro attaching it to the seat frame. (2) Pull the seat bottom cover loose from the velcro attaching it to the seat bottom. (3) Remove the cover from the seat bottom.

B. Installation (1) Position the seat bottom cover over the seat bottom. (2) Press the edges of the seat bottom cover onto the velcro attach areas. (3) Press the seat bottom into place on the velcro attached to the seat frame.

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EQUIPMENT/FURNISHINGS PASSENGER COMPARTMENT - CARPET MAINTENANCE PRACTICES

25-20-01 200200

1. PROCEDURES A. Removal and Installation Carpeting installed in the passenger compartment is secured in position along the side edges with strips of 445 2 inch-wide, double-faced tape (182, Table 1, Chapter 91-00-00) installed on the back of the carpet (Ref. Figure 201). In addition, the ends of the longer carpeting is secured to the floor with short strips of Dual-Lock (hook-loop) fastener material. The main spar cover is attached with adhesive (16, Table 1, Chapter 91-00-00). If it becomes necessary to install new tape or resecure the Dual-Lock fastener strips, all surfaces which the tape contacts must be free of dust, oil, fingerprints or other contaminants soils. Prime the carpet backing with adhesive (183, Table 1, Chapter 91-00-00) prior to the installation of the Dual-Lock fasteners. Clean all contacting metal surfaces with acetone (24, Table 1, Chapter 91-00-00), solvent (54, Table 1, Chapter 91-00-00), solvent (14, Table 1, Chapter 91-00-00) or an approved alkaline cleaning solution. The passenger compartment entryway and the ramp over the main spar are coated with an anti-skid paint coating (184, Table 1, Chapter 91-00-00).

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BAGGAGE COMPARTMENT

FS FS 143.00 150.60

FS 175.60

DUAL-LOCK FASTENER STRIPS (10 PLACES) AT FORWARD AND AFT ENDS OF CARPET AS SHOWN

FS 290.50

#445 TAPE AROUND EDGES OF CARPET (TYPICAL ALL CABIN CARPET SECTIONS)

FS 451.00

BL 00.00

A

D F CARPET ASSY

E

C

F CARPET ASSY

B FS 290.50

DUAL-LOCK FASTENER STRIPS

ANTI-SKID WALKWAY

DUAL-LOCK FASTENER STRIPS (10 PLACES) AFT FORWARD AND AFT ENDS OF CARPET AS SHOWN

SEAT TRACK

DETAIL

F

TYPICAL (2 PLACES) (UB-39 AND AFTER)

DETAIL (UB-1 THRU UB-38)

DETAIL

E

WL 84.50

FS 143.00

SPAR COVER

WL 84.50

WL 84.50 DETAIL

D

FS 290.50

DETAIL

C

DETAIL

B UC25B 061536AA.AI

Figure 201 Passenger Compartment Carpet Installation

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25-20-03 200200

EQUIPMENT/FURNISHINGS PASSENGER COMPARTMENT SIDEWALL UPHOLSTERY MAINTENANCE PRACTICES 1. ESCAPE HATCH UPHOLSTERY A. Removal (1) Perform the EMERGENCY EXIT DOOR REMOVAL procedure (Ref. Chapter 52-20-00). (2) Remove screws (5 and 12) and remove side trim (6 and 11) (Ref. Figure 201). (3) Remove screws (2 or 3) and remove top trim assembly (1 or 4) whichever is installed. (4) Remove four screws (8) and cover (9). (5) Remove escutcheon assembly (10).

B. Installation (1) Install escutcheon assembly (10) (Ref. Figure 201). (2) Install cover (9) with four screws (8). (3) Install top trim assembly (1 or 4) with screws (2 or 3) whichever is installed. (4) Install side trim (6 and 11) with screws ( 5 and 12). (5) Perform the EMERGENCY EXIT DOOR INSTALLATION procedure (Ref. Chapter 52-20-00).

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Figure 201 Escape Hatch Upholstery

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EQUIPMENT/FURNISHINGS EMERGENCY LOCATOR TRANSMITTER (ELT) DESCRIPTION AND OPERATION

25-60-00 00

1. GENERAL The Model 1900 Airliner is equipped with either a NARCO or a DORNE AND MARGOLIN emergency locator transmitter (ELT) to assist in tracking and recovery of the airplane, crew, and passengers in the event of a crash or emergency landing. The ELT is mounted on the right side of the fuselage at a point just forward of FS 598. Access to the ELT for replacement or repair is gained by removing the access panel located below the right stabilon. A spring-loaded door is installed adjacent to the transmitter to provide access for manual activation of the ELT. The antenna for the ELT is mounted on top of the fuselage under the dorsal fin at FS 570.107. The output frequencies of the ELT are 121.5 and 243.0 MHz, simultaneously. Range is approximate line of sight. The ARM-ON-OFF or TEST-AUTO-XMIT switch is located on the transmitter, and controls the operation of the set. The ON or TEST position turns the set on for testing (the TEST switch is a momentary on switch and will automatically turn off when released, the XMIT switch will continue to transmit a signal as long as the switch is in this position), and the ARM or the AUTO position actuates the set to operate the set automatically upon impact. A reset switch, located on the forward end of the transmitter, resets the transmitter in the event the impact switch is accidentally triggered. An optional remove ARM-ON-Off or TEST-AUTO-XMIT switch may be located on the instrument panel or in the aft fuselage on the right side for use during ground testing of the ELT (Ref. Figure 1).

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Figure 1 Emergency Locator Transmitter Installation

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EQUIPMENT/FURNISHINGS EMERGENCY LOCATOR TRANSMITTER (ELT) MAINTENANCE PRACTICES

200200

1. PROCEDURES A. Maintenance Maintenance on the ELT is normally limited to replacing the battery. The following is a list of various conditions which warrant battery replacement: (1) Visual inspection shows signs of leakage, corrosion or insecure leads. (2) Elapsed replacement date noted on the battery case. (This date represents 50 percent of the useful life of the battery.) NOTE: The useful life of the battery is the length of time the battery may be stored without losing its ability to continuously operate the ELT for 48 hours. (3) After any emergency use. (4) After one cumulative hour of use. (5) After operation of unknown duration. (6) If transmitter is stored in an area where the temperature is normally above 38°C (100°F.) the battery should be replaced every 12 months. CAUTION: Avoid storing the batteries at temperatures in excess of 55°C (130°F.) The information on battery life and replacement is included in the data furnished with each ELT, and is usually placarded on the battery. NOTE: Replacement batteries should be obtained only from ELT and airplane manufacturers or other acceptable suppliers, since the condition and useful life of over-the-counter batteries, such as those sold for flashlights, portable radios etc., are usually unknown. The ELT switch should not be turned to any activated position unless the ELT is connected to its associated antenna or a 50 ohm dummy load.

B. Narco Battery Replacement (1) Remove the access panel located just below the right stabilon. (2) If equipped with the ARM-OFF-ON switch, place the switch in the OFF position. If equipped with the TEST-AUTO-XMIT switch, ascertain that the switch is in the TEST position for automatic off (Ref. Figure 201). (3) Disconnect the antenna cable from the ELT. (4) Disconnect the remote switch wiring (if installed) from the terminals on the ELT. (5) Unlatch the mounting strap and remove the ELT from the airplane. (6) Extend the portable antenna.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL CAUTION: Exercise extreme care in extending the portable antenna and handling the control head during the battery replacement to avoid damage to the antenna or the plastic tab on the antenna. (7) Remove the four screws attaching the control head to the battery casing and slide the control head and battery case apart. The battery connection leads are approximately 3 inches long. (8) Disconnect the battery by unsnapping the battery terminals from the bottom of the transmitter PC board. Discard the old battery. WARNING: DO NOT DISCARD THE OLD BATTERY IN FIRE. (9) Connect the terminals of the new battery to the bottom of the transmitter PC board. (10) Using a stick, apply a small bead of sealant (supplied with battery pack) around the area of the control head which is joined with the battery case during assembly. NOTE: This sealant provides a watertight seal when the unit is assembled. (11) Insert the control head section into the battery case, being careful not to pinch the wires, and install the four attaching screws. Wipe any excess sealant from the outside of the unit. NOTE: If the four holes do not line up, rotate the battery case 180° and reinsert. (12) Stow the portable antenna. CAUTION: Exercise extreme care in order to avoid damage to the antenna or the plastic tab on the antenna. (13) Install the transmitter in the airplane and secure the mounting strap. (14) Connect the fixed antenna cable to the ELT. Ensure that the plastic contact separator is inserted between the portable antenna contact and the portable antenna. NOTE: Should the contact separator be improperly installed, a very weak signal may be transmitted. This signal may be strong enough for a functional test, but too weak for emergency use. (15) Connect the remote switch wiring, (if installed), to the terminals on the ELT. (16) Press the RESET button and place the ARM-ON-OFF or the TEST-AUTO-XMIT switch on the ELT in the ARM or the AUTO position. (17) A new replacement date must be marked on the outside of the transmitter. (18) If transmitter is stored in an area where the temperature is normally above 38°C (100°F.) the battery should be replaced every 12 months. This date represents 50% of the useful life of the battery as defined by the battery manufacturer. (19) Position the access panel over its opening and secure.

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Figure 201 Narco ELT

C. Dorne and Margolin Battery Replacement (1) Remove the access panel located just below the right stabilon. (2) Make certain the ELT is turned off. (3) Disconnect the antenna. (4) Disconnect the remote switch. (5) Remove the screws holding the ELT in place. (6) Remove the screws from the bottom of the ELT and remove the bottom. (7) Remove the screws from the bottom of the ELT and remove the bottom. (8) Disconnect the battery and discard it. NOTE: Inspect for and properly treat any corrosion that may be indicated in the area when the battery is replaced. (9) Connect the new battery. (10) Install the bottom of the ELT and the screws.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (11) Install the ELT in the airplane and install the screws which hold it in place. (12) Connect the antenna and remote switch. (13) Install and secure the access panel. (14) Test the ELT as indicated in the TESTING EMERGENCY LOCATOR TRANSMITTER procedure.

D. Testing the Emergency Locator Transmitter Generally, tests will be performed following maintenance or repairs of ELTs, other than battery replacement, to determine their operational capability. Testing the ELT, if improperly done, could trigger false alerts and create frequency jamming, and may interfere with the reception of a bonafide emergency transmission. Federal Communications Commission regulations require that testing be performed in a screened or shielded test room, or in a test enclosure that will hold the self contained ELT unit with the antenna fully extended. CAUTION: The elt switch should not be activated to any position unless the elt is connected to its associated antenna or a 50 ohm dummy load. Operational testing of installed ELTs may be accomplished as follows: NOTE: Tests should be no longer than three audio sweeps. One audio sweep may be defined as amplitude modulating the carrier with an audio frequency sweeping downward over a range of not less than 700Hz, within the range of 1600 to 300Hz, and a sweep repetition rate between two and four Hz. Tests should be conducted only in the first five minutes after the hour. The tests should be coordinated with the nearest FAA tower or flight service station. (1) Turn on the VHF transceiver Comm-1 and tune the transceiver to 121.5 MHz. (2) Turn the COMM-1 audio switch to the SPEAKER position and place the volume control in the center of its range. (3) Turn the ELT ARM-ON-OFF switch to ON or XMIT ON and monitor the ELT signal. On airplanes equipped with an optional remote switch (located in the aft fuselage on the right side or on the instrument panel), the switch may be flipped to the XMIT position and the ELT signal monitored. NOTE: A distinctive downward sweeping tone should have been heard from the monitoring receiver during the test. If the tone was heard, the ELT is functioning properly. If there was no tone, (assuming that the VHF transceiver is operational), the battery is probably disconnected, outdated or discharged. (4) Place the ARM-OFF-ON switch on the ELT to the OFF or the TEST-AUTO-XMIT switch on the ELT to the TEST position. If the remote switch is being utilized to test the unit, the switch should be returned to the AUTO position. The audio signal should disappear completely. (5) Place the ARM-ON-OFF or the TEST-AUTO-XMIT switch on the ELT to the ON or the AUTO position. There should be no audio signal present. NOTE: If a signal is heard, the impact switch has probably been activated and should be reset. (6) Firmly press the reset switch on the front of the ELT and listen to ensure the audio signal disappears from COMM-1.

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CHAPTER 26 - FIRE PROTECTION TABLE OF CONTENTS SUBJECT

PAGE

FIRE DETECTION SYSTEM 26-10-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Troubleshooting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 101 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 101 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Fire Detection Cable . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Fire Detection System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Checkout Procedures with Jet-Cal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Checkout Procedures without Jet-Cal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202

ENGINE BLEED AIR WARNING SYSTEM 26-11-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Pressure Switch . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Check For Proper Electrical Connection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Tubing Repair . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201

FIRE EXTINGUISHING SYSTEM 26-20-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Supply Cylinder . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Recharging . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Extinguisher Activation Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 203 Extinguisher Cartridge (SQUIB) and Supply Cylinder . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 206 Information . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 206 Bleed Air Warning Switches . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 206 Check For Proper Electrical Connection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 206

26-CONTENTS

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List of Effective Pages CH-SE-SU

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DATE

26-LOEP

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26-10-00

1 and 2 101 201 thru 203

Nov 1/09 Nov 1/09 Nov 1/09

26-11-00

1 201 thru 203

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26-20-00

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FIRE PROTECTION FIRE DETECTION SYSTEM DESCRIPTION AND OPERATION

26-10-00 00

1. GENERAL A fire detection system is installed to provide immediate warning in the event of a fire in the engine compartment. The main elements of the system are the two thermal sensitive fire detection cables which are joined at one end and routed about the engine in a continuous loop (Ref. Figure 1). The routing of the cable has been determined in order to monitor various sections of the engine or other temperature sensitive regions in which critical situations could develop. Arrangement of the fire detection cables in a continuous loop enables the system to be monitored for open circuit conditions and should a single break in the cable occur, the system would remain operational. Either a rotary switch or two toggle switches provide for testing of the fire detection system during pre-take off checks. Electrical power for the fire detection system is supplied by a 5-amp circuit breaker placarded FIRE DETECT, located on the right side panel. Should the airplane be equipped with the two toggle switches placarded ENG FIRE TEST - DETECT, one for the LEFT system and one for the RIGHT system, the switches are three position switches spring loaded to the center. The switch positions are placarded LOOP - OFF - AMP. When either toggle switch is placed in the LOOP position, the integrity of the appropriate fire detection cable is tested. A good test is indicated by the red lights in the appropriate FIRE PULL “T” handle being illuminated. When either toggle switch is placed in the AMP position, the integrity of the circuitry within the control amplifier is tested. A good test is indicated by the red lights in the appropriate FIRE PULL “T” handle being illuminated as in the loop test. On airplanes equipped with the rotary test switch placarded TEST SWITCH - FIRE DET FIRE EXT, the right and left systems are checked simultaneously by selecting either the FIRE position or the FAULT position. Selecting the FIRE position tests the integrity of the circuitry within the amplifier. A good test is indicated by the illumination of the red lights in the FIRE PULL “T” handles. Selecting the FAULT position tests the integrity of the fire detection loops. A good test in this position is indicated by the illumination of the red lights in the FIRE PULL “T” handles as before. The two fire detection cables, one forward and one aft, are designed as a single center wire surrounded by a semiconductive material and enclosed within a stainless steel sheath. As temperature increases, the semiconductive material separating the center wire from the outer sheath becomes more conductive and the resistance between the center wire and the outer sheath decreases. The outer sheath is grounded to the airplane structure. A control unit, mounted on the forward pressure bulkhead behind the LH subpanel, monitors resistance between ground and the center wire of the fire detection cables and is set to trip when resistance between the center wire of the cables and the sheath (ground) decreases to a certain point (approximately 100 ohms). Once the control unit trips, a warning light in the firewall fuel shutoff valve handle illuminates, signaling a fire in the engine compartment. The control unit is of the short circuit discriminating type. The short circuit discriminator is designed to actuate when the system resistance decreases rapidly from the normally high resistance of the cable.

26-10-00

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Figure 1 Fire Detection System

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FIRE PROTECTION FIRE DETECTION SYSTEM TROUBLESHOOTING

100100

1. PROCEDURES Any time a fault is indicated by the cockpit test procedure, one of the two checkout procedures under MAINTENANCE PRACTICES should be performed. Failure to perform one of these procedures may result in identifying a failure in the wrong component. Should the checkout procedure fail to reveal a fault in the fire detection loop, the fault may be assumed to be in the control unit or its associated circuitry. When it is determined that no fault exists in the circuitry, including the test switch circuits, the control unit should be replaced. While the following troubleshooting table may not indicate all possible fault conditions, the most common faults are listed in order of their most probable occurrence. Table 101 TROUBLESHOOTING - FIRE DETECTION SYSTEM PROBLEM 1. Failure of fire detection system to test.

PROBABLE CAUSE

CORRECTIVE ACTION

a. Warning lamp inoperative

a. Check test circuit continuity and warning bulb.

b. fire detection cable fails to operate properly.

b. Replace fire detection cable.

c. Control unit fails to operate properly.

c. Replace control units.

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FIRE PROTECTION FIRE DETECTION SYSTEM MAINTENANCE PRACTICES

200200

1. FIRE DETECTION CABLE A. Removal (1) Pull the FIRE DET circuit breaker on the circuit breaker panel. (2) Remove the engine cowling (Ref. Chapter 71-10-00). (3) Remove power from the airplane. (a) Perform REMOVING GROUND POWER procedure (Ref. Chapter 24-40-00). (b) Perform BATTERY DISCONNECTION procedure (Ref. Chapter 24-31-00). (4) Remove the nuts, bolts, screws, and clamps throughout the length of the fire detection cable. (5) Disconnect the electrical connector from the fire detection responder switch in the engine accessory area. (6) Remove the fire detection cable from the airplane.

B. Installation (1) Install the fire detection cable on the engine using the nuts, bolts, screws, and clamps to properly secure the cable in place. NOTE: Use care when installing the fire detection cable. Do not form loops or bends in the cable with less than a 2-inch-bend radius. Be certain that clearance is allowed between the cable and surrounding structure. (2) Install the electrical connector on the fire detection responder switch in the engine accessory area. (3) Apply power to the airplane. (a) Perform BATTERY CONNECTION procedure (Ref. Chapter 24-31-00). (b) Perform APPLYING GROUND POWER procedure (Ref. Chapter 24-40-00). (4) Reset the FIRE DET circuit breaker on the circuit breaker panel. (5) Perform FIRE DETECTION SYSTEM CHECKOUT PROCEDURES as outlined in this Chapter.

2. FIRE DETECTION SYSTEM A. Checkout Procedures with Jet-Cal (1) Remove the engine cowling. (Ref. Chapter 71-10-00). (2) Connect the Jet-Cal tester to a suitable source of electrical power, (115vac 50 to 400 Hz).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (3) Make preliminary test set up of the Jet-Cal tester using the appropriate setup procedures indicated on the cover of the unit. (4) Attach the BH1278 heater probe to the forward fire detection cable. (5) Slowly increase the temperature of the heater probe and verify that the FIRE light in the firewall fuel shutoff valve handle illuminates at a temperature of 315°C to 483°C (600°F to 900°F). (6) Reduce the temperature of the heater probe to 0°C; verify that the FIRE light in the firewall fuel shutoff valve handle extinguishes. (7) Disconnect the heater probe from the forward fire detection cable. (8) Attach the heater probe to the aft fire detection cable. (9) Slowly increase the temperature of the heater probe and verify that the FIRE light in the firewall fuel shutoff valve handle illuminates at a temperature of 248°C to 382°C, (480°F to 720°F). (10) Reduce the temperature setting of the heater probe to 0°C, and verify that the FIRE light in the firewall fuel shutoff valve handle extinguishes. (11) Disconnect the heater probe from the aft fire detection cable. (12) Repeat the test procedure for the other engine. (13) Secure the test set. (14) Install the engine cowling (Ref. Chapter 71-10-00).

B. Checkout Procedures without Jet-Cal (1) Remove the engine cowling (Ref. Chapter 71-10-00). (2) Disconnect the engine harness plug P1 from the firewall electrical connector. Refer to the 1900C Wiring Diagram Manual. (3) Using an ohmmeter, measure the resistance between pins M and P of the plug (UA-1 thru UA-3, UB-1 thru UB-53) and between pins A and B (UB-54 and after and UC-1 and after). This resistance must not exceed 10.9 ohms. Resistance in excess of 10.9 ohms would be indicative of a poor connection or an open fire detection cable center wire. (4) Locate the fire detection interconnect fittings which pass through the lower RH side of the fireseal. (5) Remove the potting material from the fittings and separate the two fire detection cables. (6) Using a good quality digital multimeter capable of reading megohms, measure the resistance between pin M of the firewall plug P1 (UA-1 thru UA-3, UB-1 thru UB-53) or pin B (UB-54 and After and UC-1 and After) and the outer sheath of the aft fire detection cable. Measure the ambient temperature. The measured resistance must not be less than the amount indicated. Refer to Table 201 for the current ambient temperature to associated resistance values. Insufficient resistance will necessitate replacing the aft fire detection cable.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (7) Measure the resistance between pin P of the firewall plug P1 (UA-1 thru UA-3, UB-1 thru UB-53) or pin A (UB-54 and After and UC-1 and After) and the outer sheath of the forward fire detection cable. The ambient temperature must be between 50°F and 100°F. This resistance must not be less than 1.7 megohms. A lower resistance will necessitate replacing the forward fire detection cable. (8) Connect the two fire detection cables and replace the sealer (158, Table 1, Chapter 91-00-00) as necessary. (9) Connect firewall plug P1 to the firewall connector. (10) Install the engine cowling (Ref. Chapter 71-10-00). Table 201 AFT FIRE DETECTION CABLE RESISTANCE Ambient Temperature Degrees Fahrenheit

Minimum Resistance Megohms

50 to 68

0.68

72

0.56

76

0.47

80

0.39

84

0.33

88

0.28

92

0.23

96

0.20

100

0.16

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FIRE PROTECTION ENGINE BLEED AIR WARNING SYSTEM DESCRIPTION AND OPERATION

26-11-00 00

1. GENERAL The engine bleed air warning system provides a visual indication of a rupture in the bleed air lines so that the bleed air valve from the affected engine can be shut down before heat of the escaping air damages the skin and structure adjacent to the break in the line. The bleed air lines are in close proximity with ethylene vinyl acetate (EVA) tubing from the engines to the cabin. Excessive heat on the tubing, caused by a ruptured bleed air line, will melt the tubing. The system has two pressure switches mounted just forward of the main spar under the center cabin floorboard. The two switches are pressurized by air tapped off the line from the bleed air manifold, mounted under the center floorboard forward of the main spar. Refer to Figure 201 in the MAINTENANCE PRACTICES section for routing of the EVA tubing. When tubing melts and the tubing pressure drops below pressure required to keep the pressure switch actuated, the normally open switch in the line will close, causing a circuit to be completed to the respective BL AIR FAIL light in the warning annunciator panel. When indication of bleed air failure becomes evident, bleed air for that side must be turned off by placing the respective bleed air valve switch in the INST ENVIR OFF position. When the switch is placed in this position, the engine firewall shutoff valve closes, stopping bleed air flow at the engine firewall. Each time the system is actuated, EVA tubing in the area where the damage occurred must be replaced when the bleed air line is repaired.

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FIRE PROTECTION ENGINE BLEED AIR WARNING SYSTEM MAINTENANCE PRACTICES

200200

1. PRESSURE SWITCH A. Removal (1) Remove center floorboards to gain access to the pressure switches located in the center section forward of main spar. (2) Remove the electrical connector from the pressure switch (Ref. Figure 201). (3) Remove the fittings from the pressure switch to be replaced. Cap fittings to prevent entry of foreign matter. (4) Remove attaching screws that secure the switch in place and remove pressure switch.

B. Installation (1) Install the pressure switch in its mounting space and secure with attaching screws (Ref. Figure 201). (2) Remove the cap from the fittings and install the fitting on the pressure switch. (3) Attach electrical connector to the pressure switch.

C. Check For Proper Electrical Connection To confirm that the left and right bleed air warning switches have not become cross connected during maintenance, perform the following check: (1) Remove the center floorboard forward of the main spar to gain access to the pressure switches. (2) Locate the left and right bleed air warning pressure switches [S128] and [S129] (Ref. Figure 201). (3) Locate the switch that connects to the plastic EVA warning line that goes into the RH wing. (4) Confirm that this switch is connected to connector P242 which should contain wire no’s. W2A22 and W4A22.

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D. Tubing Repair If a malfunction in the bleed air system, has occurred and warning system has been activated, the ethylene vinyl acetate (EVA) tubing must be repaired. The damaged area is repaired by splicing with a piece of aluminum tubing no more than four inches long. The splicing aluminum tube (128, Table 1, Chapter 91-00-00) may be 5052 or 6061, 3/16-inch O.D. with 0.028 or 0.035-inch wall thickness. If the EVA tubing has stretched because of aging or creeping, additional standoffs or clamping may need to be installed. Use the same type of standoff or cushion clamp as the original installation. A distance of two to four inches must be maintained between the EVA tube and the source of heat. If the two-inch-minimum distance can not be maintained, the tube may be insulated with the same type of insulation as used in the original installation. Cut the insulation lengthwise, install on the tubing and secure with a TYZ-28M tie wrap. The repair must be accomplished per Kit No.114-9021-1 or as follows: (1) Locate damaged EVA tubing. (2) Remove panels as necessary to gain access. (3) Cut out damaged EVA tubing (not to exceed three inches). (4) Cut aluminum tubing to proper length. NOTE: If Beading the aluminum tubing is not possible, bond tube in place with Uralane 8089. If the splice is bonded in place, exercise care to avoid plugging the tube with adhesive. (5) Bead each end of aluminum tubing. (6) Insert each end of the aluminum tube 1/2 to 3/4 inch into EVA tubing. (7) Clamp each end of tubing in place.

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Figure 201 Bleed Air Warning System

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

FIRE PROTECTION FIRE EXTINGUISHING SYSTEM DESCRIPTION AND OPERATION

26-20-00 00

1. GENERAL A fire extinguisher supply cylinder is mounted on brackets aft of the main spar in each wheel well of each main landing gear. Each cylinder is charged with 2.10 pounds of Bromotrifluoromethane (CBrF3) pressurized to 360 psi +25/ -0 psi at 70°F. The line from the cylinder runs along the side of the nacelle and branches into two spray tubes strategically located about the engine to diffuse the extinguishing agent in the event of a fire. One of the nozzles is positioned to discharge into the engine exhaust area and the other discharges into the accessory area. Once activated, the entire supply of extinguishing agent is discharged. The fire extinguisher control switches used to activate the system are located on the glareshield at each end of the warning annunciator panel. Their power is derived from the hot battery bus through the microswitches mounted on the firewall fuel shutoff valve. The push-to-activate switches incorporate three indicator lights. The red lens, placarded L or R ENG. FIRE PUSH TO EXT., indicates that the firewall fuel shutoff valve has been actuated, thus arming the fire extinguisher pushbutton circuitry. A green lens, placarded OK, is provided only for the test functions for preflight. The amber lens, placarded D, monitors the condition of the pyrotechnic cartridge and status of the cylinder charge. As long as the cartridge is intact and the cylinder is not discharged, the amber light will remain off. When the cartridge is fired, the light will come on and will remain on until the cartridge is replaced. The airplanes are equipped with either a single rotary switch placarded TEST SWITCH - FIRE DETECT FIRE EXT or two toggle switches placarded ENG FIRE TEST - EXT TEST, one for the LEFT system and one for the RIGHT system installed in the copilot's inboard subpanel. These switches are for the purpose of testing the circuitry of the fire extinguisher pyrotechnic cartridges (Ref. Figure 201, Maintenance Practices section). Should the airplane be equipped with the two ENG FIRE TEST toggle switches, the switches are moved to the EXT TEST position while verifying the illumination of the appropriate yellow D light and the appropriate green OK light on each fire extinguisher activation switch on the glareshield. The toggle switches are spring loaded and will return automatically to the center OFF position. On airplanes equipped with the rotary ENG FIRE TEST switch, the pilot should rotate the test switch to each of the two positions (RIGHT EXT and LEFT EXT) and verify the illumination of the appropriate yellow D light and the appropriate green OK light on each fire extinguisher activation switch on the glareshield.

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FIRE PROTECTION FIRE EXTINGUISHING SYSTEM MAINTENANCE PRACTICES

26-20-00 200200

1. SUPPLY CYLINDER A. Removal WARNING: Disconnect battery and disconnect external power before changing out supply cylinders. (1) Set the EXT PWR switch to OFF. (2) Set the BATT switch to OFF. (3) Perform the BATTERY DISCONNECTION procedure (Ref. Chapter 24-31-00). (4) Working through the main gear wheel well, disconnect wiring from supply cylinder terminals (Ref. Figure 201). Refer to the appropriate Model 1900 Airliner Series Wiring Diagram Manual for proper wiring of the supply cylinder (Ref. Model 1900 Airliner Series Wiring Diagram Manual, P/N 114-590032-3 (UA-1 and After) or Model 1900C Airliner Wiring Diagram Manual, P/N 114-590021-13 (UB-1 and After) or Model 1900C Airliner Wiring Diagram Manual, P/N 114-590021-61 (UC-1 and After). (5) Disconnect the tube attached to the bottom of the supply cylinder. (6) Remove the attaching bolts and remove the cylinder from the mounting bracket.

B. Installation WARNING: Disconnect battery and disconnect external power before changing out supply cylinders. NOTE: The supply cylinder comes from the vendor with two orange wires hanging loose. (1) Secure the cylinder to the mounting brackets with the attaching bolts (Ref. Figure 201). (2) Connect the wiring to the supply cylinder. Refer to the appropriate Model 1900 Airliner Series Wiring Diagram Manual for proper wiring of the supply cylinder (Ref. Model 1900 Airliner Series Wiring Diagram Manual, P/N 114-590032-3 (UA-1 and After) or Model 1900C Airliner Wiring Diagram Manual, P/N 114-590021-13 (UB-1 and After) or Model 1900C Airliner Wiring Diagram Manual, P/N 114-590021-61 (UC-1 and After). (3) Connect the tube to the bottom of the cylinder. (4) Perform the BATTERY CONNECTION procedure (Ref. Chapter 24-31-00).

C. Recharging Access to the fire extinguisher cylinder is through the wheel wells of the main landing gear. Each cylinder is charged with 2.10 pounds of Bromotrifluoromethane (CBrF3) and pressurized with dry nitrogen to 360 psi +25/ -0 psi at 70°F. Check the pressure gage on each cylinder prior to flight to ascertain that the cylinders are charged to within the pressure limits for the ambient temperatures (Ref. Table 201).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL CAUTION: Bromotrifluoromethane can cause severe corrosion in and on a hot or cold engine. If the fire extinguisher is activated, and the engine is contaminated with fire extinguishing agent, the engine must be cleaned as outlined under the heading ENGINE WASHING PROCEDURES (Ref. Chapter 12-20-00). Fully discharged cylinders, or any cylinder indicating less than the specified pressure on the pressure gage, should be returned to a recharging station certified by the vendor (HTL - Advanced Technologies, Duarte, Ca.) for proper charge of CBrF3 extinguishing agent. Such stations are located in most major cities in this country. The following is a list of charging stations: List of Charging Stations *Aviation Comprogas Services 10 Railroad Avenue East Northport, L.I., New York 11731 Phone: 516/368-2203

*United Air Lines, Inc., San Francisco Int'l Airport San Francisco, California 94128

City Sales U.S. 2 and Lake Antione Rd. P.O. Box 663 Iron Mountain, Michigan 40801 Phone: 906/774-3555 *Delta Air Lines, Inc., Protection C.F.H., Inc., Jet Maintenance Base 103 Gun Ave. Hartsfield-Atlanta Pointe Claire, Que. International Airport H9R 3X2, CANADA Atlanta, Georgia 30320 Phone: 514/694-3980 Phone: 404/346-6011 Aero Electric, Inc., *CB Enterprises 113 S. Laura 10704 Vanowen Street North Hollywood, California 91605 Wichita, Kansas 67211 Phone: 316/263-0197 Phone: 213/980-2332 *Korean Air Lines Safety Supply Company CP.O. Box 864 172 Osborne Street, South Kimpo International Airport Winnipeg, Manitoba, R3J 1Z1 Seoul, KOREA CANADA Phone: 204/453-7838 Deutsche Lufthansa *Japan Air Lines Weg Beim Jaeger Hanada International Airport D2000 Hamburg 63 Tokyo, JAPAN WEST GERMANY John Cameron Aviation Hanger 16 Eastern Aero Marine Bankstown Airport 3850 N.W. 25th Street N.S.W. 2200, AUSTRALIA Miami, Florida 33142 Phone: 7055812 Phone: 305/871-4050 Air Support International, Inc., Electronique Aerospatile (EAS) 6753 E. 47th Avenue Dr. Boite Postale No. 4 Denver, Colorado 80216 Le Bourget - 93 Phone: 303/333-5441 FRANCE Fire Fighter Sales Service Co., Tym's 1721 Main Street 414 West Arbor Vitae Pittsburgh, Pennsylvania 15215 Inglewood, California 90301 Phone: 412/782-2800 * Equipped For Hermetically-Sealed Containers

Bacon Equipment Co. 2709 North Beckley Avenue Dallas, Texas 75208 Phone: 214/742-5871

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*KLM - Royal Dutch Airlines Schiphol Airport P.O. Box 770 Amsterdam, THE NETHERLANDS Aero Mercantil, Leaver CIA Guaymaral Airport Bogota, COLUMBIA Air Asia Co., Ltd. Tainan Airfield Tainan, Taiwan 700 Republic of CHINA Phone: 234141 *Fas-Orient, (PTE) Ltd. Jalan Kayu, P.O. Box 17 SINGAPORE 28 Forind-Avio S.R.I. Via Bartolini, 1 20155 Milano, ITALY *Graviner, LTD Poyle Road, Colnbrook Slough, Buckinghamshire SL3 OHB ENGLAND General Fire Equipment Co. E. 4004 Trent Ave. Spokane, Washington 99202 Phone: 509/535-4255

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

D. Extinguisher Activation Check (1) Construct a test lamp consisting of one 327 light bulb connected to two 22 gage wires. The wires should be of sufficient length to allow the operator to sit in the cockpit, perform the procedure and view the test lamp connected to wiring in the wheel well. Install an alligator clip on the free end of each wire. (2) Remove electrical power from the airplane. (3) Disconnect but DO NOT REMOVE the battery. Perform the applicable Steps of the BATTERY REMOVAL procedure (Ref. Chapter 24-31-00). (4) Working through the main gear wheel wells, identify, tag and disconnect the wiring to the left and right fire extinguisher squibs. Refer to the appropriate Model 1900 Airliner Series Wiring Diagram Manual for proper wiring (Ref. Model 1900 Airliner Series Wiring Diagram Manual, P/N 114-590032-3 (UA-1 and After) or Model 1900C Airliner Wiring Diagram Manual, P/N 114-590021-13 (UB-1 and After) or Model 1900C Airliner Wiring Diagram Manual, P/N 114-590021-61 (UC-1 and After). (5) Temporarily connect the battery to the airplane electrical system. (6) Perform the APPLYING GROUND POWER procedure (Ref. Chapter 24-40-00). (7) Set the VOLT SELECT switch on the overhead panel to BATT and make note of the DC voltage displayed. (8) Set the EXT PWR switch to OFF. Set the BATT switch to OFF. NOTE: Two individuals are required to perform Steps (9) and (10). (9) Left Extinguisher Voltage Check: (a) Connect the negative lead of a DC voltmeter to wire W75A20N for UA/UB serials or wire W75B22 for UC serials. (b) Connect the positive lead of a DC voltmeter to wire W72A20 for UA/UB serials or wire W72B22 for UC serials. Set the meter to read a voltage of approximately 28 VDC. (c) Set the BATT switch to ON. Set the EXT PWR switch to the EXT PWR position. (d) Pull out the left FIRE PULL handle. (e) Press and hold the LEFT ENG FIRE PUSH TO EXT switch (Ref. Figure 201). The DC voltmeter should indicate approximately the voltage noted in Step (7). (f) Release the LEFT ENG FIRE PUSH TO EXT switch. The DC voltmeter should indicate 0.0 vdc. (g) Push in the left FIRE PULL handle. (h) Set the EXT PWR switch to OFF. Set the BATT switch to OFF. (i) Disconnect the voltmeter from the airplane wiring. (10) Right Extinguisher Voltage Check:

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (a) Connect the negative lead of a DC voltmeter to wire W85A20N for UA serials, wire W85A22N for UB serials or wire W85B22 for UC serials. (b) Connect the positive lead of a DC voltmeter to wire W82A20 for UA serials, wire W82A22 for UB serials or wire W82B22 for UC serials. Set the meter to read a voltage of approximately 28 VDC. (c) Set the BATT switch to ON. Set the EXT PWR switch to the EXT PWR position. (d) Pull out the right FIRE PULL handle. (e) Press and hold the RIGHT ENG FIRE PUSH TO EXT switch. The DC voltmeter should indicate approximately the voltage recorded in Step (7). (f) Release the RIGHT ENG FIRE PUSH TO EXT switch. The DC voltmeter should indicate 0.0 vdc. (g) Push in the right FIRE PULL handle. (h) Set the EXT PWR switch to OFF. Set the BATT switch to OFF. (i) Disconnect the voltmeter from the airplane wiring. (11) Left Fire Extinguisher Check: (a) Connect the negative test lamp lead to wire W75A20N for UA/UB serials or wire W75B22 for UC serials and the positive test lamp lead to wire W72A20 for UA/UB serials or wire W72B22 for UC serials. Position the test lamp so that it can be viewed by the operator in the cockpit. (b) Set the BATT switch to ON. Set the EXT PWR switch to the EXT PWR position. (c) Pull out the left FIRE PULL handle. (d) Press and hold the LEFT ENG FIRE PUSH TO EXT switch. The test lamp should illuminate. (e) Release the LEFT ENG FIRE PUSH TO EXT switch. The test light should extinguish. (f) Repeat the actions in Steps (d) and (e) an additional five times. (g) Press and hold the LEFT ENG FIRE PUSH TO EXT switch. Push in and pull out the left FIRE PULL handle six times. The test light should extinguish each time the left FIRE PULL handle is pushed in. The test light should illuminate each time the left FIRE PULL handle is pulled out. The left FIRE PULL handle is to end up in the out position with the test lamp illuminated. (h) Push in the left FIRE PULL handle. The test lamp should extinguish. (i) Release the LEFT ENG FIRE PUSH TO EXT switch. (j) Set the EXT PWR switch to OFF. Set the BATT switch to OFF. (k) Disconnect the test lamp from the left squib wiring.

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(12) Right Fire Extinguisher Check: (a) Connect the negative test lamp lead to wire W85A20N for UA serials or W85A22N for UB serials or wire W85B22 for UC serials and the positive test lamp lead to wire W82A20 for UA serials or wire W82A22 for UB serials or wire W82B22 for UC serials. Position the test lamp so that it can be viewed by the operator in the cockpit. (b) Set the BATT switch to ON. Set the EXT PWR switch to the EXT PWR position. (c) Pull out the right FIRE PULL handle. (d) Press and hold the RIGHT ENG FIRE PUSH TO EXT switch. The test lamp should illuminate. (e) Release the RIGHT ENG FIRE PUSH TO EXT switch. The test light should extinguish. (f) Repeat the actions in Steps (d) and (e) an additional five times. (g) Press and hold the RIGHT ENG FIRE PUSH TO EXT switch. Push in and pull out the right FIRE PULL handle six times. The test light should extinguish each time the right FIRE PULL handle is pushed in. The test light should illuminate each time the right FIRE PULL handle is pulled out. The right FIRE PULL handle is to end up in the out position with the test lamp illuminated. (h) Push in the right FIRE PULL handle. The test light should extinguish. (i) Release the RIGHT ENG FIRE PUSH TO EXT switch. (j) Set the EXT PWR switch to OFF. Set the BATT switch to OFF. (k) Disconnect the test lamp from the right squib wiring. (13) Shutdown and disconnect the external power unit. (14) Disconnect the battery from the airplane electrical system. (15) Identify and connect the wires to the left and right fire extinguisher squibs. Refer to the appropriate Model 1900 Airliner Series Wiring Diagram Manual for proper wiring (Ref. Model 1900 Airliner Series Wiring Diagram Manual, P/N 114-590032-3 (UA-1 and After) or Model 1900C Airliner Wiring Diagram Manual, P/N 114-590021-13 (UB-1 and After) or Model 1900C Airliner Wiring Diagram Manual, P/N 114-590021-61 (UC-1 and After). (16) Perform the applicable Steps of the BATTERY INSTALLATION procedure (Ref. Chapter 24-31-00). (17) Set the BATT switch to ON. NOTE: Perform Steps (18) and (19) for UA serials and UB-1 thru UB-56. Perform Steps (20) and (21) for UB-57 and After, UC serials and airplanes modified by Kit 114-3004. (18) Set the TEST SWITCH FIRE DET & FIRE EXT switch to EXT-LEFT. The D or DISCH and OK indicators should illuminate on the left engine fire extinguish switch. Set the TEST SWITCH FIRE DET & FIRE EXT switch to OFF, the D or DISCH and OK indicators go out (Ref. Figure 201).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (19) Set the TEST SWITCH FIRE DET & FIRE EXT switch to EXT-RIGHT. The D or DISCH and OK indicators should illuminate on the right engine fire extinguish switch. Set the TEST SWITCH FIRE DET & FIRE EXT switch to OFF, the D or DISCH and OK indicators go out. (20) Set the ENG FIRE TEST- EXT LEFT switch to TEST. The D or DISCH and OK indicators should illuminate on the left engine fire extinguish switch. Set the ENG FIRE TEST- EXT LEFT switch to OFF, the D or DISCH and OK indicators go out. (21) Set the ENG FIRE TEST- EXT RIGHT switch to TEST. The D or DISCH and OK indicators should illuminate on the right engine fire extinguish switch. Set the ENG FIRE TEST- EXT RIGHT switch to OFF, the D or DISCH and OK indicators go out. (22) Set the BATT switch to OFF.

2. EXTINGUISHER CARTRIDGE (SQUIB) AND SUPPLY CYLINDER A. Information (1) The storage temperature should not exceed 130°F. (2) The service temperature (installed) should be a nominal 200°F. (3) The life of a cartridge is determined as starting from the date (month/year) stamped on the cartridge body or marked on the plastic bag containing the cartridge.

3. BLEED AIR WARNING SWITCHES A. Check For Proper Electrical Connection To confirm that the left and right bleed air warning switches have not become cross connected during maintenance, perform the following check: (1) Remove the center floorboard forward of the main spar to gain access to the pressure switches. (2) Locate the left and right bleed air warning pressure switches [S128] and [S129] (Ref. Figure 201). (3) Locate the switch that connects to the plastic EVA warning line that goes into the RH wing. (4) Confirm that this switch is connected to connector P242 which should contain wire no’s. W2A22 and W4A22. WARNING: The fire extinguisher supply cylinder is a pressurized container. Refer to the appropriate overhaul manual for complete instructions on how to remove, disassemble, and empty a fire extinguisher. Failure to comply with the above, and to instructions given in the overhaul manual could cause injury to personnel and/or damage to equipment. A fire extinguisher past the due date for hydrostatic testing can not be used unless it is emptied and hydrostatically tested per the DOT specifications and the pressure gage has been shown to indicate proper charge. If the container is pressurized when removed from the airplane, it can be shipped to a recharge station, still pressurized, providing there is no evidence of visible damage, such as dents deeper than 1/16 inch per inch of average dent diameter or scratches deeper than 0.004 inch. The station must then hydrostatically test the container before recharging and returning it to the user.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Table 201 FIRE EXTINGUISHER CYLINDER PRESSURE LIMITS Temp °F Indicated Pressure in PSI

-40°

-20°

0°

+20°

+40°

+60°

+70°

+80°

+100°

+120°

127

148

174

207

249

304

334

367

442

532

to

to

to

to

to

to

to

to

to

to

155

180

212

251

299

354

385

417

492

582

Figure 201 Fire Extinguishing System

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

CHAPTER 27 - FLIGHT CONTROLS TABLE OF CONTENTS SUBJECT

PAGE

GENERAL INFORMATION 27-00-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Special Tools and Recommended Materials . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2

CONTROL COLUMN BEARING SUPPORT 27-00-01 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Repair . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 201 201

TRAVEL BOARD 27-00-02 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Travel Boards . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Reading a Travel Board . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Universal Travel Board . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Aileron Travel Board Installation (UA-1 and After; UB-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . Aileron Travel Board Installation (UC-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Aileron Trim Tab Travel Board Installation (UC-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Aileron Trim Tab Travel Board Installation (UA-1 and After; UB-1 and After; UC-1 and After (Alternate)) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Rudder Travel Board Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Rudder Travel Board Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Rudder Trim Tab Travel Board Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Elevator Travel Board Installation at HSS 50.00 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Elevator Trim Tab Travel Board Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Flap Travel Board Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Travel Board Certification . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Recommended Materials/Equipment For Certification* . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Certification Set-Up Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Aileron, Elevator and Rudder . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Aileron Trim Tab, Rudder Trim Tab and Elevator Trim Tabs . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Flap . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Travel Board Rework . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Aileron Travel Board Certification (UC-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Aileron Travel Board Certification (UA-1 and After; UB-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . Rudder Travel Board Certification . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Elevator Travel Board Certification . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Aileron Trim Tab and Rudder Trim Tab Travel Board Certification . . . . . . . . . . . . . . . . . . . . . . . . . . . Elevator Trim Tab Travel Board Certification . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Flap Travel Board Certification . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

27-CONTENTS

201 201 201 203 203 204 204 204 205 210 210 210 213 213 216 221 221 221 221 222 222 222 222 225 227 229 231 233 235

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AILERONS 27-10-00 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal (UA-1 and After; UB-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation (UA-1 and After; UB-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal (UC-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation (UC-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Freeplay Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Aileron Ground Adjustable Trim Tab . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Adjustment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 201 201 202 203 210 215 215

CONTROL WHEEL 27-10-01 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Control Column Cable . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 201 201 204 204 204

AILERON CABLES 27-10-02 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Fuselage Cable . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Wing Outboard Cable . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal (UA-1 and After; UB-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation (UA-1 and After; UB-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal (UC-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation (UC-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Wing Inboard Cable . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal (UA-1 and After; UB-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation (UA-1 and After; UB-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal (UC-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation (UC-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Outboard Wing Bellcrank . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal (UA-1 and After; UB-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation (UA-1 and After; UB-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Bellcrank Assembly . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal and Inspection (UC-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation (UC-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Yoke Assembly . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Checks (UC-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Bearing Removal (UC-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Bearing Installation (UC-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Taper Pin Removal (UC-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Taper Pin Installation (UC-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

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201 201 201 202 204 204 205 207 208 211 211 212 214 215 217 217 217 218 218 219 220 220 220 221 221 221

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

CHAPTER 27 - FLIGHT CONTROLS TABLE OF CONTENTS (CONTINUED) SUBJECT PAGE Removal (UC-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 221 Installation (UC-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 222

AILERON CONTROL SYSTEM 27-10-03 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Rigging (UA-1 and After; UB-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Rigging (UC-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Checks . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Operational Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Fuselage Cable Tension Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Control Column Interconnect Cable Tension Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Wing Cable Tension Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Functional Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 201 207 214 214 214 215 216 217

AILERON TRIM TAB 27-10-04 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal (UA-1 and After; UB-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation (UA-1 and After; UB-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal (UC-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation (UC-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Freeplay Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 201 201 203 203 205

AILERON TRIM TAB ACTUATORS AND CABLES 27-10-05 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Aileron Trim Tab Forward Cable . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Aileron Trim Tab Actuator and Cable . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal (UA-1 and After; UB-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Drum Cable Replacement (UA-1 and After; UB-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation (UA-1 and After; UB-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal (UC-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Drum Cable Replacement (UC-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation (UC-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 201 203 206 206 207 209 211 212 214

AILERON TRIM TAB INDICATOR 27-10-06 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Aileron Trim Tab Control . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Indicator Adjustment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

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201 201 201 201 201 202

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AILERON TRIM TAB CONTROL SYSTEM 27-10-07 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Aileron Trim Tab . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Rigging . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Operational Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Cable Tension Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Functional Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 201 204 204 205

AILERON BALANCE WEIGHTS 27-10-08 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Procedures (UC-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Clip Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201

RUDDER 27-20-00 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Freeplay Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 201 201 207

RUDDER CABLES 27-20-01 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Rudder Forward Control Cable . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Rudder Aft Control Cables . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 201 202 205 205 205

RUDDER CONTROL SYSTEM 27-20-02 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Rigging . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Rudder . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Operational Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Cable Tension Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Functional Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 201 206 206 206 208

RUDDER PEDALS 27-20-03 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Pilot Rudder Pedal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201

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CHAPTER 27 - FLIGHT CONTROLS TABLE OF CONTENTS (CONTINUED) SUBJECT PAGE Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Assembly Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Assembly Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202 Pilot Rudder Pedal Bellcrank . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 203 Pilot Rudder Pedal Tube . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 203 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 203 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 204 Copilot Rudder Pedal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 209 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 209 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 209 Assembly Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 209 Assembly Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 210 Copilot Rudder Pedal Bellcrank . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 210 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 210 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 211 Copilot Rudder Pedal Tube . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 211 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 211 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 211 Rudder Pedal Arm . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 215 Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 215 Aluminum Rudder Pedal Arms (Unbushed) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 215 Aluminum Rudder Pedal Arms With Bushings Installed in the Rudder Attach Holes . . . . . . . . . . . . . . 215 Magnesium Rudder Pedal Arms With Bushings Installed in the Rudder Attach Holes . . . . . . . . . . . . 216

RUDDER TRIM TAB 27-20-04 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Freeplay Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 201 201 203

RUDDER TRIM TAB CABLES AND ACTUATORS 27-20-05 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Rudder Trim Tab Forward Cable . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Rudder Trim Tab Middle Cable . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Rudder Trim Tab Actuator and Cable . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Drum Cable Replacement . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

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201 201 201 202 214 214 215 220 220 221 223

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RUDDER TRIM TAB INDICATOR 27-20-06 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Rudder Trim Tab Control . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Indicator- Adjustment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 201 201 201 202

RUDDER TRIM TAB CONTROL SYSTEM 27-20-07 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Rigging . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Operational Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Cable Tension Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Functional Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 201 205 205 206

FLIGHT CONTROLS ASSIST SYSTEMS 27-21-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Stability Augmentation System (Yaw Damper) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201

ELEVATOR 27-30-00 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Freeplay Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Inboard, Center and Outboard Hinge Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 201 202 203 205 205

ELEVATOR CABLES 27-30-01 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Elevator Forward Control Cables . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Elevator Aft Control Cables . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

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CHAPTER 27 - FLIGHT CONTROLS TABLE OF CONTENTS (CONTINUED) SUBJECT

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ELEVATOR CONTROL SYSTEM 27-30-02 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Preparation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Rigging . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Follow on Maintenance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 208 Operational Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 221 Cable Tension Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 221 Functional Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 222 Friction Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 223 Bobweight and Stop . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .229 Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .229 Bobweight Link Assembly . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .230 Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .230

ELEVATOR TRIM TABS 27-30-03 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Freeplay Checks . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Elevator Electric Trim . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Operational Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 201 201 203 208 208

ELEVATOR TRIM TAB CABLES 27-30-04 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Forward Elevator Trim Tab Cable . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Aft Elevator Trim Tab Cable . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Vertical Elevator Trim Tab Cable . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Horizontal Elevator Trim Tab Cable . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Elevator Trim Tab Control Wheel Chain . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 201 202 215 215 216 222 222 223 226 226 226 227 227 228

ELEVATOR TRIM TAB CONTROL SYSTEM 27-30-05 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Elevator Trim Tab . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Rigging . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201

27-CONTENTS

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

CHAPTER 27 - FLIGHT CONTROLS TABLE OF CONTENTS (CONTINUED) SUBJECT PAGE Operational Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 206 Cable Tension Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 207 Functional Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 207

ELEVATOR TRIM TAB ACTUATORS 27-30-06 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Cable Replacement . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 201 201 202

ELEVATOR ELECTRIC TRIM TAB SYSTEM 27-30-07 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Elevator Electric Trim Tab . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Cable Disconnection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Actuator Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Actuator Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Cable Connection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Left Cable Drum and Cable . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Actuator Cable Drum and Cable . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Servo Universal Joint . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Magnetic Clutch . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Torque Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Elevator Electric Trim Tab . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Cable Rigging (All Airplanes Without Collins APS-65H Autopilot System) . . . . . . . . . . . . . . . . . . . . . Actuator Speed Adjustment (All Airplanes Without Collins APS-65H Autopilot System) . . . . . . . . . . .

201 201 201 201 202 202 203 203 203 204 204 204 205 205 205 210 210 210 210 211 211 211

ELEVATOR TRIM TAB INDICATOR 27-30-08 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Adjustment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 201 201 201 202

STALL WARNING SYSTEM 27-31-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1

Page 8 Nov 1/13

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

CHAPTER 27 - FLIGHT CONTROLS TABLE OF CONTENTS (CONTINUED) SUBJECT PAGE Lift Transducer . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Lift Computer . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2 Troubleshooting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 101 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 101 Stall Warning System Calibration . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .101 Ground Calibration Section . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 101 Air Calibration Section . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 102

FLAPS 27-50-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Outboard Flap . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202 Flap Roller Bracket Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 203 Inboard Flap . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 208 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 208 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 209

FLAP CABLES 27-50-01 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Flap Drive Cables . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201

FLAP TRACKS 27-50-02 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Wear Limits . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Repair . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 201 201

FLAP MOTOR AND GEARBOX 27-50-03 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Flap Motor Gearbox Freedom of Movement Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Flap Motor . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Flap Flexible Drive Shaft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Outer Housing Inspection and Repair . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Inner Shaft Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

27-CONTENTS

201 201 201 201 202 202 202 203 204 204 205 206 207

Page 9 Nov 1/13

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

CHAPTER 27 - FLIGHT CONTROLS TABLE OF CONTENTS (CONTINUED) SUBJECT

PAGE

FLAP ACTUATORS 27-50-04 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Outboard . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Inboard . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 201 201 202 202 203

FLAP CONTROL SYSTEM 27-50-05 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Rigging - Using Travel Boards . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Rigging - Using Protractors . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Rigging - Up Position Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Flap Adjustment to Correct a Wing Heavy Condition . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Functional Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Operational Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 201 206 212 214 215 216

FLAP SAFETY SYSTEM 27-50-06 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Flap Safety Switch (Asymmetric) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal (With Kit No. 114-5057 Installed) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation (With Kit No. 114-5057 Installed) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Test (UA-1 and After; UB-1 and After; UC-1 thru UC-53) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Test (UC-54 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Adjustment (Without Kit No. 114-5057 Installed) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Adjustment (With Kit No. 114-5057 Installed) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Hub Lubrication (Without Kit No. 129-5046 Installed) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 201 201 202 202 203 204 205 206 208

FLAP POSITION SWITCHES 27-50-07 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Check Out and Rigging . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Rigging Requirements . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 201 201 201 202

GUST LOCKS AND DAMPENERS 27-70-00 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Control Locks . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201

Page 10 Nov 1/13

27-CONTENTS

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

CHAPTER 27 - FLIGHT CONTROLS TABLE OF CONTENTS (CONTINUED) SUBJECT PAGE Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202

27-CONTENTS

Page 11 Nov 1/13

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

List of Effective Pages CH-SE-SU

PAGE

DATE

CH-SE-SU

PAGE

DATE

27-LOEP

1

Nov 1/15

27-30-08

201 thru 205

Nov 1/09

27-CONTENTS

1 thru 11

Nov 1/13

27-31-00

27-00-00

1 thru 8

Nov 1/15

1 and 2 101 thru 107

Nov 1/09 May 1/11

27-00-01

201 thru 205

Nov 1/09

27-50-00

27-00-02

201 thru 237

Nov 1/09

1 thru 3 201 thru 210

Nov 1/09 Nov 1/09

27-10-00

201 thru 221

Aug 1/12

27-50-01

201

Nov 1/09

27-10-01

201 thru 204

Nov 1/09

27-50-02

201

Nov 1/09

27-10-02

201 thru 236

Nov 1/09

27-50-03

201 thru 212

Nov 1/09

27-10-03

201 thru 230

Nov 1/09

27-50-04

201 thru 207

Nov 1/09

27-10-04

201 thru 210

Nov 1/09

27-50-05

201 thru 224

Nov 1/09

27-10-05

201 thru 234

Nov 1/09

27-50-06

201 thru 213

Nov 1/09

27-10-06

201 thru 204

Nov 1/09

27-50-07

201 thru 203

Nov 1/09

27-10-07

201 thru 210

Nov 1/09

27-70-00

201 and 202

Nov 1/09

27-10-08

201 thru 208

Feb 1/10

27-20-00

201 thru 212

May 1/12

27-20-01

201 thru 210

Nov 1/09

27-20-02

201 thru 217

Nov 1/09

27-20-03

201 thru 217

Nov 1/13

27-20-04

201 thru 206

Nov 1/09

27-20-05

201 thru 230

Nov 1/09

27-20-06

201 thru 204

Nov 1/09

27-20-07

201 thru 212

Nov 1/09

27-21-00

1 and 2 201

Nov 1/09 Nov 1/09

27-30-00

201 thru 212

Nov 1/09

27-30-01

201 thru 210

Nov 1/09

27-30-02

201 thru 234

Nov 1/13

27-30-03

201 thru 208

Nov 1/09

27-30-04

201 thru 230

Nov 1/09

27-30-05

201 thru 217

Nov 1/09

27-30-06

201 thru 207

Nov 1/15

27-30-07

201 thru 212

Nov 1/09

27-LOEP

Page 1 Nov 1/15

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

FLIGHT CONTROLS GENERAL INFORMATION DESCRIPTION AND OPERATION

27-00-00 00

1. GENERAL All flight controls, with the exception of the flaps, are cable-operated conventional surfaces which require no power assistance for normal control by the pilot or copilot. The flaps and optional elevator trim are electrically powered. The yaw dampening system is to aid the pilot with directional control. All primary flight control surfaces are manually controlled through cable-pulley-bellcrank systems. Dual controls are provided for operation by either the pilot or the copilot. The ailerons and elevators are operated by conventional control wheels interconnected by a “T” shaped control column. Rudder pedals are interconnected by a linkage below the flight compartment floor. Rudder bellcranks are adjustable to two positions which move the rudder pedals approximately one inch forward or aft. Surface travel stops and linkage adjustments are incorporated into each cable-pulley-bellcrank system. Ailerons, elevators and rudder may be secured with control locks installed in the flight compartment when the airplane is on the ground and out of service. Two flaps installed on each wing are operated by an electric motor-driven gearbox mounted on the forward side of the rear spar at the centerline of the airplane. The gearbox drives four flexible drive shafts, each connected to an Acme-thread-type jackscrew at each flap. A flap limit safety switch is provided to disconnect power to the electric motor in the event of any type of failure which causes any flap to be 3° to 6° out of phase with the adjacent flap. The flaps are controlled by a lever mounted in the pedestal. Flap lever detents are provided to select UP, TAKEOFF, APPROACH and LANDING flap positions. Wing flap position is shown by an indicator located in the pedestal near the flap control lever. The indicator is controlled by a flap position potentiometer which is actuated by the right inboard flap. Trim tabs are installed on the left aileron, the rudder and on each elevator. The tabs are manually controlled by the pilot through drum-cable systems using jackscrew actuators. The tabs are driven by the actuators through an adjustable double clevis rod assembly capable of removing joint free play. Tab position indicators are provided on the pedestal tab controls. The optional electric-motor-driven elevator trim tabs are activated by a control switch on the outboard handle of each control wheel. A downspring and bob weight are incorporated into the elevator control system for improved stability. Positive stops on the primary flight control surfaces limit their travel, while traveling stops secured to the cables limit trim tab movement. Proper safetying of the trim tab cable stops prevents their loosening and moving on the cables. Because the cables are connected together with turnbuckles, each cable has a leftand right-threaded cable end. Proper winding of the cables on the pedestal and actuator drums ensures against crossing the cables and causing improper trim tab movement. Refer to the applicable rigging procedures for details regarding chain and cable tension, control wheel movement and force, down spring force and systems friction.

27-00-00

Page 1 Nov 1/15

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

A. Special Tools and Recommended Materials The special tools listed in Table 1 (shown in Figures 1, 2, 3, 4 and 5), and recommended materials listed in Table 2 as meeting federal, military or supplier specifications are provided for reference only and are not specifically prescribed by Hawker Beechcraft Corporation. Any product conforming to the specification listed may be used. The products included in these tables have been tested and approved for aviation usage by Hawker Beechcraft Corporation by the supplier, or by compliance with the applicable specifications. Generic or locally manufactured products which conform to the requirements of the specification may be used even though not included in the tables. Only the basic number of each specification is listed. No attempt has been made to update the listing to the latest revision. It is the responsibility of the technician or mechanic to determine the current revision of the applicable specification prior to usage of the product listed. This can be done by contacting the supplier of the product to be used. Table 1 SPECIAL TOOLS AND EQUIPMENT TOOL NAME 1. Trim Tab Freeplay Check Fixture 2. Dial Indicator

PART NUMBER 45-135030-9/810 or local manufacture (Ref. Figure 1). C8IQ

3. Back Plate, Dial Indicator

BK-692

4. Cable Tensiometer 5. Digital Protractor Model 3600

KS6005 or equivalent

6. Push-pull Scale 7. Flight Control Rig Pin (Elevator, Aileron, Aft Rudder) 8. Go/No-Go Scale 9. Flight Control Rig Pin (Fwd Rudder Bellcrank) 10. Flight Control Rig Pin (Rudder Pedals) 11. Flight Control Rig Pin (Aileron Wing Bellcrank, UA and UB Models Only)

Page 2 Nov 1/15

27-00-00

SUPPLIER Hawker Beechcraft Corporation, 9709 E. Central, Wichita, KS 67201 Federal Products, Providence, RI Federal Products Providence, RI Obtain locally Kell-Strom, 214 Church St., Wethersfield, CT 06109 800-851-6851 Obtain locally P/N 14-380069-0003 or fabricate locally (Ref. Figure 2). Fabricate locally (Ref. Figure 3). P/N 114-380069-0005 or fabricate locally. (Ref. Figure 4). P/N 114-380069-0007 or fabricate locally (Ref. Figure 4). P/N 114-380069-0011 or fabricate locally (Ref. Figure 5).

USE Checking flight control trim tab freeplay. Tab freeplay check measurements. Back Screw for mounting the Dial Indicator to Trim Tab Freeplay Check Fixture. Check cable tension. Flight Control System rigging.

Checking control force. Rigging flight controls. Measure forward elevator bellcrank stop bolt clearance. Rigging flight controls. Rigging flight controls. Rigging flight controls.

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Table 1 SPECIAL TOOLS AND EQUIPMENT (Continued) TOOL NAME

PART NUMBER

SUPPLIER

USE

12. Flight Control Rig Pin Set (UC Models Only)

P/N 114-380069-0001

13. Flight Control Rig Pin Set (UA and UB Models Only)

P/N 114-380069-0009

Rigging flight controls. Set contains: 2 ea. P/N 114-380069-0003 1 ea. P/N 114-380069-0005 1 ea. P/N 114-380069-0007 Rigging flight controls. Set contains: 2 ea. P/N 114-380069-0003 1 ea. P/N 114-380069-0005 1 ea. P/N 114-380069-0007 1 ea. P/N 114-380069-0011

Table 2 RECOMMENDED MATERIALS MATERIALS 1. Grease

2. Cleaning Solvent

SPECIFICATION MIL-G-23827

PRODUCT

SUPPLIER

Supermil Grease A72823

American Oil Co., 165 N. Canal Chicago, IL 60606

Aeroshell Grease 7

Shell Oil Co., Shell Plaza, P.O. Box 2463 Houston, TX 77001

PD680 Type III

3. Sealer

Obtain locally EC1239

Minnesota Mining Mfg. Co., St. Paul, MN

Petrotect Grade 2

Pennsylvania Refining Co., 1686 Lisbon Rd. Cleveland, OH 44104

Loctite 680

Loctite Corp., 705 W. Mountain Rd., Newington, CT 06111

Brayco 300

Castrol Industrial North America, 5331 E. Slauson Commerce CA 90040

7. Lubricant

Lubriplate 130A or Lubriplate Aero

Fiske Bros. Refining Co., 129 Lockwood Newark, NJ 07105-4820

8. Solvent

Methyl Propyl Ketone

Obtain locally

9. Filler

Devcon WR-11220

Devcon Corp., 59 Endicott St. Danvers, MA 01923

4. Corrosion Preventive Compound

MIL-C-16173

5. Retaining Compound

6. Lubricant, Preservative

VV-L-800

27-00-00

Page 3 Nov 1/15

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Table 2 RECOMMENDED MATERIALS (Continued) MATERIALS

SPECIFICATION

10. Filler

PRODUCT

SUPPLIER

Devcon WR-11410

Devcon Corp., 59 Endicott St. Danvers, MA 01923

11. AntiCorrosion Treatment

MIL-C-5541

Alodine 1200, 1200S or 1201

Amchem Products Inc., Ambler, PA 19002

12. Sealer

MIL-S-8802

Pro-Seal 890B 1/2

Coast Paint and Chemical Co., 1507 Grande Vista Ave. Los Angeles, CA

838 Tape

Minnesota Mining and Mfg. Co., St. Paul, MN

13. Tedlar Tape 14. Naptha

TT-N-95 Type II

15. Sealant

Obtain locally EC-2084

3M Engineered Adhesive Div. 3M Center St. Paul, MN 55144

Epoxy-Polyamide

U.S. Paint Lacquer and Chemical Co., St. Louis, MO

17. Grease

Molykote #33 Light Extreme Low Temperature or Dow Corning #33 Light Extreme Low Temperature

Dow Corning S. Saginaw Rd. Midland, MI 48641

18. Adhesive

EC-2216

3M Engineered Adhesive Div. 3M Center St. Paul, MN 55144

19. Sealant Tape

Av-dec Hi-Tak Tape P/N HT3000XXX or Dow Corning Gel Tec Sealing Strips P/N GT-1000-1-R25 or equivalent.

Obtain locally

16. Primer

Page 4 Nov 1/15

MIL-P-23377

27-00-00

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

ALL DIMENSIONS ARE IN INCHES

6,7,8 5

2.9

13

A

.5 .25

9 10 11 12

.80

C

1.25 TYP

B B

C

A TOP VIEW 125

4

15

TAB

TAP 5/8-11 IF USING ITEM 13 OTHERWISE TAP 1/2-13

3

1

2

.250

14 1.3

.500 VIEW

A-A

.7

BK-692 BACK PLATE MACHINE DETAIL VIEW ITEM NO

QUANT

C-C

.68

DESCRIPTION

.31 1.6 1 2 3/4 X 1 X 6 ALUMINUM OR EQUIV. 2 2 1 X 1 3/8 X 1 3/4 ALUMINUM OR EQUIV. 3 1 1/2 X 7 1/2 X 10 ALUMINUM OR EQUIV. 4 1 C81Q OR 281QN INDICATOR** 5 1 3/4 X 2 1/2 X 14 ALUMINUM OR EQUIV 6 1 1/4 DIA. X 2 CORROSION RES. STL. 7 1 1/4 DIA. X 1 CORROSION RES. STL. * 8 1 1/4-28 NUT 9 1 3/8 X 5 X 10 RUBBER 10 1 3/8 X 2 X 10 RUBBER VIEW 11 1 1/4 X 2 X 10 CORROSION RES. STL. * THIS GROOVE TO BE A SNUG FIT TO THE 12 2 TS 107 1/2-13 X 3 VLIER TORQUE SCREW SCREW BRACKET ON THE DIAL INDICATOR 13 2 KN813 KEENSERT OR TAP 1/2-13 14 2 1/8 X 1 X 3/4 RUBBER 15 1 BK692 BACK PLATE** **AVAILABLE THROUGH YOUR LOCAL BEECH PARTS OUTLET. TRIM TAB FREEPLAY CHECK FIXTURE

B-B

UC27B 042320AC.A

Figure 1 Trim Tab Freeplay Fixture

27-00-00

Page 5 Nov 1/15

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

HANDLE

10 INCHES

MAKE FROM STEEL ROD OR DRILL ROD

FLIGHT CONTROL RIG PIN (ELEVATOR, AILERON, AFT RUDDER)

0.184 - 0.189 INCH DIAMETER UC27B 041846AC.AI

Figure 2 Flight Control Rig Pin

TAPE

TAPE

1

2

3

4

5

6

7

8

9

1

2

3

4

5

6

7

8

9

1

SCALE

1

2

3

4

5

6

7

8

9

1

2

3

4

5

6

7

8

9

2

1

2

3

4

5

6

7

8

9

1

2

3

4

5

6

7

8

9

3

1

2

3

4

5

6

7

8

9

1

2

3

4

5

6

7

8

9

4

1

2

3

1

2

3

.XX .XX

NOTE: THIS SCALE IS AN EXAMPLE ONLY. DIMENSIONS MAY VARY. UC27B 041845AB.AI

Figure 3 Go/No-Go Scale

Page 6 Nov 1/15

27-00-00

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

HANDLE

MAKE FROM STEEL ROD OR DRILL ROD

10 INCHES

0.315 - 0.321 INCH DIAMETER FLIGHT CONTROL RIG PIN (FWD RUDDER BELLCRANK)

0.375 - 0.385 INCH DIAMETER 13 INCHES

MAKE FROM STEEL ROD OR DRILL ROD 2.75 - 3.00 INCHES HANDLE

FLIGHT CONTROL RIG PIN (RUDDER PEDALS)

UC27B 042386AB.AI

Figure 4 Rudder Rig Pins

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Page 7 Nov 1/15

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

HANDLE

MAKE FROM STEEL ROD OR DRILL ROD

7.0-7.5 INCHES

0.246-0.249 INCH DIAMETER FLIGHT CONTROL RIG PIN (AILERON WING BELLCRANK UA-1 AND AFTER; UB-1 AND AFTER)

Figure 5 Aileron Wing Bellcrank Rig Pin (UA-1 and After; UB-1 and After) Page 8 Nov 1/15

27-00-00

UA27B 044814AA.AI

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

FLIGHT CONTROLS CONTROL COLUMN BEARING SUPPORT MAINTENANCE PRACTICES

27-00-01 200200

1. PROCEDURES A. Inspection (1) Remove the access panel (3) located aft of the nose wheel well (Ref. Chapter 6-50-00, FUSELAGE ACCESS PANELS). NOTE: If discrepancies are noted during this inspection, perform the CONTROL COLUMN BEARING SUPPORT REPAIR procedure, in this section. CAUTION: If the gap between the support arm (5) and the bearing support (6) exceeds 0.020-inch and/or loose, working or missing rivets are found, inspect adjacent attachment intercostal/keel (7) area for deformation and cracks (Ref. Figure 201). (2) Inspect the left and right control column bearing supports (6) for loose, working or missing rivets. (3) Check the gap between left and right support arm (5) and bearing support (6) to verify the gap does not exceed 0.020-inch. (4) If applicable, install the access panel.

B. Repair This procedure should only be performed to correct discrepancies found in use or during a detailed periodic inspection (Ref. Chapter 5-20-04, Third 200-Hour Interval Detailed Inspection). WARNING: Whenever any part of this system is dismantled, adjusted, repaired or replaced, a detailed investigation must be made on completion to make sure that distortion, tools, rags, or any other loose articles, or foreign matter that could impede the free movement and safe operation of the system are not present, and the systems and installation in the work area are clean. (1) Remove all power from the airplane and disconnect the battery. Display warning notices prohibiting reconnection of airplane electrical power. (2) Remove the forward cabin partitions and any associated cabinetry (as required), if installed. (3) Remove the flight compartment seats (Ref. Chapter 25-10-00, SEAT REMOVAL). (4) Remove the extrusion on the right side of the pedestal. (5) Remove the control column boot and plates forward of the pedestal and on each side of the pedestal. (6) Remove the forward flight compartment floorboard panels and carpet. (7) Disconnect the link from the bob weight. (8) Remove the access panel (3) located aft of the nose wheel well (Ref. Chapter 6-50-00, FUSELAGE ACCESS PANELS).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (9) Working through the access hole, disconnect the push-pull tube and if applicable, the pitch change transducer rod end. NOTE: Install suitable cable blocks forward of floor member at approximately FS 107 prior to disconnecting the aileron control column cables. (10) Disconnect the aileron control column cables. (11) Remove the left and right taper pins (12) from the aft side of the universal joint (11) and disconnect the torque tubes (1) (Ref. Figure 201). NOTE: The radio panel, in the center of the instrument panel, may need to be removed for clearance (Ref. Chapter 39-10-00, INSTRUMENT AND CONTROL PANELS). (12) Remove the control column mounting bolts (8). Raise the control column (10) and support in place using a dowel rod inserted into the lower tube (support arm crossmember just forward of the bearing supports). (13) Carefully remove rivets and remove the bearing support (6). Refer to the Model 1900 Airliner Series Structural Repair Manual. (14) Inspect the bearing supports (6) and the support arm (5) for damage. Inspect for damage in the adjacent area and ensure that the proper edge margin for the bearing support and the support arm is maintained. (15) If necessary, drill the support arm and bearing support attach holes to 0.159 to 0.164-inch diameter. NOTE: It is recommended that the bearing support attach rivets be installed with the manufactured (resultant) head of the rivet on the inboard side of the support arm. (16) Attach each bearing support (6) to the support arm (5) on the control column with 6 CR3213-5 Cherry Max rivets or MS20470AD5 solid rivets (Ref. Model 1900 Airliner Series Structural Repair Manual, Chapter 51-40-00). (17) Ensure the gap between the bearing support (6) and the support arm (5) does not exceed 0.020-inch. NOTE: The maximum unshimmed total gap between the bearing supports (6) and the adjacent attachment intercostals/keels (7) measured prior to bolt installation is not to exceed 0.156-inch. Shim as required with washers (NAS1149F0516P, NAS1149F0532P, NAS1149F0563P, AN960-516, or AN960-516L) between the bearing supports and the adjacent attachment intercostals/keels to reduce the total gap to a maximum of 0.030-inch (0.015-inch maximum per side). Shim equally on each side of the control column. It is acceptable to substitute bolts up to two grip lengths longer as required to allow washers to be added while maintaining at least 1 1/2 thread protrusion through the nut. It is recommended that control column mounting bolts be installed with the bolt head on the outside of the attachment intercostal/keel and the nut on the inside of the support arm. (18) Carefully lower the control column into place and insert mount bolts into the left and right attachment intercostal/keel and the adjacent bearing support.

Page 202 Nov 1/09

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (19) Position the control column to either the left, or right side until the bearing support touches the adjacent attachment (intercostal/keel). On the same side, install the washer (9) and nut (3) on the mounting bolt (8) and torque to 50 to 65 inch-pounds. (20) On the opposite side measure the resultant gap between the bearing and adjacent attachment intercostal/keel. NOTE: A total gap of 0.030-inch maximum (0.015-inch per side) between the bearing supports (6) and adjacent attachment intercostal/keel (7) is acceptable. It is acceptable to reduce gaps of less than 0.030-inch if desired, provided the total interference fit does not exceed 0.020-inch (0.010-inch per side). A total gap of between 0.030 and 0.156-inch must be reduced by installing washers as shims. Use either NAS1149F0516P, NAS1149F0532P, NAS1149F051663P, AN960-516, or AN960-516L washers as shims. The washers on each side of the column must be of equal thickness, and must reduce the gap of each side to 0.015-inch or less. No more than two washers of any thickness may be installed per side. The total interference fit not to exceed 0.020-inch (0.10 per side) is also acceptable. If the gap exceeds 0.156-inch call Hawker Beechcraft Corporation for assessment. (21) Shim equally as required to achieve a maximum gap of 0.030-inch. Ensure bolts are of suitable grip length to obtain a proper fit. Install bolt(s) (8), washer(s) (9) and nut(s) (3) and torque to 50 to 65 inch-pounds. (22) Install cotter pins (4). CAUTION: Excessive driving of the taper pin can cause cracks in the tubing. A lightweight rawhide or nylon mallet should be used to set the taper pin when it is installed. (23) Install taper pins (12), taper pin washers (2) and nuts (3) to secure the torque tubes (1) to the universal joints (11). Ensure that the small end of the taper pin is at least flush with, but not more than 0.06-inch above the surface of the universal joint and torque to 15 to 20 inch pounds. (24) Install cotter pins (4). (25) Connect the aileron control column cables and remove cable blocks forward of floor member at approximately FS 107. (26) Connect the push-pull tube and if applicable, the pitch change transducer rod end. (27) Connect the link to the bob weight. (28) Rig the cables (Ref. 27-10-03, Control System Rigging). Rig any other systems disturbed during accomplishment of this repair in accordance with the applicable chapter of this maintenance manual. (29) Install the boot and plates forward of the pedestal and on each side of the pedestal. (30) Install the extrusion on the right side of the pedestal. (31) Install the forward cabin partitions and any associated cabinetry. (32) Install the flight compartment seats (Ref. Chapter 25-10-00, SEAT INSTALLATION). (33) Install the forward flight compartment floorboard panels and carpet.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (34) Install the access panel (3) located aft of the nose wheel well (Ref. Chapter 6-50-00, FUSELAGE ACCESS PANELS). (35) If applicable, install the radio panel in the center of the instrument panel (Ref. Chapter 39-10-00, INSTRUMENT AND CONTROL PANELS). (36) Connect the battery and remove warning notices prohibiting reconnection of airplane electrical power.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 201 Control Column Bearing Support Repair

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

FLIGHT CONTROLS TRAVEL BOARD MAINTENANCE PRACTICES

27-00-02 200200

1. PROCEDURES A. Travel Boards Travel boards are used to aid in rigging of the primary and secondary control surfaces. A list of Hawker Beechcraft Corporation approved travel boards are found in Table 201. Travel boards for the primary control surfaces are constructed of aluminum with phenolic or nylon locating stops to set the travel board on the control surface. Once these stops are located on the surface, zero position and surface limits can be read on the respective boards. Travel boards for the secondary surfaces (trim tabs) are made from stainless steel. These “wedge” style travel boards are designed to nest on the primary control surface adjacent to the trim tab. The trim tab zero position and the trim tab surface limits can be read on the respective boards. Travel boards provided as ship sets may have Hawker Beechcraft Corporation internal model designation assigned. Model 1900 is equivalent to UA serial aircraft. Model 1900C is equivalent to UB serial aircraft. Model 1900C-1 is equivalent to UC serial aircraft. Table 201 SPECIAL TOOLS AND EQUIPMENT TOOL NAME

PART NO.

REVISION

SUPPLIER

USE

1. Aileron Travel Board* Alternate: See Item 9.

99-524000/810 for UA-1 and After, UB-1 and After at WS 236.923

D or Later

Hawker Beechcraft Corporation, 9709 E. Central, Wichita, KS 67201

Measure aileron control system.

2. Aileron Travel Board* Alternate: See Item 9.

118-130000-3/810-1 for UC-1 and After at WS 276.010

D or Later

Hawker Beechcraft Corporation, 9709 E. Central, Wichita, KS 67201

Measure aileron control system.

3. Aileron Trim Tab Travel Board*

118-130000-1/810-2 at Aileron Station 41.660

None or Later

Hawker Beechcraft Corporation, 9709 E. Central, Wichita, KS 67201

Measure aileron trim tab control system.

4. Rudder Travel Board*

114-630000-1/810

C or Later

Hawker Beechcraft Corporation, 9709 E. Central, Wichita, KS 67201

Measure rudder control system

5. Rudder Trim Tab Travel Board*

114-630000-1/810-2

None or Later

Hawker Beechcraft Corporation, 9709 E. Central, Wichita, KS 67201

Measure rudder trim tab control system

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Table 201 SPECIAL TOOLS AND EQUIPMENT (Continued) TOOL NAME

PART NO.

REVISION

SUPPLIER

USE

6. Elevator Travel Board* Alternate: See Item 9.

101-610000-1/807-2 at HSS 50.00

C or Later

Hawker Beechcraft Corporation, 9709 E. Central, Wichita, KS 67201

Measure elevator control systems.

7. Elevator Trim Tab Travel Board*

101-610000-1/810-2, 101-610000-2/810-2

C or Later

Hawker Beechcraft Corporation, 9709 E. Central, Wichita, KS 67201

Measure elevator trim tab control systems.

8. Flap Travel Board*

118-521046-1/810

E1 or Later

Hawker Beechcraft Corporation, 9709 E. Central, Wichita, KS 67201

Measure flap control system

9. Universal Travel Board

TE-100TB

2/11/03 or Later

Hawker Beechcraft Corporation, 9709 E. Central, Wichita, KS 67201

Measure control systems

* The following list of Hawker Beechcraft Corporation travel boards have been identified/superseded to those in Table 7, Chapter 91-00-00, but are approved: 810 99-524000-1 (Aileron, WS 236.9, UA, UB), 810-1 118-130000-1 (Aileron, WS 276, UC), 810 99-524000 (Aileron Trim Tab, WS 236.9, UA, UB), D810-1 101-130001-1 (Aileron Trim Tab, WS 276, UC), D807 101-610000-1, -2 (Elevator, HSS 35), D807-2 101-610000-1 (Elevator, HSS 50), 807 101-610000-1, -2 (Elevator Trim Tab), D807 114-630000-1 (Rudder), 807 114-630000-1 (Rudder Trim Tab), 810 114-521046-1 (Flap).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

B. Reading a Travel Board To obtain accurate travel board readings, place a straight edge against the travel board parallel to the degree markings and align the top corner of the straight edge with the center of the control surface trailing edge (Ref. Figure 219).

C. Universal Travel Board The Universal Travel Board (9, Table 201) is an adjustable tool designed to find the zero position of various control surfaces. Refer to Table 202 and the Universal Travel Board manual P/N 98-32928E or subsequent for proper use. NOTE: Calibration required. Refer to the Universal Travel Board Manual P/N 98-32928E or subsequent, for information on using and calibrating the travel board with Required settings shown below. The Universal Travel Board must be calibrated annually. Table 202 Universal Travel Board Placement SURFACE

LOCATION

AILERON

WS 276.010 (Ref. Note 1)

1st rib outboard from trim tab cutout, ELEVATOR HSS 50.00 (Ref. Note 2)

SPARS

SETTINGS (INCHES) A1

A2

COMMENTS

A3

N

13 3/4 ± 1/32

14 ± 1/32

4 27/32 ± 1/32

N

10 1/2 ± 1/32

8 1/4 ± 1/32

5 1/2 ± 1/32

Digital Protractor can be used to level both control wheels.

NOTE 1: Locate A2 leg 1.25 inches forward of the upper aft wing skin edge at the aileron cove at WS 276.010. The horizontal beam should be perpendicular to the forward spar. WS 276.010 is located 13.1 inches outboard of the centerline of the aileron central hinge at the rear spar. NOTE 2: Locate A2 leg 1.25 inches forward of the upper aft horizontal stabilizer skin edge at the elevator cove at HSS 50.00. The horizontal beam should be in line with the rivets at HSS 50.00.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

D. Aileron Travel Board Installation (UA-1 and After; UB-1 and After) The Aileron Travel Board (1, Table 201) is mounted on top of the wing and is used to measure the travel of the aileron through its full range (Ref. Figure 201). This procedure defines the installation of the Aileron Travel Board at WS 236.923. (1) Determine WS 236.923 on the trailing edge by measuring 13.60 inches from the center line of the inboard hinge (5) outboard along the rear spar (4) and mark the location. (2) Determine WS 236.923 on the forward spar (2) by measuring 80.52 inches from the inboard edge of the light dam (1) on the center of the forward spar (2) inboard and mark the location. NOTE: One travel board may be used and moved from one side to the other. (3) Place the Aileron Travel Board (6) at WS 236.923 (± 0.25 inch) on the top of the wing centered on the locations identified in Steps (1) and (2). (4) Ensure all stops (3) contact the wing surface.

E. Aileron Travel Board Installation (UC-1 and After) The Aileron Travel Board (2, Table 201) is mounted on top of the wing and is used to measure the travel of the aileron through its full range (Ref. Figure 202). This procedure defines the installation of the Aileron Travel Board at WS 276.010. (1) Determine WS 276.010 on the trailing edge by measuring 13.1 inches outboard from the center line of the middle hinge (5) along the rear spar (4) and mark the location. (2) Determine WS 276.010 on the forward spar (2) by measuring 42.9 inches inboard from the inboard edge of the light dam (1) on the center of the forward spar (2) and mark the location. NOTE: One travel board may be used and moved from one side to the other. (3) Place the Aileron Travel Board (6) at WS 276.010 (± 0.25 inch) on the top of the wing centered on the locations identified in Steps (1) and (2). (4) Ensure all stops (3) contact the wing surface.

F. Aileron Trim Tab Travel Board Installation (UC-1 and After) The Aileron Trim Tab Travel Board (3, Table 201) is mounted on the aileron. The Aileron Trim Tab Travel Board is used to measure the travel of the aileron trim tab through its full range (Ref. Figure 203). Due to variances in the aileron surface, it may be difficult to align zero on the Aileron Trim Tab Travel Board to the aileron trailing edge. It is permissible to install the Aileron Trim Tab Travel Board onto the aileron trim tab facing outboard (readings would be upside down). The travel board would move with the aileron trim tab and readings would come from the aileron trailing edge. Perform the AILERON TRIM TAB TRAVEL BOARD INSTALLATION (UA-1 AND AFTER; UB-1 AND AFTER; UC-1 AND AFTER (ALTERNATE)) procedure in this section. (1) Position the Aileron Trim Tab Travel Board (4) on the aileron (1) within 0.25 inch of the aileron trim tab cut out (2) on the aileron (1).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (2) Ensure the Aileron Trim Tab Travel Board (4) fits tightly on the aileron (1) and the trailing edge of the aileron trim tab (3) is within 0.25 inch of the stop (5) on the aileron trim tab travel board (4). (3) Verify that the 0° reading on the Aileron Trim Tab Travel Board (4) is aligned with the trailing edge of the aileron (1).

G. Aileron Trim Tab Travel Board Installation (UA-1 and After; UB-1 and After; UC-1 and After (Alternate)) The Aileron Trim Tab Travel Board (3, Table 201) is mounted on the aileron trim tab facing outboard (readings will be upside down). The travel board will move with the aileron trim tab and readings will come from the aileron trailing edge. The Aileron Trim Tab Travel Board is used to measure the travel of the aileron trim tab through its full range (Ref. Figure 204). (1) Position the Aileron Trim Tab Travel Board (4) on the aileron trim tab (3) within 0.25 inch of the aileron trim tab outboard edge (2). (2) Ensure the aileron trim tab travel board (4) fits tightly on the aileron trim tab (3) and the trailing edge of the aileron trim tab (3) is within 0.25 inch of the stop (5) on the Aileron Trim Tab Travel Board (4). (3) Verify that the 0° reading on the Aileron Trim Tab Travel Board (4) is aligned with the trailing edge of the aileron trim tab (3).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

1. INBOARD EDGE OF LIGHT DAM 2. FORWARD SPAR 3. STOP 4. REAR SPAR 5. CENTER LINE OF INBOARD HINGE 6. AILERON TRAVEL BOARD

3 2

3

1

80.52 INCHES

4

13.60 INCHES 5

6 UC27B 043119AA.AI

Figure 201 Aileron Travel Board Installation WS 236.923 (UA-1 and After; UB-1 and After) Page 206 Nov 1/09

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

1. INBOARD EDGE OF LIGHT DAM 2. FORWARD SPAR 3. STOP 4. REAR SPAR 5. CENTER LINE OF MIDDLE HINGE 6. AILERON TRAVEL BOARD

3 2

3

1

42.9 INCHES

4

13.1 INCHES 5

6 UC27B 042560AA.AI

Figure 202 Aileron Travel Board Installation WS 276.010 (UC-1 and After)

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

FWD WITHIN 0.25 INCH

1

A

OUTBD

1 2 3

B

3 VIEW LOOKING OUTBOARD ALONG TRAILING EDGE OF AILERON

4 5

VIEW LOOKING DOWN DETAIL

A FWD

4

5 1

1. AILERON 2. AILERON TAB CUT OUT 3. AILERON TRIM TAB 4. AILERON TRIM TAB TRAVEL BOARD 5. STOP

WITHIN 0.25 INCH VIEW LOOKING INBOARD DETAIL

B

Figure 203 Aileron Trim Tab Travel Board Installation (UC-1 and After)

Page 208 Nov 1/09

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UC27B 042566AB.AI

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

FWD WITHIN 0.25 INCH

A

1

2 OUTBD

3

1

B VIEW LOOKING INBOARD ALONG TRAILING EDGE OF AILERON

5

4

VIEW LOOKING DOWN DETAIL

A FWD 3 4

5

1. AILERON 2. AILERON TRIM TAB OUTBOARD EDGE 3. AILERON TRIM TAB 4. AILERON TRIM TAB TRAVEL BOARD 5. STOP

WITHIN 0.25 INCH VIEW LOOKING OUTBOARD DETAIL

B

UC27B 043203AB.AI

Figure 204 Aileron Trim Tab Travel Board Alternate Installation (UA-1 and After, UB-1 and After, UC-1 and After (Alternate))

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

H. Rudder Travel Board Installation The Rudder Travel Board (4, Table 201) is mounted on the right side of the vertical stabilizer using existing bolt holes for the VOR/LOC antenna (Ref. Figure 205). It is used to measure the travel of the rudder through its full range. (1) Remove the VOR/LOC antenna from the right side of the vertical stabilizer (5) (Ref. Chapter 23-10-00, ANTENNA REMOVAL). (2) Using a stiff plastic scraper, remove the old sealant until the antenna mounting area is flush and smooth with the surface skin. (3) Install the Rudder Travel Board mounting angle (2), with three bolts (3), using the three antenna mounting holes as shown in Figure 205. (4) With assistance, install the rudder travel board (1) on the mounting angle (2) with adjustment bolts (4). Do not tighten adjustment bolts (4). (5) Adjust the rudder travel board (1) to the contour of the vertical stabilizer (5). (6) Ensure all stops (6) contact the vertical stabilizer (5) surface. Tighten the four adjustment bolts (4) on the Rudder Travel Board (1). NOTE: If all stops do not contact the vertical stabilizer surface, contact Hawker Beechcraft Customer Support for consultation 1.800.429.5372.

I. Rudder Travel Board Removal (1) With assistance, remove adjustment bolts (4) and Rudder Travel Board (1) (4, Table 201) from the mounting angle (2) (Ref. Figure 205). (2) Remove the mounting angle (2) from the vertical stabilizer (5). (3) Install the VOR/LOC antenna on the right side of the vertical stabilizer (5) (Ref. Chapter 23-10-00, ANTENNA INSTALLATION).

J. Rudder Trim Tab Travel Board Installation The Rudder Trim Tab Travel Board (5, Table 201) is mounted on the rudder (Ref. Figure 206). The Rudder Trim Tab Travel Board is used to measure the travel of the rudder trim tab through its full range. (1) Position the Rudder Trim Tab Travel Board (4) on the rudder (2) within 0.25 inch above the rudder trim tab cut out (5). (2) Ensure the Rudder Trim Tab Travel Board (4) fits tightly on the rudder (2) and the trailing edge of the rudder trim tab (3) is within 0.12 inch of the stop (1) on the Rudder Trim Tab Travel Board (4). (3) Verify that the 0° reading on the Rudder Trim Tab Travel Board (4) is aligned with the trailing edge of the rudder (2).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

1

2 6

A

B

5 DETAIL

1. RUDDER TRAVEL BOARD 2. MOUNTING ANGLE 3. BOLT 4. ADJUSTMENT BOLT 5. VERTICAL STABILIZER 6. STOPS

A 4

3 2

4 3 DETAIL

B UC27B 042558AA.AI

Figure 205 Rudder Travel Board Installation

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

1. STOP 2. RUDDER 3. RUDDER TRIM TAB 4. RUDDER TRIM TAB TRAVEL BOARD 5. RUDDER TRIM TAB CUT OUT

2

A 4

VIEW LOOKING UP LEFT HAND SIDE

WITHIN 0.12 INCH 2

1

4

5

WITHIN 0.25 INCH

3 FWD VIEW LOOKING AT LEFT SIDE DETAIL

A

Figure 206 Rudder Trim Tab Travel Board Installation

Page 212 Nov 1/09

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UC27B 042610AB.AI

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

K. Elevator Travel Board Installation at HSS 50.00 The Elevator Travel Board (6, Table 201) is mounted on the top of the horizontal stabilizer and is used to measure the travel of the elevator through its full range (Ref. Figure 207). This procedure defines the installation for the Elevator Travel Board that is designed for HSS 50.00. NOTE: One travel board may be used and moved from one side to the other. (1) Locate the rivet center line at Horizontal Stabilizer Station (HSS) 50.00 (1) (Ref. Chapter 6-30-00, STATION LOCATIONS). (2) Mark the location of the rivet center line at HSS 50.00 (1) (Ref. Figure 207). (3) Position the center line of the Elevator Travel Board (2) over the rivet center line at HSS 50.00 (1). (4) Ensure all stops (4) contact the horizontal stabilizer (3) surface. NOTE: Ensure that the stop locations are clearly marked to ensure that the readings are consistent when moving the travel board from one side to the other. (5) Mark the location of the stops (4) on the horizontal stabilizer (3).

L. Elevator Trim Tab Travel Board Installation The Elevator Trim Tab Travel Boards (one left and one right) (7, Table 201) are mounted on the elevators (Ref. Figure 208). The Elevator Trim Tab Travel Boards are used to measure the travel of the elevator trim tabs through their full range. (1) Position the Elevator Trim Tab Travel Board (1) on the elevator (2) between the elevator trim tab (4) and the drain hole grommet (6). The Elevator Trim Tab Travel Board (1) must be within 0.25 inch of the elevator trim tab cut out (3). (2) Ensure the Elevator Trim Tab Travel Board (1) fits tightly on the elevator (2) and the trailing edge of the elevator (2) is within 0.20 inch of the stop (5) on the Elevator Trim Tab Travel Board (1). (3) Verify that the 0° reading on the Elevator Trim Tab Travel Board (1) is aligned with the trailing edge of the elevator (2).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

C

1. RIVET CENTER LINE AT HSS 50.00 2. ELEVATOR TRAVEL BOARD 3. HORIZONTAL STABILIZER 4. STOP

A B

4 3 2

4

FWD

2

4

3 1

VIEW LOOKING DOWN DETAIL

A

2

4 VIEW LOOKING FWD

4

DETAIL

4

B

3

VIEW LOOKING OUTBOARD DETAIL

C

UC27B 042675AA.AI

Figure 207 Elevator Travel Board Installation

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

2 1

A

VIEW LOOKING UP AT LEFT HAND HORIZONTAL STABILIZER

1. ELEVATOR TRIM TAB TRAVEL BOARD 2. ELEVATOR 3. ELEVATOR TRIM TAB CUT OUT 4. ELEVATOR TRIM TAB 5. STOP 6. DRAIN HOLE GROMMET

1 INBD

WITHIN 0.20 INCH 5 FWD

6 4

2

3

WITHIN 0.25 INCH VIEW LOOKING UP (LH SHOWN, RH OPPOSITE) DETAIL

A

UC27B 042609AB.AI

Figure 208 Elevator Trim Tab Travel Board Installation

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M. Flap Travel Board Installation The Flap Travel Board (8, Table 201) is mounted on the top of the wing and is used to measure the travel of the inboard flaps through its full range (Ref. Figure 209, 210 and 211). (1) Determine WS 60.52 on the trailing edge wing skin by measuring 14.02 inches outboard from the center line of the inboard flap, inboard flap track (2), along the trailing edge of the wing skin (3). Mark the location to identify WS 60.52 (Ref. Figure 209). (2) Determine WS 60.52 on the main spar using the following Steps (Ref. Figure 210): (a) For the right side, locate the line of screws (4) at approximately 29 inches outboard from the fuselage (5). Measure from the center of the line of screws 0.5 inch inboard and mark the location. (b) For the left side, locate the line of screws (4) at approximately 28 inches outboard from the fuselage (5). Measure from the center of the line of screws 0.5 inch outboard and mark the location. (3) Starting from WS 60.52, position the Flap Travel Board (8) by moving it outboard until the mounting brackets (9) align with the first set of screw holes along the main spar (3) (1.50 inches offset maximum). Verify the aft edge of the Flap Travel Board (7) is moved outboard the same distance as the forward edge of the Flap Travel Board (10). (4) Mark the new location of the Flap Travel Board (8) as noted in Step (3). (5) If wing access panels (57 and 56, UA-1 and After, UB-1 and After) (18 and 25, UC-1 and After) (Ref. Chapter 6-50-00, WING ACCESS PANELS) are not installed proceed to Step (6). If panels (57 and 56, UA-1 and After; UB-1 and After) (18 and 25, UC-1 and After) are installed, perform the following Steps: (a) Remove the screws from panel (57 and 56, UA-1 and After; UB-1 and After) (18 and 25, UC-1 and After) that align with the screws holes identified in Step (3). NOTE: Do not over tighten the retaining screws in the mounting brackets. Over tightening will not allow aft stop to seat. Do not use spacer plate (6) when wing panel is installed. Tolerance build-up of the wing access panels (57 and 56, UA-1 and After, UB-1 and After) (18 and 25, UC-1 and After) may cause the forward stop not to seat. In this case, remove wing access panels (57 and 56, UA-1 and After, UB-1 and After) (18 and 25, UC-1 and After) and proceed to Step (6). (b) Secure the Flap Travel Board (8) with the retaining screws (1). Proceed to Step (7). (6) If wing access panels (57 and 56, UA-1 and After; UB-1 and After or 18 and 25, UC-1 and After, Ref. Chapter 6-50-00, WING ACCESS PANELS) are not installed perform the following Steps (Ref. Figure 211): (a) Position the spacer plate (4) attached to the Flap Travel Board (1) between the mounting bracket (5) and main spar (6). (b) Place the travel board mounting bracket (5) on the spacer plate (4).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL NOTE: Do not over tighten the retaining screws in the mounting brackets. Over tightening will not allow aft stop to seat. (c) Secure the Flap Travel Board (1) with the retaining screws (3) that align screw holes identified in Step (3). Proceed to Step (7). (7) Ensure that all stops (2) are seated on the wing surface.

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1. FLAP TRAVEL BOARD 2. CENTER LINE OF INBOARD TRACK 3. TRAILING EDGE OF WING SKIN

1 3

A

2

B

WS 60.52 AFT 1.50 INCHES OFFSET MAXIMUM

14.02 INCHES

VIEW LOOKING FORWARD LEFT SIDE DETAIL

A 1

2

AFT

WS 60.52

14.02 INCHES

1.50 INCHES OFFSET MAXIMUM

3 VIEW LOOKING FORWARD RIGHT SIDE DETAIL

B

Figure 209 Flap Travel Board Installation

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UC27B 042559AA.AI

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

5

A

7 2

1

8

B

1

3 6

10

28 INCHES APPROXIMATELY

9 AFT

0.50 INCH

4 WS 60.52

1.50 INCHES OFFSET MAXIMUM

VIEW LOOKING AFT LEFT SIDE DETAIL A

8 10

9

7

2

1

1 3 5

29 INCHES APPROXIMATELY

1. RETAINING SCREW 2. STOP 3. MAIN SPAR 4. LINE OF SCREWS 5. FUSELAGE 6. SPACER PLATE 7. AFT EDGE OF TRAVEL BOARD 8. FLAP TRAVEL BOARD 9. MOUNTING BRACKETS 10. FORWARD EDGE OF TRAVEL BOARD

6

WS 60.52 4 1.50 INCHES 0.50 INCH OFFSET MAXIMUM VIEW LOOKING AFT RIGHT SIDE DETAIL

B

AFT

UC27B 042650AB.AI

Figure 210 Flap Travel Board Installation (With Panels Installed, UC Model Shown, Other Models Similar)

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1. FLAP TRAVEL BOARD 2. STOP 3. RETAINING SCREWS 4. SPACER PLATE 5. MOUNTING BRACKET 6. MAIN SPAR

A 3

2

1

2

B

3

2

6

AFT

4

5 VIEW LOOKING AFT LEFT SIDE DETAIL 3

2

A 1

2 3

6

2

AFT 5

4 VIEW LOOKING AFT RIGHT SIDE DETAIL

B

UC27B 043136AA.AI

Figure 211 Flap Travel Board Installation (Without Panels Installed, UC Model Shown, Other Models Similar)

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2. TRAVEL BOARD CERTIFICATION Travel boards are certified for 24 months or ten uses whichever occurs later from the date of original certification, and are required to be recertified every 24 months or 10 uses, whichever occurs later, thereafter. The following certification requirements are intended for Hawker Beechcraft Corporation approved travel boards only (Ref. Table 201). Once travel boards are certified, it is the responsibility of the owner/operator to keep proper records of certification and subsequent recertification. If a travel board encounters a “catastrophic event” (i.e. dropped or damaged in such a way that could affect the measurements), it is required that the travel board be recertified immediately. These procedures are intended for Hawker Beechcraft Corporation approved travel boards only and are not to be used for fabricating travel boards.

A. Recommended Materials/Equipment For Certification* (1) Surface Table. (2) Height Gage. (3) 12” Knee Blocks. (4) Clamps (to hold down boards). (5) Adjustable Parallel bar. (6) Gage Blocks. (7) Stable Drawing Material (Mylar). (8) Drawing Tools (ruler, compass, protractor). * Equivalent equipment is allowed as long as the certification requirements are met.

B. Certification Set-Up Procedures Various methods can be used for verifying the dimensional requirements of the travel boards. The following methods may be used and have been demonstrated for accuracy for in-service airplanes.

C. Aileron, Elevator and Rudder Method 1 a) Establish a reference line by marking the center of the forward stop (dimension “C”) and the zero degree mark (graduation) on the travel board (Ref. Figures 212, 213, 214 and 215). b) Make all subsequent measurements from the established reference line or the forward stop location as applicable. Method 2 a) Using a stable medium (Mylar), fabricate an inspection aid by drawing the dimensions shown on the respective certification procedures (Ref. Figures 212, 213, 214 and 215). b) From the hinge points, mark all applicable graduations.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL c) Using a flat surface and spacers, lay the travel board onto the inspection aid. NOTE: Dimensions of the inspection aid shall be verified before each use.

D. Aileron Trim Tab, Rudder Trim Tab and Elevator Trim Tabs (1) Using a stable medium (Mylar), fabricate an inspection aid by drawing the dimensions shown on the respective certification procedures (Ref. Figures 216 and 217). (2) From the hinge points, mark all applicable graduations. (3) Using a flat surface and spacers, lay the travel board onto the inspection aid. NOTE: Dimensions on the inspection aid shall be verified before each use.

E. Flap Method 1 a) Draw a reference line approximately the length of the Flap Travel Board (Ref. Figure 218). b) Mark a line the distance the forward stop is above the reference line (Dimension “A”). c) Locate the flap travel board such that the “A” dimensions on the forward stop and the reference line crosses the zero degree mark on the travel board. d) Make all subsequent measurements from the reference line or forward spar reference mark as applicable. Method 2 a) Using a stable medium (Mylar), fabricate an inspection aid by drawing the dimensions shown on the respective certification procedures (Ref. Figure 218). b) From the hinge points, mark all applicable graduations. c) Using a flat surface and spacers, lay the travel board onto the inspection aid. NOTE: Dimensions on the inspection aid shall be verified before each use.

F. Travel Board Rework Rework is only allowed in areas specifically designated in the certification procedures and figures. Any rework shall be documented in the tool records.

G. Aileron Travel Board Certification (UC-1 and After) Requirements (1) Aileron Travel Board (2, Table 201) to be flat with a permanent set not exceeding 0.125 inch along the length of the board. (2) All stops are in good condition, do not show excessive wear and are securely attached. (3) General condition: (a) Identification. (b) Graduation marks are legible.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (c) Other discrepancies noted. (4) Dimensional Requirements (Ref. Table 203 and Figure 212). NOTE: The center of the forward locator (stop) is 0.090 below the reference line. (5) Verify the location and that the following minimum graduation marks exist: (a) Trailing edge up: 0.5°,23°, 24°, 25°, 26°. (b) Trailing edge down: 16°, 17°, 18°, 19°. (c) Zero degrees.

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Table 203 Aileron Travel Board Dimensions (UC-1 and After) Dimension

G

Nominal

Tolerance

A

3.445

± 0.010

B

2.435

± 0.010

C

13.435

± 0.030

D

29.435

± 0.030

E (Reference)

37.435

N/A

F (E+R)

47.435

± 0.060

G

0.090

N/A

R

10.00

± 0.060

A

B

REFERENCE LINE R

C D E F

UC27B 042744AA.AI

Figure 212 Aileron Travel Board Arrangement (UC-1 and After)

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H. Aileron Travel Board Certification (UA-1 and After; UB-1 and After) Requirements (1) Aileron Travel Board (1, Table 201) to be flat with a permanent set not exceeding 0.125 inch along the length of the board. (2) All stops are in good condition, do not show excessive wear and are securely attached. (3) General condition: (a) Identification. (b) Graduation marks are legible. (c) Other discrepancies noted. (4) Dimensional Requirements (Ref. Table 204 and Figure 213). (5) Verify the location and that the following minimum graduation marks exist: (a) Trailing edge up: 0.5°,23°, 24°, 25°, 26°. (b) Trailing edge down: 15°, 16°, 17°, 18°. (c) Zero degrees.

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Table 204 Aileron Travel Board Dimensions (UA-1 and After; UB-1 and After) Dimension

Nominal

Tolerance

A

4.460

± 0.010

B

2.525

± 0.010

C

13.60

± 0.030

D

38.46

± 0.030

E (Reference)

44.07

N/A

F (E+R)

55.57

± 0.060

R

11.50

± 0.060

A B

REFERENCE LINE

R

C D E F

UC27B 042746AA.AI

Figure 213 Aileron Travel Board Arrangement (UA-1 and After; UB-1 and After)

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I. Rudder Travel Board Certification Requirements (1) Rudder Travel Board (4, Table 201) to be flat with a permanent set not exceeding 0.125 inch along the length of the board. (2) All stops are in good condition, do not show excessive wear and are securely attached. (3) General condition: (a) Identification. (b) Graduation marks are legible. (c) Other discrepancies noted. (d) Attachment hardware is present and in good condition. (4) Dimensional Requirements (Ref. Table 205 and Figure 214). (5) Verify the location and that the following minimum graduation marks exist: (a) Trailing edge left: 25°, 26°. (b) Trailing edge right: 25°, 26°. (c) Zero degrees. (6) Rework information: Permissible to rework the four holes (or slots) in the adjustment handle to allow locating stops to rest against the vertical stabilizer as required.

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Table 205 Rudder Travel Board Dimensions Dimension

Nominal

Tolerance

A

4.980

± 0.010

B

4.120

± 0.010

C

11.010

± 0.030

D

35.450

± 0.030

E (Reference)

44.50

N/A

F (E+R)

66.10

± 0.060

R

21.60

± 0.060

ADJUSTMENT ANGLE

A

B

C

REFERENCE LINE

R D E F

UC27B 043138AA.AI

Figure 214 Rudder Travel Board Arrangement

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

J. Elevator Travel Board Certification Requirements (1) Elevator Travel Board (6, Table 201) to be flat with a permanent set not exceeding 0.125 inch along the length of the board. (2) All stops are in good condition, do not show excessive wear and are securely attached. (3) General condition: (a) Identification. (b) Graduation marks are legible. (c) Other discrepancies noted. (4) Dimensional Requirements (Ref. Table 206 and Figure 215). (5) Verify the location and that the following minimum graduation marks exist: (a) Trailing edge up: 20°, 21°. (b) Trailing edge down: 7°,14°, 15°. (c) Zero degrees.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Table 206 Elevator Travel Board Dimensions Dimension

Nominal

Tolerance

A

1.970

± 0.010

B

1.885

± 0.010

C

6.075

± 0.030

D

25.240

± 0.030

E (Reference)

31.975

N/A

F (E+R)

46.475

± 0.060

R

14.50

± 0.060

A

B

R C D E F UC27B 042748AA.AI

Figure 215 Elevator Travel Board Arrangement

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

K. Aileron Trim Tab and Rudder Trim Tab Travel Board Certification Requirements (1) Aileron Trim Tab and Rudder Trim Tab Travel Board (3 and 5, Table 201) general condition: (a) Identification. (b) Graduation marks are legible. (c) Other discrepancies noted. (2) Dimensional Requirements (Ref. Table 207 and Figure 216). (3) Rework information: (a) Permissible to replace loose rivets as long as dimensional requirements are maintained.

Table 207 Aileron Trim Tab and Rudder Trim Tab Travel Board Dimensions

Trim Tab

Minimum Graduations Marks Required*

A Value

B Value

Tolerance

Aileron

0.500

3.419

± 0.015

Up: 16.5°, 15°, 13.5° Down: 16.5°, 15°, 13.5° Zero Degrees

Rudder

1.093

6.900

± 0.015

Left: 16.5°, 15°, 11°, 10°, 9° Right: 16.5°, 15°, 11°, 10°, 9° Zero Degrees

*Half degree increments are not required if the whole degree marks are adjacent (i.e. if 16° and 17° exist, 16.5° is not required).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

HINGE POINT CHART 6 MARKS A

0 DEG A

ANGLE

B

TRAVEL BOARD OVERLAID FOR REFERENCE

Figure 216 Aileron Trim Tab and Rudder Trim Tab Travel Board Arrangement

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UC27B 042753AA.AI

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

L. Elevator Trim Tab Travel Board Certification Requirements (1) Elevator Trim Tab Travel Board (7, Table 201) general condition: (a) Identification. (b) Graduation marks are legible. (c) Other discrepancies noted. (2) Dimensional Requirements (Ref. Table 208 and Figure 217). (3) Rework information: (a) Permissible to replace loose rivets as long as dimensional requirements are maintained. (b) Elevator Trim Tab Travel Board: Permissible to trim rubber pad and angle to clear the drain grommet (scupper) on the lower surface of the elevator.

Table 208 Elevator Trim Tab Travel Board Dimensions

Trim Tab

A Value

B Value

C Value

Tolerance

Elevator

0.407

4.168

3.502

± 0.015

Minimum Graduations Marks Required* Up: 5.5°, 5°, 2.5°, 1.5° Down: 16°, 15°, 7°, 6°, 5° Zero Degrees

*Half degree increments are not required if the whole degree marks are adjacent (i.e. if 16° and 17° exist, 16.5° is not required).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

HINGE POINT

A A ZERO DE

GR E E

C

B TRAVEL BOARD OVERLAID FOR REFERENCE

UC27B 043273AA.AI

Figure 217 Elevator Trim Tab Travel Board Arrangement (LH Shown, RH Opposite)

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

M. Flap Travel Board Certification Requirements (1) Flap Travel Board (8, Table 201) to be flat with a permanent set not exceeding 0.125 inch along the length of the board. (2) All stops are in good condition, do not show excessive wear and are securely attached. (3) General condition: (a) Identification. (b) Graduation marks are legible. (c) Other discrepancies noted. (d) Attachment hardware is present and in good condition. (4) Dimensional Requirements (Ref. Table 209 and Figure 218). (5) Verify the that the following minimum graduation marks exist: (a) Trailing edge down: 0.5°, 1°, 1.5°, 2°, 9°, 10°, 11°, 16.5°, 17°, 17.5°, 18°, 18.5°, 19°, 20°, 20.5°, 21°, 33°, 34°, 35°, 36°. (b) Zero degrees. (6) Rework information: Permissible to rework the adjustment angle to allow the two locating stops to contact the wing surface. This can be done by slotting the angle, milling the lower portion of the adjustment angle (side touching the wing) or using smaller diameter (# 10) bolts.

Table 209 Flap Travel Board Dimensions Dimension

Nominal

Tolerance

A

10.373

± 0.010

B

5.688

± 0.010

C

1.541

± 0.060

D

37.113

± 0.060

F

61.253

± 0.060

G

10.443

± 0.030

H (10°)

3.466

± 0.030

J (20°)

6.291

± 0.030

K (35°)

9.986

± 0.030

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

ADJUSTMENT ANGLE D C

A G

A B REFERENCE LINE F

H J K

B

FORWARD SPAR REFERENCE (DATUM)

"MILL STEP" MARK

REFERENCE LINE

EDGE OF BOARD

LINE PERPENDICULAR TO REFERENCE LINE

DETAIL

DETAIL

B

A UC27B 042751AA.AI

Figure 218 Flap Travel Board Arrangement

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UC27C 060912AA.AI

Figure 219 Reading a Travel Board

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FLIGHT CONTROLS AILERONS MAINTENANCE PRACTICES

27-10-00 200200

1. PROCEDURES A. Removal (UA-1 and After; UB-1 and After) Only the left aileron incorporates an aileron trim tab. The aileron trim tab pushrod connects directly to the actuator mounted forward of the rear spar. The hinge brackets inserted between the aileron skin and spar are secured with four screws. The aileron pushrod is connected to one hinge bracket to transmit movements to the aileron. (1) Install a red tag to the pilots control wheel with the words “Do not Operate, Maintenance in Progress”. (2) Perform STATIC DISCHARGER REMOVAL procedure (Ref. Chapter 23-60-00) on the aileron to be removed. (3) On the LH aileron, remove the bolt, washers and nut connecting the underside of the aileron trim tab to the push-pull rod from the aileron trim tab actuator. The double-clevis ends on the trim tab actuator push-pull rod are designed to tighten the outer clevis with a binding action against the clevis bolt and the inner clevis, which removes freeplay of the bolt in the hole. Loosen the outer clevis first (the larger nut) before removing the clevis bolt. (4) Remove the double clevis end of the push-pull rod so that the push-pull rod will pass though the left side aileron during removal. (5) Remove screws (11), washers (2) and nuts (1) attaching bonding jumpers (7) to aileron (13) at the outboard, middle and inboard hinge points (Ref. Figure 201). (6) While providing support at each end of the aileron, remove screws (9) from the top and bottom surface of aileron (13) at the outboard, middle and inboard hinge points. (7) Remove the aileron (13) from the wing. Take care not to damage the skin while removing the trim tab actuator push-pull rod from the LH aileron and aileron trim tab assembly.

B. Installation (UA-1 and After; UB-1 and After) WARNING: Failure to follow the correct procedure for AILERON INSTALLATION could result in partial/complete loss of the aileron resulting in degradation of flying qualities and or loss of control of the aircraft. NOTE: Any repair, modification, painting or replacement of the aileron or aileron trim tab requires balancing (Ref. Chapter 57-50-00). (1) Prepare the area where the bonding cables attach for electrical bonding (Ref. Chapter 20-00-01, PREPARATION OF SURFACE). (2) On the LH aileron (13), carefully guide the aileron trim tab push-pull rod through the aileron (13) (Ref. Figure 201).

27-10-00

Page 201 Aug 1/12

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (3) With assistance, align the inboard, middle and outboard hinge points of aileron (13) with hinge halves (6) on the wing. Then loosely install two screws (9) (one screw in the top and one in the bottom) through each hinge point. Pull on the aileron to confirm that the screws (9) are properly installed. NOTE: If there is no movement when the aileron is pulled on, the bolts are properly installed. (4) Install screws (9) to the top and bottom surface of the aileron (13) at the outboard, middle and inboard hinge points. (5) Install screws (11), washers (2) and nuts (1) attaching the bonding jumpers (7) to the aileron (13) at the outboard, middle and inboard hinge points. WARNING: Any time the aileron trim tab push-pull rods are installed or adjusted, the inspection holes near the ends of the rods must be checked to ascertain that the threads of the end fittings are visible. (6) Install the double clevis end of the trim tab actuator push-pull rod, so that the push-pull rod assembly is 8.58 ± 0.06 inches from the centerline of the mounting hole of both clevis ends. (7) On the left-hand aileron, connect the trim tab actuator push-pull rod to the horn on the aileron trim tab. NOTE: The double-clevis ends on the trim tab actuator push-pull rod are designed to tighten the outer clevis with a binding action against the clevis bolt and the inner clevis, which removes freeplay of the bolt in the hole. Install the clevis bolt and tighten the large jam nut on the outer clevis. With the flaps fully retracted and the aileron in the neutral position, the clearances noted must be maintained (Ref. Figure 203). The gap between the aileron and the outboard flap should be constant within ± 0.06 inch from the leading edge to the trailing edge. These dimensions do not apply to the lower forward area where the aileron tapers outboard. (8) Remove the red tag from the pilots control wheel. (9) Perform the AILERON FUNCTIONAL CHECK procedure (Ref. 27-10-03). (10) Perform the AILERON TRIM TAB FUNCTIONAL CHECK procedure (Ref. 27-10-07). (11) Perform the AILERON FREEPLAY CHECK procedure in this section. (12) Perform STATIC DISCHARGER INSTALLATION procedure (Ref. Chapter 23-60-00).

C. Removal (UC-1 and After) The aileron is operated by an arm on the aileron bellcrank located at the inboard end of each aileron. The arm is connected through a tapered pin to a hinged yoke mounted in the aileron just inboard of the inboard aileron hinge. Only the left aileron incorporates an aileron trim tab. The aileron trim tab push-pull rod connects directly to the actuator mounted on the aft side of the rear spar. (1) Install a red tag to the pilots control wheel with the words “Do not Operate, Maintenance in Progress”.

Page 202 Aug 1/12

27-10-00

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (2) Perform STATIC DISCHARGER REMOVAL procedure (Ref. Chapter 23-60-00) on the aileron to be removed. NOTE: Do not disturb the adjustment of the rod end on the push-pull rod. (3) On the LH aileron, remove the bolt, washers and nut connecting the underside of the aileron trim tab to the push-pull rod from the aileron trim tab actuator. The double-clevis ends on the trim tab actuator push-pull rod are designed to tighten the outer clevis with a binding action against the clevis bolt and the inner clevis, which removes freeplay of the bolt in the hole. Loosen the outer clevis first (the larger nut) before removing the clevis bolt. (4) Remove screws (9), washers (5) and nuts (4) attaching the bonding jumpers (6) to the aileron (13) at the outboard, middle and inboard hinge points (Ref. Figure 202). (5) While providing support at each end of the aileron, remove the safety wire, bolt (8) and washer (7) securing the inboard hinge bracket on the wing. (6) Remove six screws (11) attaching the aileron (13) to the middle hinge clevis (10). (7) Remove four screws (11) attaching the aileron (13) to the outboard hinge clevis (10). (8) Remove aileron (13) from the wing, taking care not to cause damage while disengaging the yoke tapered pin (14) on either aileron (13), or while removing the trim tab actuator push-pull rods from the LH aileron and trim tab assembly.

D. Installation (UC-1 and After) NOTE: Any repair, modification, painting or replacement of the aileron or aileron trim tab requires balancing (Ref. Chapter 57-50-00). (1) Prepare the area where the bonding cables attach for electrical bonding (Ref. Chapter 20-00-01, PREPARATION OF SURFACE). (2) Lightly lubricate the aileron yoke (12) tapered pin (14) and bellcrank arm bearing (16) with grease (1, Table 2, 27-00-00) (Ref. Figure 202). (3) On the LH aileron (13), carefully guide the aileron trim tab push-pull rod through the aileron (13). NOTE: On the inboard aileron hinge fitting (2), the hat portion of the bushing (3) does not rest against the fitting (2). The extra grip length bushing (3) is used intentionally to avoid aileron hinge bearing becoming pinched if the hinge bolt is over-torqued (Ref. Figure 202, Detail E). (4) With assistance, align the aileron with the middle and outboard hinge clevises (10) and the inboard hinge bracket on the wing, engage the actuator tapered pin (14) into the bellcrank arm bearing (16). (5) Loosely install six screws (11) securing the aileron (13) to the middle hinge clevis (10). (6) Loosely install four screws (11) securing the aileron (13) to the outboard hinge clevis (10). (7) Align and install the bolt (8) and washer (7) securing the inboard hinge to the wing (15). Safety wire the bolt (8). (8) Tighten the screws (11) at the outboard and middle hinge points.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (9) Install screws (9), washers (5) and nuts (4) attaching the bonding jumpers (6) to the aileron (13) at the outboard, middle and inboard hinge points. WARNING: Any time the aileron trim tab push-pull rods are installed or adjusted, the inspection holes near the ends of the rods must be checked to ascertain that the threads of the end fittings are visible. (10) On the left-hand aileron, connect the trim tab actuator push-pull rod to the horn on the aileron trim tab. NOTE: The double-clevis ends on the trim tab actuator push-pull rod are designed to tighten the outer clevis with a binding action against the clevis bolt and the inner clevis, which removes freeplay of the bolt in the hole. Install the clevis bolt and tighten the large jam nut on the outer clevis. With the flaps fully retracted and the aileron in the neutral position, the clearances noted must be maintained (Ref. Figure 203). The gap between the aileron and the outboard flap should be constant within ± 0.06 inch from the leading edge to the trailing edge. These dimensions do not apply to the lower forward area where the aileron tapers outboard. (11) Remove the red tag from the pilots control wheel. (12) Perform the AILERON FUNCTIONAL CHECK procedure (Ref. 27-10-03). (13) Perform the AILERON TRIM TAB FUNCTIONAL CHECK procedure (Ref. 27-10-07). (14) Perform the AILERON FREEPLAY CHECK procedure in this section. (15) Perform STATIC DISCHARGER INSTALLATION procedure (Ref. Chapter 23-60-00).

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Figure 201 Aileron Installation (UA-1 and After; UB-1 and After)

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2

1. BUSHING 2. FITTING 3. BUSHING 4. NUTS 5. WASHERS 6. BONDING JUMPERS 7. WASHER 8. BOLT 9. SCREWS 10. CLEVIS 11. SCREWS 12. YOKE 13. AILERON 14. TAPERED PIN 15. WING 16. BEARING

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B Figure 202 Aileron Installation (UC-1 and After)

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Figure 203 Aileron Clearance

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E. Freeplay Check NOTE: Movement or jarring of the airplane will invalidate the aileron freeplay readings. The airplane should be hangared and no personnel in or on the airplane during the freeplay check. (1) For UA-1 and After; UB-1 and After perform the following Steps: (a) Remove left and right lower wing access panels 44 and 45 (Ref. Chapter 6-50-00, WING ACCESS PANELS). (b) Install rig pin (3) (11, Table 1, 27-00-00) through the aileron outboard wing bellcrank (1) (Ref. Figure 206). (2) For UC-1 and After perform the following Steps: (a) Remove the passenger seat(s) as required to gain access to floor access panel 16E (Ref. Chapter 25-20-00, SEAT REMOVAL). (b) Remove the carpet as required to gain access to floor access panel 16E (Ref. Chapter 25-20-01). (c) Remove floor access panel 16E (Ref. Chapter 6-50-00, FLOOR ACCESS PANELS). (d) Install rig pin (1) (7, Table 1, 27-00-00) through the aileron quadrant support bracket (2) and the aileron quadrant (3) (Ref. Figure 204). (3) Perform the AILERON TRAVEL BOARD INSTALLATION procedure (Ref. 27-00-02). (4) Visually inspect the aileron for any damage, for integrity of the hinge attach points, and for tightness of the actuating system. (5) Apply a small piece of masking tape (for paint protection) one inch forward of the aileron trailing edge just left or right of the aileron travel board (6). This will be the point of pressure against the aileron (3) by the push-pull scale (Ref. Figure 205). (6) Apply another piece of masking tape in the corresponding position on the bottom surface of the aileron (3) for the same purpose. (7) Attach a scale (8) or dial indicator (1) to the aileron travel board (6) so the up and down movement can be measured at the aileron (3) trailing edge. (8) If using the dial indicator (1) perform the following Steps: (a) Position the dial indicator (1) so the stem (2) is 0.50 inch from the trailing edge of the aileron (3) and is depressed 0.10 inch when in contact with the aileron (3) surface initially. Turn the rotating face of the dial indicator (1) to zero. Do not reset during the checking procedure. (b) With a push-pull scale (6, Table 1, 27-00-00) against the top of the aileron (3), apply four pounds of downward load. Record the dial reading. (c) Apply a four pound upwards load on the bottom surface of the aileron (3). Record the dial reading.

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(9) If using the scale (8) perform the following Steps: (a) Attach the scale (8) to the aileron travel board (6) with strips of tape (7). (b) With a push-pull scale (6, Table 1, 27-00-00) against the top of the aileron (3), apply four pounds of downward load. Record the scale reading. (c) Apply a four pound upward load on the bottom surface of the aileron (3). Record the scale reading. (10) For UA and UB, bladder wing airplanes, the maximum freeplay limit is 0.12 inch. For UC airplanes, the maximum freeplay travel limit is 0.06 inch. Excess movement must be corrected. (11) Perform Steps (3) through (10) for the opposite aileron. (12) If freeplay limits are exceeded, inspect all components for cracks and wear, and repair or replace as required. (13) For UC-1 and After perform the following Steps: (a) Remove the rig pin (1) from the aileron quadrant (3) (Ref. Figure 204). (b) Install floor access panel 16E (Ref. Chapter 6-50-00, FLOOR ACCESS PANELS). (c) Install the carpet as required (Ref. Chapter 25-20-01). (d) Install the passenger seat(s) as required (Ref. Chapter 25-20-00, SEAT INSTALLATION). (14) For UA-1 and After; UB-1 and After perform the following Steps: (a) Remove rig pin (3) (11, Table 1, 27-00-00) from the aileron outboard wing bellcrank (1) (Ref. Figure 206). (b) Install left and right lower wing panels 44 and 45 (Ref. Chapter 6-50-00, WING ACCESS PANELS). (15) Remove the aileron travel board.

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1. RIG PIN 2. AILERON QUADRANT SUPPORT BRACKET 3. AILERON QUADRANT

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Figure 204 Aileron Quadrant Rig Pin Installation

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Figure 205 Aileron Freeplay Fixture

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1. BELLCRANK 2. LOWER SUPPORT BRACKET 3. RIG PIN 4. UPPER SUPPORT BRACKET

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Figure 206 Aileron Outboard Wing Bellcrank Rig Pin Installation (UA-1 and After; UB-1 and After)

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2. AILERON GROUND ADJUSTABLE TRIM TAB A. Adjustment NOTE: If wing heavy condition cannot be corrected using this procedure alone the FLAP ADJUSTMENT TO CORRECT A WING HEAVY CONDITION procedure (Ref. 27-50-05) may be used in conjunction with this procedure. (1) Perform the AILERON FUNCTIONAL CHECK procedure to ensure proper aileron system operation prior to performing the following procedure (Ref. 27-10-03). (2) Perform the FLAP SYSTEM FUNCTIONAL CHECK procedure to ensure proper flap alignment prior to performing the following procedure (Ref. 27-50-05). NOTE: The aileron ground adjustable trim tab is located on the right aileron only and is used to correct aileron load with the aileron trim set to zero (Ref. Figure 207). (3) The line of rivets on the trailing edge of the aileron should be stabilized while the bendable trim tab is being adjusted. This will reduce the risk of bending the aileron trailing edge out of symmetry as well as causing paint to crack. A shop aid used to stabilize the trailing edge can be built by performing the following Steps: (a) Figure 208 shows a top and bottom rivet brace that fits over the rivet line and are clamped by toggle pliers. Fabricate both rivet braces by cutting two iron bars to the dimensions shown in Figure 208. The top and bottom rivet braces are identical in form. (b) Obtain two pair of toggle pliers and weld the lower jaw of each pliers to a corresponding cut out in one of the rivet braces as shown in Figure 209 Sheet 1. The rivet brace must be oriented with the slot facing up and the aft surface facing the pliers (Ref. Figure 209 Sheets 1 thru 3). (c) Drill out and tap the ends of the toggle pliers upper jaws to accept a three inch bolt with a stop nut (Ref. Figure 209, Sheet 1 of 3). (d) Obtain two swivel fittings then drill and tap the fittings to accept the three inch bolts. Weld each of the swivel fittings to the cut outs in the remaining rivet brace (Ref. Figure 210). The swivel fittings must be able to rotate freely within the cut outs such that they can be screwed onto the bolts to make contact with the upper jaw ends. Use the stop nut to keep the bolt from turning as the swivel fitting is torqued onto the jaw ends. (e) Surfaces of the shop aid that come in contact with the aileron should be covered with duct tape to prevent abrasion of rivet heads and paint. (4) Secure the shop aid to the line of rivets that fasten the ground adjustable aileron trim tab to the right aileron (Ref. Figure 211). NOTE: Do not bend the Ground Adjustable Aileron Trim Tab more than 30° up or down. (5) Use a modified hand seamer (Ref. Figure 212) to bend the ground adjustable aileron trim tab to correct aileron load imbalance (Ref. Figure 213).

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(6) Perform flight test. (7) Repeat Steps (4) thru (6) as necessary until aileron load is equal during flight with aileron trim set approximately to zero.

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Figure 207 Ground Adjustable Trim Tab

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Figure 208 Top and Bottom Rivet Braces

Figure 209 (Sheet 1 of 3) Bottom Rivet Brace Attachment to Toggle Pliers

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Figure 209 (Sheet 2 of 3) Bottom Rivet Brace Attachment to Toggle Pliers

Figure 209 (Sheet 3 of 3) Bottom Rivet Brace Attachment to Toggle Pliers

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Figure 210 (Sheet 1 of 3) Top Rivet Brace Attachment to Toggle Pliers

Figure 210 (Sheet 2 of 3) Top Rivet Brace Attachment to Toggle Pliers

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Figure 210 (Sheet 3 of 3) Top Rivet Brace Attachment to Toggle Pliers

Figure 211 Stabilize the Aileron Trailing Edge Page 220 Aug 1/12

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Figure 212 Hand Seamer

Figure 213 Bend Aileron Ground Adjustable Trim Tab Using Hand Seamer

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FLIGHT CONTROLS CONTROL WHEEL MAINTENANCE PRACTICES

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1. PROCEDURES A. Removal NOTE: This procedure is typical for both the pilot’s and copilot’s control wheel. (1) Disconnect subpanel connector (3) at subpanel (2) (Ref. Figure 201). (2) Remove faceplate (4) from control wheel (1) by removing the screws (5). (3) If the control wheel has the optional digital or analog clock faceplate installed, perform the following Step: (a) Disconnect the electrical connector (6) from the optional clock faceplate (4). (4) Remove the three attaching nuts (8), washers (9) and screws (7) securing the control wheel to the control column (10).

B. Installation NOTE: This procedure is typical for both the pilot’s and copilot’s control wheel. (1) Position the control wheel (1) on the control column (10) and install screws (7), washers (9) and nuts (8) (Ref. Figure 201). (2) If the control wheel has the optional digital or analog clock faceplate installed, perform the following Step: (a) Connect the electrical connector (6) for the optional clock faceplate (4). (3) Position faceplate (4) on control wheel (1) and secure with screws (5). (4) Connect subpanel connector (3) at subpanel (2).

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1. CONTROL WHEEL 2. SUBPANEL 3. SUBPANEL CONNECTOR

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Figure 201 (Sheet 1 of 2) Control Wheel Installation

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Figure 201 (Sheet 2 of 2) Control Wheel Installation

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2. CONTROL COLUMN CABLE A. Removal (1) Disconnect the aileron cables from the control column cable at the turnbuckles. (2) Paint one tooth on each of the control actuating sprockets and its corresponding chain link to ensure proper alignment of the control wheels at reinstallation of the chain. (3) Loosen the cable turnbuckle in the center of the control column horizontal cross member, remove the cable retaining pins, unsafety the chain from the sprockets and remove the control column cable and chain assembly.

B. Installation (1) Place the control column cable and chain assembly on the cross member of the control column with the painted links of the chains engaging the corresponding painted sprocket teeth. Install the cable retaining pins and safety wire the chain to the sprocket. (2) Connect the aileron cables at the turnbuckles on the control column. (3) Perform the AILERON CONTROL SYSTEM RIGGING procedure (Ref. 27-10-03).

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FLIGHT CONTROLS AILERON CABLES MAINTENANCE PRACTICES

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1. FUSELAGE CABLE A. Removal (1) Attach a red tag to the control wheels with the words “Do Not Operate, Maintenance In Progress”. (2) Remove both flight compartment seats (Ref. Chapter 25-10-00, SEAT REMOVAL). (3) Remove the flight compartment carpet as required to access flight compartment floor access panels 3, 4 and 5. (4) Remove the control column boot located forward of the pedestal. (5) Remove pedestal side panels. (6) Remove the control panels from the pedestal to gain access to the aileron fuselage cables. Refer to the applicable maintenance procedure. (7) Remove flight compartment floor access panels 3, 4 and 5 (Ref. Chapter 6-50-00, FLOOR ACCESS PANELS). (8) Remove the passenger seats as required to access floor board access panels 16A, 16B, 16C, 16D and 16E (Ref. Chapter 25-20-00, SEAT REMOVAL). (9)

Remove passenger compartment carpet as required to access floor access panels 16A, 16B, 16C, 16D and 16E (Ref. Chapter 25-20-01).

(10) Remove the center aisle spar aft ramp. (11) Remove floor access panels 16A, 16B, 16C, 16D and 16E (Ref. Chapter 6-50-00, FLOOR ACCESS PANELS). (12) If equipped, identify, tag and disconnect the Supplemental Type Certificate (STC) Flight Data Recorder (FDR) sensor bridle clamp located on the aileron fuselage cable. Refer to the STC holders instructions. NOTE: Each turnbuckle (4 and 5) barrel has a groove (7) at one end to identify the left-hand threaded end (Ref. Figure 203). (13) Remove safety clips (6) from turnbuckles (4 and 5). (14) Attach a tag with the words “left-hand threads terminal end” to the lower end of the left turnbuckle (5). (15) Disconnect the left-hand threads terminal end (3) from the left turnbuckle (5) and attach a feed line to the terminal end. Label the feed line with the words “left-hand threads terminal end”. (16) Attach a tag with the words “right-hand threads terminal end” to the lower end of the right turnbuckle (4).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (17) Disconnect the right-hand threads terminal end (2) from the right turnbuckle (4) and attach a feed line to the terminal end. Label the feed line with the words “right-hand threads terminal end”. (18) Disconnect the aileron cable locking plates on the aileron quadrant (Ref. Figure 201, UA-1 and After; UB-1 and After or Figure 202, UC-1 and After). (19) Remove the cable guard pins from pulley brackets. Refer to Figure 201, UA-1 and After; UB-1 and After or Figure 202, UC-1 and After for general pulley locations. NOTE: It may be necessary to remove the pulleys if cable passage is restricted. If cable is to be reused, take care to keep the cable clean and free from damage. Coil cable loosely no tighter than a 19 inch diameter. (20) With assistance, feed the aileron cables through the fuselage pulleys and pull the cables through the access openings at the aileron quadrant. (21) Detach the feed lines from the aileron fuselage cable terminal ends leaving the feed lines in place.

B. Installation WARNING: Perform all Steps of this procedure, in the order listed. Do not skip any Steps of this procedure. Failure to do so may result in injury to personnel and damage to equipment. CAUTION: Where the cables pass through structure, the areas of possible contact between the control cables and adjacent structure must be protected with protective elements such as grommets, rub strips, blocks or guide fairings. Where contact of control cables does occur with the protective elements, a force no greater than eight (8) ounces shall be required to move the cable to a position of no contact. At no time should flight control cables contact metal structure with the protective elements removed. (1) If a used cable is to be installed, clean with solvent (2, Table 2, 27-00-00) and check for corrosion and damage. Replace cable if necessary. Dip the cable in corrosion preventive compound (4, Table 2, 27-00-00). Excess should be removed by wiping with a clean rag. (2) Attach a tag labeled “left-hand threads terminal end” to the aileron fuselage cable left-hand threads terminal end. (3) Attach a tag labeled “right-hand threads terminal end” to the aileron fuselage cable right-hand threads terminal end. (4) Attach the right-hand threads terminal end to the feed line labeled “right-hand threads terminal end”. (5) Attach the left-hand threads terminal end to the feed line labeled “left-hand threads terminal end”. NOTE: More than one person will be required to route the aileron fuselage cable. Take precautions to keep the cable clean and free from damage. (6) Route the cables through the fuselage as follows: NOTE: It is permissible to install cable guard pins and any removed pulleys as the cable is being routed.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (a) From the aileron quadrant route cable (7) left-hand thread, through the left-hand pulley at FS 293 (Ref. Figure 201 UA-1 and After; UB-1 and After or Figure 202 UC-1 and After). (b) Route cable (7) through the left-hand pulley at FS 161. (c) Route cable (7) through the left-hand pulley at FS 145. (d) Route cable (7) through the left-hand pulley at FS 127. (e) Route cable (7) through the left-hand pulley at FS 103. (f) From the aileron quadrant route cable (8) right-hand thread, through the right-hand pulley at FS 293. (g) Route cable (8) through the right-hand pulley at FS 161. (h) Route cable (8) through the right-hand pulley at FS 127. (i) Route cable (8) through the right-hand pulley at FS 103. CAUTION: Do not over torque the cable locking plate attachment screws or damage to the quadrant may occur. Maximum torque will not exceed 15 inch-pounds. (j) Connect the cables to the aileron quadrant by installing the cable locking plates (34 and 37) and attachment screws (36 and 39). Safety wire the screws with 0.032-inch diameter safety wire. (k) Ensure all cable guard pins and pulleys are installed in the pulley brackets. (7) Lubricate turnbuckles (4 and 5) with grease (1, Table 2, 27-00-00) for corrosion protection prior to installation (Ref. Figure 203). (8) Remove feed line, and attach the left-hand threads terminal end (3) to the left turnbuckle (5). (9) Remove feed line, and attach the right-hand threads terminal end (2) to the right turnbuckle (4). (10) Tension the aileron fuselage cables sufficient to prevent slack. (11) Ensure that the fuselage cables are routed properly by verifying that the cables have been routed exactly as described in Step (6). Ensure the cables are engaged in the pulley grooves and all pulley guard pins are installed. (12) Remove all tape from the cables and turnbuckles. (13) Perform the AILERON OPERATIONAL CHECK procedure (Ref. 27-10-03). (14) Perform the AILERON CONTROL SYSTEM RIGGING procedure (Ref. 27-10-03). (15) Ensure safety clips are installed on all turnbuckles. (16) Install the Supplemental Type Certificate (STC) Flight Data Recorder (FDR) sensor bridle clamp, if equipped. Refer to the STC holders instructions for installation and calibration. (17) Install floor access panels 16A, 16B, 16C, 16D and 16E (Ref. Chapter 6-50-00, FLOOR ACCESS PANELS).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (18) Install the center aisle spar aft ramp. (19) Install passenger compartment carpet (Ref. Chapter 25-20-01). (20) Install the passenger seats (Ref. Chapter 25-20-00, SEAT INSTALLATION). (21) Install flight compartment floor access panels 3, 4 and 5 (Ref. Chapter 6-50-00, FLOOR ACCESS PANELS). (22) Install the control column boot located forward of the pedestal. (23) Install the control panels removed from the pedestal. (24) Install pedestal side panels. (25) Install the flight compartment carpet (Ref. Chapter 25-20-01). (26) Install both flight compartment seats (Ref. Chapter 25-10-00, SEAT INSTALLATION). (27) Remove the red tag from the control wheels.

2. WING OUTBOARD CABLE A. Removal (UA-1 and After; UB-1 and After) (1) Attach a red tag to the control wheels and the flap lever with the words “Do Not Operate, Maintenance In Progress”. (2) Remove lower wing access panels 7, 35, 36, 37, 38, 43, 44, 45, 46 and 47 (Ref. Chapter 6-50-00, WING ACCESS PANELS). (3) Remove upper wing access panels 4, 29 and 30 (Ref. Chapter 6-50-00, WING ACCESS PANELS). NOTE: The aileron wing cables from the aileron quadrant to the turnbuckles and from the turnbuckles to the outboard bellcranks are not interchangeable. Each cable should be identified for correct location prior to removal. Each turnbuckle (2 and 8) barrel has a groove (9) at one end to identify the left-hand threaded end (Ref. Figure 204). (4) Move aileron wing cables to ensure easy access to the turnbuckles. (5) Install cable block (11) on the aileron wing inboard cables (1 and 10) to prevent loss of cable tension. (6) Remove safety clips (7) from both turnbuckles (2 and 8). (7) Attach a tag with the words “outboard cable left-hand threads terminal end” to the outboard end of the forward turnbuckle (2). (8) Disconnect the left-hand threads terminal end (3) from the forward turnbuckle (2) and attach a feed line to the terminal end. Label the feed line with the words “left-hand threads terminal end”. (9) Attach a tag with the words “outboard cable right-hand threads terminal end” to the outboard end of the aft turnbuckle (8). Page 204 Nov 1/09

27-10-02

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (10) Disconnect the right-hand threads terminal end (5) from the aft turnbuckle (8) and attach a feed line to the terminal end. Label the feed line with the words “right-hand threads terminal end”. NOTE: The cable guard pin for the pulley at WS 182 is installed head up and cotter pinned. (11) Remove the cable guard pins from pulley brackets. (12) Remove cotter pin (18), nut (17), washer (16) and bolt (15) from the outboard wing bellcrank (Ref. Figure 201). (13) Remove cotter pin (19), nut (20), washer (21) and bolt (22) from the outboard wing bellcrank. (14) Remove cables (4 and 5) from the outboard wing bellcrank. NOTE: If cable is to be reused, take care to keep the cable clean and free from damage. Coil cable loosely no tighter than a 29 inch diameter. (15) Pull the cables (4 and 5) through the access openings at the outboard wing bellcrank. (16) Detach the feed lines from the aileron wing outboard cable terminal ends leaving the feed lines in place.

B. Installation (UA-1 and After; UB-1 and After) WARNING: Perform all Steps of this procedure, in the order listed. Do not skip any Steps of this procedure. Failure to do so may result in injury to personnel and damage to equipment. CAUTION: Where the cables pass through structure, the areas of possible contact between the control cables and adjacent structure must be protected with protective elements such as grommets, rub strips, blocks or guide fairings. Where contact of control cables does occur with the protective elements, a force no greater than eight (8) ounces shall be required to move the cable to a position of no contact. At no time should flight control cables contact metal structure with the protective elements removed. (1) If a used cable is to be installed, clean with solvent (2, Table 2, 27-00-00) and check for corrosion and damage. Replace cable if necessary. Dip the cable in corrosion preventive compound (4, Table 2, 27-00-00). Excess should be removed by wiping with a clean rag. NOTE: More than one person may be required to route the aileron wing cable. Take precautions to keep the cable clean and free from damage. (2) Attach a tag labeled “left-hand threads terminal end” to the aileron outboard wing cable left-hand threads terminal end (1 or 4) (Ref. Figure 201). (3) Attach a tag labeled “right-hand threads terminal end” to the aileron outboard wing cable right-hand threads terminal end (5 or 10). (4) Attach the right-hand threads terminal end to the feed line labeled “right-hand threads terminal end”. (5) Attach the left-hand threads terminal end to the feed line labeled “left-hand threads terminal end”. NOTE: More than one person may be required to route the aileron wing cable. Take precautions to keep the cable clean and free from damage.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (6) Route the cables through the wing as follows: NOTE: It is permissible to install cable guard pins, fairleads and pulleys as the cable is being routed. (a) From the aileron bellcrank route cable (5 or 10) right-hand thread terminal end, through the forward pulley at WS 211 (Ref. Figure 201). NOTE: The cable guard pins for the pulleys at WS 182 are installed head up and cotter pinned. (b) Route cable (5 or 10) through the pulley at WS 182. (c) Route cable (5 or 10) through the pulley at WS 144. (d) Route cable (5 or 10) through the wing and wheel well to the aft turnbuckle at WS 103. (e) From the aileron bellcrank route cable (1 or 4) left-hand thread terminal end, through the aft pulley at WS 211. (f) Route cable (1 or 4) through the pulley at WS 147. (g) Route cable (1 or 4) through the wing and wheel well to the forward turnbuckle at WS 103. (h) Ensure all cable guard pins, pulleys and fairleads have been installed. (7) Install the aileron outboard cables on the bellcrank as follows (a) Place the right-hand thread cable (5) clevis end over the outboard wing bellcrank forward arm. Install bolt (22), washer (21), nut (20) and cotter pin (19) (Ref. Figure 201). (b) Insert the left-hand thread cable (4) clevis end into the outboard wing bellcrank aft arm. Install bolt (15), washer (16), nut (17) and cotter pin (18). (8) Lubricate turnbuckles with grease (1, Table 2, 27-00-00) for corrosion protection prior to installation. (9) Remove feed line, and attach the left-hand threads terminal end (3) to the forward turnbuckle (2) (Ref. Figure 204). (10) Remove feed line, and attach the right-hand threads terminal end (5) to the aft turnbuckle (8). (11) Tighten the turnbuckles (2 and 8) to tension the aileron wing cables. (12) Remove cable block (11) from the aileron wing inboard cables (1 and 10). (13) Ensure that the aileron wing outboard cables are routed properly by verifying that the cable has been routed exactly as described in Step (6). Ensure cables are engaged in the pulley grooves and all guard pins are installed. (14) Remove all tape from the cables and turnbuckles. (15) Perform the AILERON OPERATIONAL CHECK procedure (Ref. 27-10-03). (16) Perform the AILERON CONTROL SYSTEM RIGGING procedure (Ref. 27-10-03).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (17) Ensure safety clips are installed on all turnbuckles. (18) Install lower wing access panels 7, 35, 36, 37, 38, 43, 44, 45, 46 and 47 ((Ref. Chapter 6-50-00, WING ACCESS PANELS). (19) Install upper wing access panels 4, 29 and 30 (Ref. Chapter 6-50-00, WING ACCESS PANELS). (20) Remove the red tag from the control wheels and the flap lever.

C. Removal (UC-1 and After) (1) Connect external electrical power to the airplane. (2) Select the BATT switch to the ON position. (3) Select the EXT PWR switch to the EXT PWR position. (4) Lower flaps to the full down position. (5) Select the EXT PWR switch to the OFF position. (6) Select the BATT switch to the OFF position. (7) Disconnect external electrical power from the airplane. (8) Attach a red tag to the control wheels and the flap lever with the words “Do Not Operate, Maintenance In Progress”. (9) Perform the AILERON REMOVAL procedure (Ref. 27-10-00). (10) Remove the four cove panels and the two bellcrank covers from the wings at the inboard end of the ailerons. (11) Remove the lower wing access panels 17 and 19 as required (Ref. Chapter 6-50-00, WING ACCESS PANELS). NOTE: The aileron wing cables from the aileron quadrant to the turnbuckles and from the turnbuckles to the outboard bellcranks are not interchangeable. It is possible to wrap the cable around the wing bellcrank in either direction, but there is only ONE direction that will yield correct aileron movement. Each cable should be identified for correct location prior to removal. If aileron cable kit 118-5001-1 is installed the turnbuckles will be accessed though wing access panels 19 inboard of the main landing gear, instead of in the wheel well. Each turnbuckle (2 and 8) barrel has a groove (9) at one end to identify the left-hand threaded end (Ref. Figure 204). (12) Install cable block(s) (11) on the aileron wing inboard cables to prevent loss of cable tension. (13) Remove safety clips (7) from both turnbuckles (2 and 8). (14) Attach a tag with the words “aileron wing outboard cable right-hand (left-hand for kit 118-5001-1) threads terminal end” to the outboard end of the aft turnbuckle (8).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (15) Disconnect the right-hand (left-hand for kit 118-5001-1) threads terminal end (5) from the aft turnbuckle (8) and attach a feed line to the terminal end. Label the feed line with the words “right-hand (left-hand for kit 118-5001-1) threads terminal end”. (16) Attach a tag with the words “aileron wing outboard cable left-hand (right-hand for kit 118-5001-1) threads terminal end” to the outboard end of the forward turnbuckle (2). (17) Disconnect the left-hand (right-hand for kit 118-5001-1) threads terminal end (3) from the forward turnbuckle (2) and attach a feed line to the terminal end. Label the feed line with the words “left-hand (right-hand for kit 118-5001-1) threads terminal end”. NOTE: The cable guard pins for the pulleys at WS 150 are installed head down and cotter pinned. (18) Remove the cable guard pins from pulley brackets. Refer to Figure 202 for general pulley locations. (19) Cut safety wire (8) (Ref. Figure 205). (20) Remove bolt and washer (6) and remove the bellcrank (5). NOTE: If cable is to be reused, take care to keep the cable clean and free from damage. Coil cable loosely no tighter than a 29 inch diameter. (21) Pull the cables through the access openings at the outboard wing bellcrank. (22) Detach the feed lines from the aileron wing outboard cable terminal ends leaving the feed lines in place. (23) Prior to removal of the cable from the wing bellcrank, identify the cable position by marking the cable and bellcrank where the cable meets the top of the bellcrank. A visible line of ink, paint, etc. will aid in installation. (24) Cut the safety wire (2) and drive out the two roll pins (1) from the bellcrank (5). Remove the cable ball (3) and cable from the bellcrank (5).

D. Installation (UC-1 and After) WARNING: Perform all Steps of this procedure, in the order listed. Do not skip any Steps of this procedure. Failure to do so may result in injury to personnel and damage to equipment. The aileron wing outboard cable ball (3) is in the middle of the cable. It is possible to incorrectly wrap the cable on the aileron wing bellcrank (5). The right-hand thread terminal end (11) wraps over the top of the bellcrank (5) and the left-hand thread terminal end (12) wraps under the bottom of the bellcrank (5) (Ref. Figure 205). If Kit 118-5001-1 is installed the following warning applies: The aileron wing outboard cable ball (3) “IS NOT” in the middle of the cable. It is possible to incorrectly wrap the cable on the aileron wing bellcrank (5) but there is only ONE direction that will yield correct aileron movement. The short end (left-hand thread terminal end (11)) wraps over the top of the bellcrank (5) and the long end (right-hand thread terminal end (12)) wraps under the bottom of the bellcrank (5) (Ref. Figure 205).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL CAUTION: Where the cables pass through structure, the areas of possible contact between the control cables and adjacent structure must be protected with protective elements such as grommets, rub strips, blocks or guide fairings. Where contact of control cables does occur with the protective elements, a force no greater than eight (8) ounces shall be required to move the cable to a position of no contact. At no time should flight control cables contact metal structure with the protective elements removed. (1) If a used cable is to be installed, clean with solvent (2, Table 2, 27-00-00) and check for corrosion and damage. Replace cable if necessary. Dip the cable in corrosion preventive compound (4, Table 2, 27-00-00). Excess should be removed by wiping with a clean rag. (2) Install the aileron outboard cable on the bellcrank as follows: (a) Locate the right-hand (left-hand for kit 118-5001-1) threads terminal end (11) and wrap it over the top of the bellcrank (5) (Ref. Figure 205). (b) Place the cable ball (3) in the cable ball slot (10) in the bellcrank (5). (c) Install two roll pins (1) into the holes in the bellcrank (5). (d) Install 0.032-inch diameter safety wire (2). (3) Install the bellcrank (5) with bolt and washer (6). (4) Torque Bolt (6) to 50 to 70 inch pounds and safety wire with 0.032 safety wire (8). (5) Attach a tag labeled “left-hand (right-hand for kit 118-5001-1) threads terminal end” to the aileron outboard wing cable left-hand (right-hand for kit 118-5001-1) threads terminal end (12). (6) Attach a tag labeled “right-hand (left-hand for kit 118-5001-1) threads terminal end” to the aileron outboard wing cable right-hand (left-hand for kit 118-5001-1) threads terminal end (11). (7) Attach the right-hand threads terminal end to the feed line labeled “right-hand threads terminal end”. (8) Attach the left-hand threads terminal end to the feed line labeled “left-hand threads terminal end”. NOTE: More than one person may be required to route the aileron wing cable. Take precautions to keep the cable clean and free from damage. (9) Route the cables through the wing as follows: NOTE: It is permissible to install cable guard pins as the cable is being routed. (a) From the aileron bellcrank route cable (5 or 10) left-hand (right-hand for kit 118-5001-1) thread terminal end, through the pulley at WS 193 (Ref. Figure 202). NOTE: The cable guard pins for the pulleys at WS 150 are installed head down and cotter pinned. (b) Route cable (5 or 10) through the bottom pulley at WS 150. (c) Route cable (5 or 10) through the bottom pulley at WS 127.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (d) Route cable (5 or 10) through the wing into the wheel well to the forward turnbuckle at WS 98. (e) From the aileron bellcrank route cable (1 or 4) right-hand (left-hand for kit 118-5001-1) thread terminal end, through the top pulley at WS 150. (f) Route cable (1 or 4) through the top pulley at WS 127. (g) Route cable (1 or 4) through the wing into the wheel well to the aft turnbuckle at WS 98. (h) Ensure all cable guard pins are installed in the pulley brackets. (10) Lubricate turnbuckles with grease (1, Table 2, 27-00-00) for corrosion protection prior to installation. (11) Remove feed line, and attach the right-hand (left-hand for kit 118-5001-1) threads terminal end (5) to the aft turnbuckle (8) (Ref. Figure 204). (12) Remove feed line, and attach the left-hand (right-hand for kit 118-5001-1) threads terminal end (3) to the forward turnbuckle (2). (13) Tension the aileron wing outboard cable sufficient to prevent slack. (14) Remove the cable block(s) (11) from the aileron wing inboard cables. (15) Ensure that the aileron wing outboard cables are routed properly by verifying that the cable has been routed exactly as described in Step (9). Ensure cables are engaged in the pulley grooves and all guard pins are installed. (16) Remove all tape from the cables and turnbuckles. (17) Visually check to ensure that aileron wing bellcrank travel responds properly to the control wheel movement by performing the following: (a) Move pilot’s control wheel counterclockwise to the left position and make sure that the left and right wing aileron bellcranks rotate clockwise smoothly with no unusual noise or binding. (b) Move pilot’s control wheel clockwise to the right position and make sure that the left and right wing aileron bellcranks rotate counterclockwise smoothly with no unusual noise or binding. (18) Install the two aileron bellcrank covers and the four cove panels on the wing trailing edges. (19) Perform the AILERON INSTALLATION procedure (Ref. 27-10-00). (20) Perform the AILERON CONTROL SYSTEM RIGGING procedure (Ref. 27-10-03). (21) Ensure safety clips are installed on all turnbuckles. (22) Install lower wing access panels 17 and 19 (Ref. Chapter 6-50-00, WING ACCESS PANELS). (23) Remove the red tag from the control wheels and the flap lever.

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3. WING INBOARD CABLE A. Removal (UA-1 and After; UB-1 and After) (1) Attach a red tag to the control wheels and the flap lever with the words “Do Not Operate, Maintenance In Progress”. (2) Remove passenger seats as required to access floor access panels 16E, 10 (UA-1 and After only), 11 (UB-1 and After; UC-1 and After) and 17E (Ref. Chapter 25-20-00, SEAT REMOVAL). (3) Remove passenger compartment carpet as required to access floor access panels 16E, 10 (UA-1 and After only), 11 (UB-1 and After; UC-1 and After) and 17E (Ref. Chapter 25-20-01). (4) Remove the center aisle spar aft ramp. (5) Remove floor access panels 16E, 10 (UA-1 and After only), 11 (UB-1 and After; UC-1 and After) and 17E (Ref. Chapter 6-50-00, FLOOR ACCESS PANELS). (6) Remove lower wing access panels 7, 8, and 9 (Ref. Chapter 6-50-00, WING ACCESS PANELS). NOTE: The aileron cables from the aileron quadrant to the turnbuckles and from the turnbuckles to the outboard bellcranks are not interchangeable. Each cable should be identified for correct location prior to removal. Each turnbuckle (2 and 8) barrel has a groove (9) at one end to identify the left-hand threaded end (Ref. Figure 204). (7) Move aileron wing cables to ensure easy access to the turnbuckles. (8) Install cable block (4) on the wing outboard cables (3 and 5) to prevent loss of cable tension. (9) Remove safety clips (7) from both turnbuckles (2 and 8). (10) Attach a tag with the words “inboard cable left-hand threads terminal end” to the inboard end of the aft turnbuckle (8). (11) Disconnect the left-hand threads terminal end (10) from the aft turnbuckle (8) and attach a feed line to the terminal end. Label the feed line with the words “left-hand threads terminal end”. (12) Attach a tag with the words “inboard cable right-hand threads terminal end” to the inboard end of the forward turnbuckle (2). (13) Disconnect the right-hand threads terminal end (1) from the forward turnbuckle (2) and attach a feed line to the terminal end. Label the feed line with the words “right-hand threads terminal end”. (14) Disconnect the aileron cable locking plates (37) on the aileron quadrant just forward of the rear spar (Ref. Figure 201). (15) Remove the cable guard pins from pulley brackets and pressure seals from the fuselage. NOTE: It may be necessary to remove the pulleys if cable passage is restricted. If cable is to be reused, take care to keep the cable clean and free from damage. Coil cable loosely no tighter than a 29 inch diameter. (16) With assistance pull the cables through the access opening at the aileron quadrant in the fuselage.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (17) Detach the feed lines from the aileron wing inboard cable terminal ends leaving the feed lines in place.

B. Installation (UA-1 and After; UB-1 and After) WARNING: Perform all Steps of this procedure, in the order listed. Do not skip any Steps of this procedure. Failure to do so may result in injury to personnel and damage to equipment. CAUTION: Where the cables pass through structure, the areas of possible contact between the control cables and adjacent structure must be protected with protective elements such as grommets, rub strips, blocks or guide fairings. Where contact of control cables does occur with the protective elements, a force no greater than eight (8) ounces shall be required to move the cable to a position of no contact. At no time should flight control cables contact metal structure with the protective elements removed. (1) If a used cable is to be installed, clean with solvent (2, Table 2, 27-00-00) and check for corrosion and damage. Replace cable if necessary. Dip the cable in corrosion preventive compound (4, Table 2, 27-00-00). Excess should be removed by wiping with a clean rag. NOTE: More than one person may be required to route the aileron wing cable. Take precautions to keep the cable clean and free from damage. (2) Attach a tag labeled “left-hand threads terminal end” to the inboard cable left-hand threads terminal end. (3) Attach a tag labeled “right-hand threads terminal end” to the inboard cable right-hand threads terminal end. (4) Attach the right-hand threads terminal end to the feed line labeled “right-hand threads terminal end”. (5) Attach the left-hand threads terminal end to the feed line labeled “left-hand threads terminal end”. NOTE: More than one person will be required to route the aileron wing cable. Take precautions to keep the cable clean and free from damage. It is permissible to install cable guard pins and pulleys as the cable is being routed. (6) Route the cables through the wing and fuselage as follows: (a) From the aileron quadrant route cable (6 or 9) right-hand thread, through the forward pulley at BL 26 (Ref. Figure 201). NOTE: When routing the cable through the forward pressure seal hole ensure the cable passes through the pressure seal retainer. (b) Route cable (6 or 9) through the forward pressure seal hole at BL 27. (c) Route cable (6 or 9) through the wing to the forward turnbuckle at WS 103. (d) From the aileron quadrant route cable (2 or 3) left-hand thread, through the aft pulley at BL 26.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL NOTE: When routing the cable through the pressure seal hole ensure the cable passes through the pressure seal retainer. (e) Route cable (2 or 3) through the aft pressure seal hole at BL 27. (f) Route cable (2 or 3) through the wing to the aft turnbuckle at WS 103. CAUTION: Do not over torque the cable locking plate attachment screws or damage to the quadrant may occur. Maximum torque will not exceed 15 inch-pounds. (g) Connect the cables to the aileron quadrant by installing the cable locking plates (37) and attaching screws (39). Safety wire the screws with 0.032-inch diameter safety wire. (h) Ensure all cable guard pins and pulleys have been installed. (7) Lubricate turnbuckles with grease (1, Table 2, 27-00-00) for corrosion protection prior to installation. (8) Remove feed line, and attach the left-hand threads terminal end (10) to the aft turnbuckle (8) (Ref. Figure 204). (9) Remove feed line, and attach the right-hand threads terminal end (1) to the forward turnbuckle (2). (10) Tension the aileron wing inboard cable sufficient to prevent slack. (11) Remove cable block (4) from the wing outboard cables (3 and 5). (12) Ensure that the aileron wing inboard cables are routed properly by verifying that the cables have been routed exactly as described in Step (6). Ensure the cables are engaged in the pulley grooves and all guard pins are installed. (13) Remove all tape from the cables and turnbuckles. (14)

Lubricate the aileron wing inboard cable to one inch beyond the length of travel through the pressure seal with grease (1, Table 2, 27-00-00).

(15) Fill the cable pressure seal with grease (1, Table 2, 27-00-00), and install the seal in the hole at BL 27 and seal with sealant (12, Table 2, 27-00-00). (16) Perform the AILERON OPERATIONAL CHECK procedure (Ref. 27-10-03). (17) Perform the AILERON CONTROL SYSTEM RIGGING procedure (Ref. 27-10-03). (18) Ensure safety clips are installed on all turnbuckles. (19) Install floor access panels 16E, 10 (UA-1 and After only), 11 (UB-1 and After; UC-1 and After) and 17E (Ref. Chapter 6-50-00, FLOOR ACCESS PANELS). (20) Install the center aisle spar aft ramp. (21) Install passenger compartment carpet (Ref. Chapter 25-20-01). (22) Install passenger seats (Ref. Chapter 25-20-00, SEAT INSTALLATION). (23) Install lower wing access panels 7, 8, and 9 (Ref. Chapter 6-50-00, WING ACCESS PANELS).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (24) Remove the red tag from the control wheels and the flap lever.

C. Removal (UC-1 and After) (1) Attach a red tag to the control wheels with the words “Do Not Operate, Maintenance In Progress”. (2) Remove passenger seats as required to access floor access panels 16E, 11 and 17E (Ref. Chapter 25-20-00, SEAT REMOVAL). (3) Remove passenger compartment carpet as required to access floor access panels 16E, 11 and 17E (Ref. Chapter 25-20-01). (4) Remove the spar aft ramp. (5) Remove floor access panels 16E, 11 and 17E (Ref. Chapter 6-50-00, FLOOR ACCESS PANELS). (6) Remove the lower wing access panels 17 and 19 as required (Ref. Chapter 6-50-00, WING ACCESS PANELS). NOTE: The aileron cables from the aileron quadrant to the turnbuckles and from the turnbuckles to the outboard bellcranks are not interchangeable. Each cable should be identified for correct location prior to removal. If aileron cable Kit 118-5001-1 is installed the turnbuckles will be accessed though wing access panels 19 inboard of the main landing gear, instead of in the wheel well. Each turnbuckle (2 and 8) barrel has a groove (9) at one end to identify the left-hand threaded end (Ref. Figure 204). (7) Install cable block(s) (4) on the wing outboard cables to prevent loss of cable tension. (8) Remove safety clips (7) from both turnbuckles (2 and 8). (9) Attach a tag with the words “right-hand (left-hand for Kit 118-5001-1) threads terminal end” to the inboard end of the forward turnbuckle (2). (10) Disconnect the right-hand (left-hand for Kit 118-5001-1) threads terminal end (1) from the forward turnbuckle (2) and attach a feed line to the terminal end. Label the feed line with the words “right-hand (left-hand for kit 118-5001-1) threads terminal end”. (11) Attach a tag with the words “left-hand (right-hand for Kit 118-5001-1) threads terminal end” to the inboard end of the aft turnbuckle (8). (12) Disconnect the left-hand (right-hand for Kit 118-5001-1) threads terminal end from the aft turnbuckle (8) and attach a feed line to the terminal end. Label the feed line with the words “left-hand (left-hand for kit 118-5001-1) threads terminal end”. (13) Disconnect the aileron cable locking plates (34) on the aileron quadrant just forward of the rear spar (Ref. Figure 202, Detail D). (14) Remove the cable guard pins from pulley brackets and pressure seals from the fuselage. Refer to Figure 202, UC-1 and After for general locations. NOTE: It may be necessary to remove the pulleys if cable passage is restricted.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL If cable is to be reused, take care to keep the cable clean and free from damage. Coil cable loosely no tighter than a 29 inch diameter. (15) With assistance pull the cables through the access opening at the aileron quadrant in the fuselage. (16) Detach the feed lines from the aileron wing inboard cable terminal ends leaving the feed lines in place.

D. Installation (UC-1 and After) WARNING: Perform all Steps of this procedure, in the order listed. Do not skip any Steps of this procedure. Failure to do so may result in injury to personnel and damage to equipment. CAUTION: Where the cables pass through structure, the areas of possible contact between the control cables and adjacent structure must be protected with protective elements such as grommets, rub strips, blocks or guide fairings. Where contact of control cables does occur with the protective elements, a force no greater than eight (8) ounces shall be required to move the cable to a position of no contact. At no time should flight control cables contact metal structure with the protective elements removed. (1) If a used cable is to be installed, clean with solvent (2, Table 2, 27-00-00) and check for corrosion and damage. Replace cable if necessary. Dip the cable in corrosion preventive compound (4, Table 2, 27-00-00). Excess should be removed by wiping with a clean rag. (2) Attach a tag labeled “left-hand threads terminal end” to the aileron wing inboard cable left-hand threads terminal end. (3) Attach a tag labeled “right-hand threads terminal end” to the aileron wing inboard cable right-hand threads terminal end. (4) Attach the right-hand threads terminal end to the feed line labeled “right-hand threads terminal end”. (5) Attach the left-hand threads terminal end to the feed line labeled “left-hand threads terminal end”. NOTE: More than one person will be required to route the aileron wing cable. Take precautions to keep the cable clean and free from damage. It is permissible to install cable guard pins as the cable is being routed. (6) Route the cables through the wing and fuselage as follows: NOTE: It is permissible to install cable guard pins and any removed pulleys as the cable is being routed. (a) From the aileron quadrant route cable (6 or 9) right-hand (left-hand for Kit 118-5001-1) threads terminal end, through the forward pulley at BL 26 (Ref. Figure 202). NOTE: When routing the cable through the pressure seal hole ensure the cable passes through the pressure seal retainer (13). (b) Route cable (6 or 9) through the forward pressure seal hole at BL 27. (c) Route cable (6 or 9) through the wing to the forward turnbuckle at WS 98.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (d) From the aileron quadrant route cable (2 or 3) left-hand (right-hand for Kit 118-5001-1) threads terminal end, through the aft pulley at BL 26. NOTE: When routing the cable through the pressure seal hole ensure the cable passes through the pressure seal retainer (13). (e) Route cable (2 or 3) through the aft pressure seal hole at BL 27. (f) Route cable (2 or 3) through the wing to the aft turnbuckle at WS 98. CAUTION: Do not over torque the cable locking plate attachment screws or damage to the quadrant may occur. Maximum torque will not exceed 15 inch-pounds. (g) Connect the cables to the aileron quadrant by installing the cable locking plates (34) and attaching screws (36). Safety wire the screws with 0.032-inch diameter safety wire (Ref. Figure 202, Details A, D and G). (h) Ensure all cable guard pins and pulleys are installed in the pulley brackets. NOTE: The interior of all turnbuckles should be coated or filled with grease (1, Table 2, 27-00-00) for corrosion protection. (7) Lubricate turnbuckles with grease (1, Table 2, 27-00-00) for corrosion protection prior to installation. (8) Remove feed line, and attach the right-hand (left-hand for Kit 118-5001-1) threads terminal end (1) to the forward turnbuckle (2) (Ref. Figure 204). (9) Remove feed line, and attach the left-hand (right-hand for Kit 118-5001-1) threads terminal end (10) to the aft turnbuckle (8). (10) Tension the aileron wing inboard cable sufficient to prevent slack. (11) Remove cable block(s) (4) from the wing outboard cables. (12) Ensure that the aileron wing inboard cables are routed properly by verifying that the cables have been routed exactly as described in Step (6). Ensure the cables are engaged in the pulley grooves and all guard pins are installed. (13) Remove all tape from the cables and turnbuckles. (14)

Lubricate the aileron wing inboard cable to one inch beyond the length of travel through the pressure seal with grease (1, Table 2, 27-00-00).

(15) Fill the cable pressure seal with grease (1, Table 2, 27-00-00), and install the seal in the hole at BL 27 and seal with sealant (12, Table 2, 27-00-00). (16) Perform the AILERON OPERATIONAL CHECK procedure (Ref. 27-10-03). (17) Perform the AILERON CONTROL SYSTEM RIGGING procedure (Ref. 27-10-03). (18) Ensure all turnbuckles have safety clips installed. (19) Install cabin floorboard panels 16E, 11 and 17E (Ref. Chapter 6-50-00, FLOOR ACCESS PANELS).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (20) Instal the center aisle spar aft ramp. (21) Install passenger compartment carpet (Ref. Chapter 25-20-01). (22) Install passenger seats (Ref. Chapter 25-20-00, SEAT INSTALLATION). (23) Install lower wing access panels 17 and 19 (Ref. Chapter 6-50-00, WING ACCESS PANELS). (24) Remove the red tag from the control wheels.

4. OUTBOARD WING BELLCRANK A. Removal (UA-1 and After; UB-1 and After) (1) Remove lower wing access panels 43, 44 and 46 (left) or 38, 45, and 47 (right) (Ref. Chapter 6-50-00, WING ACCESS PANELS). (2) Remove upper wing access panel 29 (left) or 30 (right) (Ref. Chapter 6-50-00, WING ACCESS PANELS). (3) Install cable blocks to the aileron cables through panel 43 or 38. (4) Remove cotter pin (5), nut (4), washer (3) and bolt (1) from the aft arm of the outboard wing bellcrank (18) (Ref. Figure 208). (5) Remove cotter pin (12), nut (13), washer (14) and bolt (15) from the forward arm of the outboard wing bellcrank (18). (6) Move the outboard wing bellcrank (18) to access the aileron push-pull rod (6) and remove cotter pin (8), nut (10), washer (11) and bolt (2). (7) Remove safety wire from the outboard wing bellcrank mount bolt (9) and remove bolt (9). (8) Remove the outboard wing bellcrank (18) from the airplane.

B. Installation (UA-1 and After; UB-1 and After) (1) Position the outboard wing bellcrank (18) and install bolt (9). Safety wire the bolt (9) (Ref. Figure 208). (2) Position the aileron push-pull rod (6) to the outboard wing bellcrank (18) and install bolt (2), washer (11), nut (10) and cotter pin (8). (3) Position the aft aileron cable (17) clevis end to the aft arm of the outboard wing bellcrank (18) and install bolt (1), washer (3), nut (4) and cotter pin (5). (4) Position the forward aileron cable (16) clevis end to the forward arm of the outboard wing bellcrank (18) and install bolt (15), washer (14), nut (13) and cotter pin (12). (5) Remove cable blocks from the aileron cables through panel 43 or 38. (6) Perform the AILERON CONTROL SYSTEM RIGGING procedure (Ref. 27-10-03). (7) Install upper wing access panel 29 (left) or 30 (right) (Ref. Chapter 6-50-00, WING ACCESS PANELS).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (8) Install lower wing access panels 43, 44 and 46 (left) or 38, 45, and 47 (right) (Ref. Chapter 6-50-00, WING ACCESS PANELS).

5. BELLCRANK ASSEMBLY NOTE: The AILERON YOKE ASSEMBLY and the AILERON BELLCRANK ASSEMBLY may be checked at the same interval to satisfy the inspection requirements outlined in Chapter 5.

A. Removal and Inspection (UC-1 and After) (1) Connect external electrical power to the airplane. (2) Select the BATT switch to the ON position. (3) Select the EXT PWR switch to the EXT PWR position. (4) Lower flaps to the full down position. (5) Select the EXT PWR switch to the OFF position. (6) Select the BATT switch to the OFF position. (7) Disconnect external electrical power from the airplane. (8) Attach a red tag to the control wheels and the flap lever with the words “Do Not Operate, Maintenance In Progress”. (9) Perform the AILERON REMOVAL procedure (Ref. 27-10-00). NOTE: If aileron cable Kit 118-5001-1 is installed the turnbuckles will be accessed through wing access panels 19 inboard of the main landing gear, instead of in the wheel well. (10) Remove the four cove panels and the two bellcrank covers from the wings at the inboard end of the ailerons as required. (11) Remove the lower wing access panels 17 and 19 as required (Ref. Chapter 6-50-00, WING ACCESS PANELS). (12) Install cable block(s) (11) on the aileron wing inboard cables to prevent loss of cable tension (Ref. Figure 204). (13) Remove safety clips (7) from turnbuckles (2 and 8). (14) Loosen turnbuckles (2 and 8) to relieve outboard cable tension. (15) Cut safety wire (8) and remove bolt and washer (6) (Ref. Figure 205). NOTE: The bellcrank will still be attached to the Aileron wing outboard cable. (16) Remove bellcrank (5) from aft wing area. (17) Check bearing (9) in the bellcrank arm for shifting, corrosion and smoothness of rotation and ability to pivot. (18) Check the bearings (13) in the bellcrank (5) for shifting, corrosion and smoothness of rotation.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (19) If replacement or repairs are required, perform the following Steps: (a) Cut safety wire (2) and drive out roll pins (1). (b) Remove bellcrank (5) from the outboard cable. (20) Inspect pulleys, brackets and hardware for wear. NOTE: Bearings should be staked if replaced. Inspect bearings for looseness and smoothness of rotation after installation. (21) Replace or repair damaged bearings, and/or pulleys, brackets and hardware as required.

B. Installation (UC-1 and After) WARNING: Perform all Steps of this procedure, in the order listed. Do not skip any Steps of this procedure. Failure to do so may result in injury to personnel and damage to equipment. The aileron wing outboard cable ball (3) is in the middle of the cable. It is possible to incorrectly wrap the cable on the aileron wing bellcrank (5). The right-hand thread terminal end (11) wraps over the top of the bellcrank (5) and the left-hand thread terminal end (12) wraps under the bottom of the bellcrank (5) (Ref. Figure 205). If Kit 118-5001-1 is installed the following warning applies: The aileron wing outboard cable ball (3) “IS NOT” in the middle of the cable. It is possible to incorrectly wrap the cable on the aileron wing bellcrank (5) but there is only ONE direction that will yield correct aileron movement. The short end (left-hand thread terminal end (11)) wraps over the top of the bellcrank (5) and the long end (right-hand thread terminal end (12)) wraps under the bottom of the bellcrank (5) (Ref. Figure 205). (1) Install the aileron outboard cable on the bellcrank as follows: (a) Locate the right-hand (left-hand for kit 118-5001-1) threads terminal end (11) and wrap it over the top of the bellcrank (5) (Ref. Figure 205). (b) Place the cable ball (3) in the cable ball slot (10) in the bellcrank (5). (c) Install two roll pins (1) into the holes in the bellcrank (5). (d) Install 0.032-inch diameter safety wire (2). (2) Install the bellcrank (5) with bolt and washer (6). (3) Torque Bolt (6) from 50 to 70 inch-pounds and safety wire with 0.032 safety wire (8). (4) Tension the aileron wing cables sufficient to prevent slack. (5) Visually check to ensure the cables are properly seated in all pulleys. (6) Remove the cable blocks (11) from the aileron wing inboard cables (Ref. Figure 204). (7) Visually check to ensure that aileron wing bellcrank travel responds properly to the control wheel movement by performing the following:

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (a) Move pilot’s control wheel counterclockwise to the left position and make sure that the left and right wing aileron bellcranks rotate clockwise smoothly with no unusual noise or binding. (b) Move pilot’s control wheel clockwise to the right position and make sure that the left and right wing aileron bellcranks rotate counterclockwise smoothly with no unusual noise or binding. (8) Install the four cove panels and the two bellcrank covers from the wings at the inboard end of the ailerons. (9) Perform the AILERON INSTALLATION procedure (Ref. 27-10-00). (10) Perform the AILERON CONTROL SYSTEM RIGGING procedure (Ref. 27-10-03). (11) Ensure safety clips (7) are installed on all turnbuckles (2 and 8) (Ref. Figure 204). (12) Install the lower wing access panels 17 and 19 (Ref. Chapter 6-50-00, WING ACCESS PANELS).

6. YOKE ASSEMBLY NOTE: The AILERON YOKE ASSEMBLY and the AILERON BELLCRANK ASSEMBLY may be inspected at the same interval to satisfy the inspection requirements outlined in Chapter 5.

A. Checks (UC-1 and After) (1) Perform the AILERON REMOVAL procedure (Ref. 27-10-00). (2) Move the yoke assembly side to side to check bearings (4 and 7) for corrosion, looseness and smoothness of rotation. If bearings need to be replaced, perform AILERON-YOKE BEARING REMOVAL and AILERON-YOKE BEARING INSTALLATION procedures in this section (Ref. Figure 206 and 207). (3) Measure the diameter of the taper pin (12) in the area that engages the bellcrank bearing. Measure the inside diameter of the bearing. The inside diameter of the bearing minus the diameter of the taper pin should not exceed 0.005 inch. If the difference exceeds 0.005 inch, it must be determined which part(s) needs to be replaced by comparing the existing part dimensions to new part dimensions. A new pin measures 0.3742 / 0.3737 inch diameter in the area that engages into the bellcrank bearing. A new bearing inside diameter measures 0.3750 / 0.3745 inch. It is permissible to allow these parts to wear beyond the new part tolerances; however, the 0.005 inch difference between the two diameters should not be exceeded. (4) If taper pin needs to be replaced, perform AILERON-YOKE TAPER PIN REMOVAL and AILERON-YOKE TAPER PIN INSTALLATION procedure in this section.

B. Bearing Removal (UC-1 and After) (1) Perform AILERON YOKE ASSEMBLY REMOVAL in this section. NOTE: Bearings (4 and 7) may need to be pressed out of the yoke (6) (Ref. Figure 206 and 207). (2) Remove the two snap rings (3). (3) Remove bearings (4 and 7) and spacer (5). Discard bearings as required.

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C. Bearing Installation (UC-1 and After) (1) Clean the interior surfaces of the yoke (6) with solvent (8, Table 2, 27-00-00) (Ref. Figure 206 and 207). (2) Ensure that bearing seats are 0.9007 +0.0000/-0.0005 inch in diameter. NOTE: Spacer (5) must be inserted between the bearings (4 and 7) before they are seated. (3) Install bearing (7) in the aileron yoke assembly with retaining compound (5, Table 2, 27-00-00). (4) Install the two snap rings (3). (5) Insert spacer (5) before installing bearing (4). (6) Install bearing (4) with retaining compound (5, Table 2, 27-00-00). NOTE: After installation inspect bearings (4 and 7) for looseness and smoothness of rotation. (7) Perform the AILERON YOKE ASSEMBLY INSTALLATION procedure in this section.

D. Taper Pin Removal (UC-1 and After) (1) Perform AILERON YOKE ASSEMBLY REMOVAL in this section. (2) Remove cotter pin (15), nut (14) and washer (13) (Ref. Figure 206 and 207). (3) Remove taper pin (12) and discard old pin, if required.

E. Taper Pin Installation (UC-1 and After) (1) Inspect the condition of the taper pin per Step (3) in the AILERON YOKE ASSEMBLY CHECKS procedure in this section. Replace pin if required. (2) Install the taper pin (12), taper pin washer (13) and nut (14). Torque nut (14) to 22 inch-pounds. If cotter pin (15) cannot be installed continue to tighten nut (14) until castellation and cotter pin hole align. Maximum torque will not exceed 45 inch-pounds. Install cotter pin (15) (Ref. Figure 206 and 207). (3) Ensure that the taper pin (12) protrudes 1.23 to 1.35 inches from the end of the yoke (6). (4) Perform AILERON YOKE ASSEMBLY INSTALLATION in this section.

F. Removal (UC-1 and After) (1) Perform the AILERON REMOVAL procedure (Ref. 27-10-00). (2) On the forward inboard end of the aileron, remove the plugs to obtain access to the aileron yoke assembly (Ref. Figure 206 and 207). (3) Remove cotter pin (11), nut (10), washer (9) and bolt (1) from the yoke assembly. (4) Remove yoke assembly from the aileron. (5) Note position of bushings prior to removal. Remove bushings (2) and (8) from the fitting assembly.

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G. Installation (UC-1 and After) NOTE: On airplanes UC-1 and After not modified by Service Bulletin 2489, install the lower bushing (8) with the flange on the outside of the fitting assembly (Ref. Figure 206). On airplanes UC-1 and After that have Service Bulletin 2489 modifications, install the lower bushing (8) with the flange on the inside of the fitting assembly (Ref. Figure 207). (1) Install bushing (2) in upper fitting and bushing (8) in lower fitting (Ref. Figure 206 and 207). NOTE: Ensure the yoke assembly is correctly inserted into the aileron with bearing (7) facing down. (2) Place the yoke assembly inside the aileron and insert the bolt (1) through the fitting into the yoke (6) and spacer (5), then secure with washer (9), nut (10) and cotter pin (11). (3) Install upper and lower plugs on the forward inboard end of the aileron. (4) Perform the AILERON INSTALLATION procedure (Ref. 27-10-00).

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11

E

12

1

1 WS 216

WS 147

10 WS 211

14 13

1

10

10 WS 182

B

4

C

10 R

WS 144

10

1

L

OUTBOARD BELLCRANK

2 DETAIL

L

E

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R

WS 103 BL 27

22

2 9 BL 26

D

AILERON SERVO ACTUATOR

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3 FS 322

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6 BL 27

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FS 127

DETAIL

5

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WS 147

4 5

WS 182

FS 145 1. AILERON RIGHT WING OUTBOARD LEFT-HAND THREAD CABLE 2. AILERON RIGHT WING INBOARD LEFT-HAND THREAD CABLE 3. AILERON LEFT WING INBOARD LEFT-HAND THREAD CABLE 4. AILERON LEFT WING OUTBOARD LEFT-HAND THREAD CABLE 5. AILERON LEFT WING OUTBOARD RIGHT-HAND THREAD CABLE 6. AILERON LEFT WING INBOARD RIGHT-HAND THREAD CABLE 7. AILERON FUSELAGE LEFT, LEFT-HAND THREAD CABLE 8. AILERON FUSELAGE RIGHT, RIGHT-HAND THREAD CABLE 9. AILERON RIGHT WING INBOARD RIGHT-HAND THREAD CABLE 10. AILERON RIGHT WING OUTBOARD RIGHT-HAND THREAD CABLE 11. UP STOP BOLT

F

5

7 7

FS 103

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L

FS 161

L

OUTBOARD BELLCRANK

C 6

8

8

14

21 20

8

FS 127

16 17 18

5

12. DOWN STOP BOLT 13. RIG PIN HOLE 14. PUSH-PULL ROD 15. BOLT 16. WASHER 17. NUT 18. COTTER PIN 19. COTTER PIN 20. NUT 21. WASHER 22. BOLT

4

F

5

WS 211

5 WS 216

E UA27B 050013AA.AI

Figure 201 (Sheet 1 of 2) Aileron Control System (UA-1 and After; UB-1 and After)

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23. BRACKET 24. SEAL 25. RETAINER 26. SPRING 27. AILERON LEFT WING FORWARD CABLE 28. AILERON LEFT HAND FUSELAGE CABLE 29. AILERON RIGHT HAND FUSELAGE CABLE 30. AILERON RIGHT WING FORWARD CABLE 31. AILERON RIGHT WING AFT CABLE 32. AILERON LEFT WING AFT CABLE 33. AILERON FUSELAGE CABLE GROOVE 34. LEFT WING AILERON CABLE GROOVE 35. RIGHT WING AILERON CABLE GROOVE 36. AILERON SERVO ACTUATOR CABLE GROOVE 37. CABLE LOCKING PLATE 38. SAFETY WIRE 39. SCREW

37

23

38 24

39

26

25

AILERON RIGHT WING FORWARD CABLE SAFETY WIRE DETAIL

PRESSURE SEALS

A

DETAIL

C

TURNBUCKLE SAFETYCLIPS

28

A

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DETAIL

29

B

30

27

37 38

39

F 33

FWD

32

34 35

31

AILERON QUADRANT (TOP VIEW) DETAIL

D

AILERON LEFT WING FORWARD CABLE SAFETY WIRE (TYPICAL)

36 DETAIL

G

AILERON QUADRANT (SIDE VIEW) DETAIL

F

UC27B 050014AA.AI

Figure 201 (Sheet 2 of 2) Aileron Control System (UA-1 and After; UB-1 and After

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E

1. AILERON RIGHT WING OUTBOARD RIGHT-HAND THREAD CABLE LEFT HAND THREAD WITH KIT 118-5001 INSTALLED 2. AILERON RIGHT WING INBOARD LEFT-HAND THREAD CABLE RIGHT HAND THREAD WITH KIT 118-5001 INSTALLED 3. AILERON LEFT WING INBOARD LEFT-HAND THREAD CABLE RIGHT HAND THREAD WITH KIT 118-5001 INSTALLED 4. AILERON LEFT WING OUTBOARD RIGHT-HAND THREAD CABLE LEFT HAND THREAD WITH KIT 118-5001 INSTALLED 5. AILERON LEFT WING OUTBOARD LEFT-HAND THREAD CABLE RIGHT HAND THREAD WITH KIT 118-5001 INSTALLED 6. AILERON LEFT WING INBOARD RIGHT-HAND THREAD CABLE LEFT HAND THREAD WITH KIT 118-5001 INSTALLED 7. AILERON FUSELAGE LEFT, LEFT-HAND THREAD CABLE 8. AILERON FUSELAGE RIGHT, RIGHT-HAND THREAD CABLE 9. AILERON RIGHT WING INBOARD RIGHT-HAND THREAD CABLE LEFT HAND THREAD WITH KIT 118-5001 INSTALLED 10. AILERON RIGHT WING OUTBOARD LEFT-HAND THREAD CABLE RIGHT HAND THREAD WITH KIT 118-5001 INSTALLED

1

WS 208

10

1 WS 193

10 WS 150

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R L

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AILERON SERVO ACTUATOR

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3 BL 26

BL 27

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7 FS 103

FS 127

L R

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4

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8 L

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3 FS 322

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FS 161 FS 145

THE AILERON CABLES RUN FROM THE BELLCRANK ON A VERTICAL PLANE AND ROTATE TO A HORIZONTAL PLANE BETWEEN WING STATIONS 104 AND 124, THEN CONTINUE ON A HORIZONTAL PLANE UNTIL CONNECTING TO THE AILERON QUADRANT.

WS 98

4 5

E

4 WS 127

5 WS 150

5 WS 193

5 WS 208 UC27B 050010AA.AI

Figure 202 (Sheet 1 of 2) Aileron Control System (UC-1 and After)

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24. BELLCRANK 25. TAPER PIN 26. ARM 27. YOKE 28. YOKE HINGE POINT 29. AILERON HINGE POINT 30. AILERON FUSELAGE CABLE GROOVE 31. LEFT WING AILERON CABLE GROOVE 32. RIGHT WING AILERON CABLE GROOVE 33. AILERON SERVO ACTUATOR CABLE GROOVE 34. CABLE LOCKING PLATE 35. SAFETY WIRE 36. SCREW

11. BRACKET 12. SEAL 13. RETAINER 14. SPRING 15. AILERON LEFT WING FORWARD CABLE 16. AILERON LEFT HAND FUSELAGE CABLE 17. AILERON RIGHT HAND FUSELAGE CABLE 18. AILERON RIGHT WING FORWARD CABLE 19. AILERON LEFT WING AFT CABLE 20. AILERON RIGHT WING AFT CABLE 21. BELLCRANK COVER 22. UPPER STOP BOLT 23. HINGE POINT

11 TURNBUCKLE SAFETYCLIPS DETAIL

25

B

12

14

13

26 27

24

PRESSURE SEALS 28 DETAIL

AILERON RIGHT WING FORWARD CABLE SAFETY WIRE DETAIL

C

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A 22

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15 21

OUTBOARD BELLCRANK (VIEW LOOKING DOWN) DETAIL

35

36

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E

F 30

FWD

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31 32

20

AILERON QUADRANT (TOP VIEW) DETAIL

D

AILERON LEFT WING FORWARD CABLE SAFETY WIRE (TYPICAL)

33 DETAIL

G

AILERON QUADRANT (SIDE VIEW) DETAIL

F

UC27B 050011AA.AI

Figure 202 (Sheet 2 of 2) Aileron Control System (UC-1 and After)

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1. "T" COLUMN 2. RIGHT-HAND THREAD TERMINAL END 3. LEFT-HAND THREAD TERMINAL END 4. RIGHT TURNBUCKLE 5. LEFT TURNBUCKLE 6. SAFETY CLIPS 7. LEFT-HAND THREAD GROOVES 8. "T" COLUMN CABLES

A

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8

7

6

5 7

4 3

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VIEW LOOKING RIGHT FWD DETAIL

A

UC27B 046216AA.AI

Figure 203 Aileron Fuselage Cable Turnbuckles

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1. WING INBOARD CABLE RIGHT-HAND THREAD TERMINAL END 2. FORWARD TURNBUCKLE 3. WING OUTBOARD CABLE LEFT-HAND THREAD TERMINAL END 4. CABLE BLOCK 5. WING OUTBOARD CABLE RIGHT-HAND THREAD TERMINAL END 6. LEFT MAIN LANDING GEAR 7. SAFETY CLIPS 8. AFT TURNBUCKLE 9. LEFT-HAND THREAD GROOVE 10. WING INBOARD CABLE LEFT-HAND THREAD TERMINAL END 11. CABLE BLOCK

A 2

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3 4

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11 10

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9 8

7 OUTBD

VIEW LOOKING AFT (LEFT SIDE SHOWN, RIGHT SIDE SIMILAR) DETAIL

A

Figure 204 Aileron Wing Cable Turnbuckles

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

1

2

A

3

FWD

OUT BD 1. ROLL PINS 2. SAFETY WIRE 3. CABLE BALL 4. STAKE POINTS 5. AILERON WING BELLCRANK 6. BOLT AND WASHER 7. BEARING 8. SAFETY WIRE 9. BEARING 10. CABLE BALL SLOT 11. RIGHT-HAND THREAD TERMINAL END LEFT HAND WITH KIT 118-5001 INSTALLED 12. LEFT-HAND THREAD TERMINAL END RIGHT HAND WITH KIT 118-5001 INSTALLED 13. BEARING

RIGHT SIDE SHOWN LEFT SIDE TYPICAL 4

5

6

7

4

9

8

OUTBD 5

10

11 3

12

DETAIL

13

A

UC27B 046217AA.AI

Figure 205 Aileron Wing Bellcrank (UC-1 and After)

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Figure 206 Aileron Yoke Assembly (UC-1 and After)

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Figure 207 Aileron Yoke Assembly (UC-1 and After Modified by Service Bulletin 2489)

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1. BOLT 2. BOLT 3. WASHER 4. NUT 5. COTTER PIN 6. AILERON PUSH-PULL ROD 7. JAMNUT 8. COTTER PIN 9. BOLT 10. NUT

11. WASHER 12. COTTER PIN 13. NUT 14. WASHER 15. BOLT 16. FORWARD AILERON CABLE 17. AFT AILERON CABLE 18. OUTBOARD WING BELLCRANK 19. ROD END

A

17 1 16 15

2

18

6

3 4 5

14

7

13

19

11 10 8

12 9

DETAIL

A UB27B 051939AA.AI

Figure 208 Aileron Outboard Wing Bellcrank (UA-1 and After; UB-1 and After)

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FLIGHT CONTROLS AILERON CONTROL SYSTEM MAINTENANCE PRACTICES

27-10-03 200200

1. PROCEDURES A. Rigging (UA-1 and After; UB-1 and After) WARNING: Perform all Steps of this procedure, in the order listed. Do not skip any Steps of this procedure. Failure to do so may result in injury to personnel and damage to equipment. CAUTION: Where the cables pass through structure, the areas of possible contact between the control cables and adjacent structure must be protected with protective elements such as grommets, rub strips, blocks or guide fairings. Where contact of control cables does occur with the protective elements, a force no greater than eight (8) ounces shall be required to move the cable to a position of no contact. At no time should flight control cables contact metal structure with the protective elements removed. NOTE: Do not connect the actuating components of the autopilot until all flight control systems have been properly rigged. This will permit aileron forces to be measured properly. (1) Remove the pilot’s seat (Ref. Chapter 25-10-00, SEAT REMOVAL). (2) Remove the control column boot located forward of the pedestal. (3) Remove passenger compartment seats as required to access floor access panels 16C and 16E (Ref. Chapter 25-20-00, PASSENGER SEAT REMOVAL). (4) Remove passenger compartment carpet as required to access floor access panels 16C and 16E (Ref. Chapter 25-20-01). (5) Remove floor access panels 16C and 16E (Ref. Chapter 6-50-00, FLOOR ACCESS PANELS). (6) Remove left and right lower wing access panels 7, 44, 45, 46 and 47 as required (Ref. Chapter 6-50-00, WING ACCESS PANELS). (7) Remove left and right upper wing access panels 29 and 30 as required (Ref. Chapter 6-50-00, WING ACCESS PANELS). NOTE: One travel board may be used and moved from one side to the other. (8) Perform the AILERON TRAVEL BOARD INSTALLATION procedure (Ref. 27-00-02). Apply tape to the wing surface to mark the location of the travel board feet ensuring proper location when moving the travel board from one side to the other. (9) If equipped, identify, tag and disconnect the Supplemental Type Certificate (STC) Flight Data Recorder (FDR) sensor bridle clamp located on the aileron fuselage cables. Refer to the STC holders instructions. (10) If equipped, disconnect the autopilot aileron servo cable from the aileron quadrant (Ref. Chapter 22-10-00, AILERON SERVO BRIDLE CABLE REMOVAL).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (11) Install rig pin (1) (7, Table 1, 27-00-00) through the aileron quadrant support bracket (2) and the aileron quadrant (3) (Ref. Figure 204). (12) Ensure the pilot’s control wheel is level within ± 0.5°. If no adjustment is required proceed to Step (13). If adjustment is required perform the following Steps: (a) Locate the turnbuckles (3) on the control column (1) (T-Column) forward of the pedestal and remove safety clips (4) (Ref. Figure 205). (b) Adjust the fuselage aileron control cable turnbuckles (3) to level the pilot’s control wheel. (13) Check the fuselage cable tension between the control column and the aileron quadrant by performing the following Steps: NOTE: Do not measure the tension when the temperature is varying rapidly or with the airplane located in direct sunlight. (a) Allow the temperature of the airframe to stabilize before measuring and adjusting the cable tension. (b) Measure the temperature in the compartment under floor access panel 16C. (c) Refer to Aileron Fuselage Cable Tension Graph, Figure 201, and read the pounds of tension required for the measured temperature. NOTE: Fuselage cable tension tolerance is +10/-0 pounds of the tension found in Figure 201. (d) Position a cable tensiometer (4, Table 1, 27-00-00) on the aileron fuselage cables at least three inches from turnbuckles and pulleys and measure the tension of both cables. Cable diameter is noted in Figure 201. (e) If no adjustment is required, proceed to Step (14). If adjustment is required, perform the following Steps: 1 Locate the turnbuckles (3) on the control column (1) (T-Column) forward of the pedestal. If not previously removed, remove safety clips (4) from both turnbuckles (3) (Ref. Figure 205). WARNING: If cable tension at any time is below 10 pounds, check all aileron fuselage cable system pulleys for proper cable engagement. 2 Adjust the turnbuckles until both fuselage cables have equal tensions needed at the current temperature found in Figure 201. Cable diameter is noted in Figure 201. 3 Remove rig pin (1) (7, Table 1, 27-00-00) from the aileron quadrant (3) (Ref. Figure 204). 4 Using either control wheel, move the aileron system through three cycles to equalize tension in the aileron fuselage system cables. 5 Install rig pin (1) (7, Table 1, 27-00-00) through the aileron quadrant support bracket (2) and the aileron quadrant (3).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL 6 Verify that the tension of both fuselage cables is within acceptable limits per Figure 201. If cable tension is out of limits repeat Step (13). 7 Verify the pilot’s control wheel is still level within ± 0.5°. If not within limits repeat Steps (12) and (13). (14) Verify safety clips (4) are installed on all turnbuckles (3) in the system (Ref. Figure 205). (15) Check the cable tension of the control column interconnect (chain and cable assembly) between the pilot’s and copilot’s control wheel by performing the following Steps: NOTE: The control wheels must align within ± 0.5° of each other. (a) Tension of the control column interconnect must be between 10 to 25 pounds. Cable diameter is 1/8 inch. (b) Position a cable tensiometer (4, Table 1, 27-00-00) on the cable (3) and measure the tension (Ref. Figure 206). (c) If no adjustment is required, proceed to Step (16). If adjustment is required, perform the following Steps: 1 Remove safety clips (4) from both turnbuckles (2). 2 Adjust the turnbuckles (2) to meet the required cable tension in Step (15) (a). 3 Remove rig pin (1) (7, Table 1, 27-00-00) from the aileron quadrant (3) (Ref. Figure 204). 4 Using either control wheel, move the aileron system through three cycles to equalize tension in the chain and cable system. 5 Install rig pin (1) (7, Table 1, 27-00-00) through the aileron quadrant support bracket (2) and the aileron quadrant (3) (Ref. Figure 204). 6 Verify that the tension of the control column interconnect is within acceptable limits per Step (15) (a). If cable tension is out of limits repeat Step (15). (16) Verify the control wheels are aligned within ± 0.5° of each other. If no adjustment is required proceed to Step (17). If adjustment is required perform the following Steps (Ref. Figure 206): (a) If not previously removed, remove safety clips (4) from turnbuckles (2). (b) While maintaining cable tension of 10 to 25 pounds, adjust the turnbuckles (2) until the control wheels are level, then repeat Steps (15) (c) 3 thru (16). (17) Verify safety clips (4) are installed on all turnbuckles (2) in the system. (18) Install rig pin (3) (11, Table 1, 27-00-00) through the aileron outboard wing bellcrank (1) (Ref. Figure 211).

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(19) Check the wing cable tension between the aileron quadrant and the wing outboard bellcrank by performing the following Steps NOTE: Do not measure the tension when the temperature is varying rapidly or with the airplane located in direct sunlight. (a) Allow the temperature of the airframe to stabilize before measuring and adjusting the cable tension. (b) Measure the temperature in the compartment with the aileron cables in the lower wing area. (c) Refer to Aileron Wing Cable Tension Graph, Figure 202, Sheet 1 and read the pounds of tension required for the measured temperature. NOTE: Wing cable tension tolerance is ± 8 pounds of the tension found in Figure 202, Sheet 1. Wing cable diameter is 3/16 inch and the fuselage cable diameter is 1/8 inch. (d) Position a cable tensiometer (4, Table 1, 27-00-00) on the aileron wing cable at least three inches from turnbuckles and pulleys and measure the tension of both cables. Cable diameter is noted in Figure 202, Sheet 1. (e) If no adjustment is required, proceed to Step (20). If adjustment is required, perform the following Steps: WARNING: If cable tension at any time is below 10 pounds, check all aileron wing cable system pulleys for proper cable engagement. 1 Remove safety clips (3) from both turnbuckles (2) (Ref. Figure 207). 2 Adjust the turnbuckles (2) until both cables have equal tensions needed at the current temperature found in Figure 202, Sheet 1. Cable diameter is noted in Figure 202, Sheet 1. 3 Remove rig pin (1) (7, Table 1, 27-00-00) from the aileron quadrant (3) (Ref. Figure 204). 4 Remove rig pin (3) (11, Table 1, 27-00-00) from the aileron wing outboard bellcrank (1) (Ref. Figure 211). 5 Using either control wheel, move the aileron system through three cycles to equalize tension in the aileron wing system cables. 6 Install rig pin (1) (7, Table 1, 27-00-00) through the aileron quadrant support bracket (2) and the aileron quadrant (3) (Ref. Figure 204). 7 Install rig pin (3) (11, Table 1, 27-00-00) through the aileron outboard bellcrank (1) (Ref. Figure 211). 8 Verify that the tension of both cables is within acceptable limits per Figure 202, Sheet 1. If cable tension is out of limits repeat Step (19). 9 Install safety clips (3) on turnbuckles (2) (Ref. Figure 207).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (20) Verify safety clips (3) are installed on all turnbuckles (2) in the system. (21) Check the position of the aileron using the aileron travel board. Aileron must be at neutral position 0° to 0.5° up. (22) If no adjustment is required, proceed to Step (23). If adjustment is required, perform the following Steps: WARNING: If cable tension drops below 10 pounds at any time, check all aileron wing cable system pulleys for proper cable engagement. (a) Remove rig pin (1) (7, Table 1, 27-00-00) from the aileron quadrant (3) (Ref. Figure 204). (b) Remove rig pin (3) (11, Table 1, 27-00-00) from the aileron wing outboard bellcrank (1) (Ref. Figure 211). (c) Remove cotter pin (14), nut (5), washers (4) and bolt (7) attaching the push-pull rod (3) to the outboard wing bellcrank (Ref. Figure 208). (d) Loosen the jamnut (8) on the aileron push-pull rod (3). (e) Rotate rod end (6) as necessary until aileron is at neutral 0° to 0.5° up deflection. (f) Verify that the threads of the rod end (6) are visible through the inspection hole (9) at the end of the push-pull rod (3) after adjustment is completed. (g) Tighten the jamnut (8) on the aileron push-pull rod (3). (h) Install bolt (7), washers (4), nut (5) and new cotter pin (14) attaching the push-pull rod (3) to the outboard wing bellcrank. (i) Install rig pin (1) (7, Table 1, 27-00-00) through the aileron quadrant support bracket (2) and the aileron quadrant (3) (Ref. Figure 204). (j) Install rig pin (3) (11, Table 1, 27-00-00) through the aileron outboard bellcrank (1) (Ref. Figure 211). (k) Verify that the aileron surface is at neutral 0° to 0.5° up deflection. If aileron surface is not within limits repeat Step (22). (23) Perform Steps (18) thru (22) for the opposite aileron. (24) Remove rig pin (1) (7, Table 1, 27-00-00) from the aileron quadrant (3) (Ref. Figure 204). (25) Remove rig pin (3) (11, Table 1, 27-00-00) from the aileron outboard bellcrank (1) (Ref. Figure 211). NOTE: The difference between the full-up and full-down travel of each aileron should be 6° to 9° (i.e. 24° up 17° down = 7°). The difference between the full-up limit of one aileron and the full-down limit of the other (crisscross differential) should be 6° to 9° (i.e. LH 25° up RH 17° down = 8°).

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(26) Check the maximum travel of the aileron surfaces by performing the following Steps: (a) Using minimum force manually move aileron down and check for surface travel of 15° -0°/+3° down on the aileron travel board. Repeat for opposite aileron. (b) Using minimum force manually move aileron up and check for surface travel of 23° -0°/+3° up on the aileron travel board. Repeat for opposite aileron. (c) Verify the difference between the full-up and full-down travel of each aileron is 6° to 9° and the difference between the full-up limit of one aileron and the full-down limit of the other (crisscross differential) is 6° to 9°. (d) Verify all stops make contact and that opposite (crisscross) stops contact at the same time. (e) If no adjustment is required, proceed to Step (27). If adjustment is required, perform the following Steps: NOTE: Adjusting one stop requires the opposite (crisscross) stop be adjusted also. 1 Through left wing access panel 46, loosen jamnut (12) on up stop adjustment bolt (13). Adjust the bolt (13) as necessary to obtain maximum upward deflection of 23° -0°/+3° (Ref. Figure 208). 2 Through right wing access panel 30, loosen jamnut (11) on down stop adjustment bolt (10). Adjust the bolt (10) as necessary to obtain maximum downward deflection of 15° -0°/+3°. 3 Through left wing access panel 29, loosen jamnut (11) on down stop adjustment bolt (10). Adjust the bolt (10) as necessary to obtain maximum downward deflection of 15° -0°/+3°. 4 Through right wing access panel 47, loosen jamnut (12) on up stop adjustment bolt (13). Adjust the bolt (13) as necessary to obtain maximum upward deflection of 23° -0°/+3°. 5 Verify the difference between the full-up and full-down travel of each aileron is 6° to 9° and the difference between the full-up limit of one aileron and the full-down limit of the other (crisscross differential) is 6° to 9°. If not within limits, repeat Steps (26) (e) 1 thru (26) (e) 5. 6 When adjustment is complete, tighten the jamnut(s) (11 and 12) on the stop bolts (10 and 13). 7 Ensure all stops make contact and that opposite (crisscross) stops contact at the same time. (27) Verify safety clips are installed on all turnbuckles in the system. (28) If equipped, connect and rig the aileron autopilot servo cables (Ref. Chapter 22-10-00, AILERON SERVO CABLE RIGGING). (29) If the Supplemental Type Certificate (STC) Flight Data Recorder (FDR) sensor bridle clamp was installed, install bridle clamp. Refer to the STC holders instructions. Refer to the STC holders instructions for installation and calibration. Page 206 Nov 1/09

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (30) Perform the AILERON OPERATIONAL CHECK procedure in this section. (31) When the system is completely rigged, use a force gauge (2) (6, Table 1, 27-00-00) against the control wheel (1) at a point 5 inches out from the center. Verify that the force required to move the control wheel 10° to the left or right of neutral does not exceed 6 pounds (Ref. Figure 210). (32) Connect external electrical power to the airplane. (33) Select the BATT switch to the ON position. (34) Select the EXT PWR switch to the EXT PWR position. (35) Raise flaps to the full up position. (36) Select the EXT PWR switch to the OFF position. (37) Select the BATT switch to the OFF position. (38) Disconnect external electrical power from the airplane. (39) With the flaps fully retracted and the aileron in the neutral position a maximum step of 0.50 inch is allowable between the outboard flap and the aileron at the trailing edge with the flap trailing edge below the aileron trailing edge. (40) Remove the aileron travel board and tape from the wing surface. (41) Install left and right lower wing access panels 7, 44, 45, 46 and 47 (Ref. Chapter 6-50-00, WING ACCESS PANELS). (42) Install left and right upper wing access panels 29 and 30 (Ref. Chapter 6-50-00, WING ACCESS PANELS). (43) Install floor access panels 16C and 16E (Ref. Chapter 6-50-00, FLOOR ACCESS PANELS). (44) Install passenger compartment carpet (Ref. Chapter 25-20-01). (45) Install passenger INSTALLATION).

compartment

seats

(Ref.

Chapter

25-20-00,

PASSENGER

SEAT

(46) Install the control column boot located forward of the pedestal. (47) Install the pilot’s seat (Ref. Chapter 25-10-00, SEAT INSTALLATION).

B. Rigging (UC-1 and After) WARNING: Perform all Steps of this procedure, in the order listed. Do not skip any Steps of this procedure. Failure to do so may result in injury to personnel and damage to equipment. CAUTION: Where the cables pass through structure, the areas of possible contact between the control cables and adjacent structure must be protected with protective elements such as grommets, rub strips, blocks or guide fairings. Where contact of control cables does occur with the protective elements, a force no greater than eight (8) ounces shall be required to move the cable to a position of no contact. At no time should flight control cables contact metal structure with the protective elements removed.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL NOTE: Do not connect the actuating components of the autopilot until all flight control systems have been properly rigged. This will permit aileron forces to be measured properly. (1) Connect external electrical power to the airplane. (2) Select the BATT switch to the ON position. (3) Select the EXT PWR switch to the EXT PWR position. (4) Lower flaps to the full down position. (5) Select the EXT PWR switch to the OFF position. (6) Select the BATT switch to the OFF position. (7) Disconnect external electrical power from the airplane. (8) Remove the pilot’s seat (Ref. Chapter 25-10-00, SEAT REMOVAL). (9) Remove the control column boot located forward of the pedestal. (10) Remove passenger compartment seats as required to access floor access panels 16C and 16E (Ref. Chapter 25-20-00, PASSENGER SEAT REMOVAL). (11) Remove passenger compartment carpet as required to access floor access panels 16C and 16E (Ref. Chapter 25-20-01). (12) Remove floor access panels 16C and 16E (Ref. Chapter 6-50-00, FLOOR ACCESS PANELS). (13) Remove left and right lower wing access panels 19, as required (Ref. Chapter 6-50-00, WING ACCESS PANELS). NOTE: One travel board may be used and moved from one side to the other. (14) Perform the AILERON TRAVEL BOARD INSTALLATION procedure (Ref. 27-00-02). Apply tape to the wing surface to mark the location of the travel board feet ensuring proper location when moving the travel board from one side to the other. (15) If equipped, identify, tag and disconnect the Supplemental Type Certificate (STC) Flight Data Recorder (FDR) sensor bridle clamp located on the aileron fuselage cables. Refer to the STC holders instructions. (16) If equipped, disconnect the autopilot aileron servo cable from the aileron quadrant (Ref. Chapter 22-10-00, AILERON SERVO BRIDLE CABLE REMOVAL). (17) Install rig pin (1) (7, Table 1, 27-00-00) through the aileron quadrant support bracket (2) and the aileron quadrant (3) (Ref. Figure 204). (18) Ensure the pilot’s control wheel is level within ± 0.5°. If no adjustment is required proceed to Step (19). If adjustment is required perform the following Steps: (a) Locate the turnbuckles (3) on the control column (1) (T-Column) forward of the pedestal and remove safety clips (4) (Ref. Figure 205). (b) Adjust the fuselage aileron control cable turnbuckles (3) to level the pilot’s control wheel.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (19) Check the fuselage cable tension between the control column and the aileron quadrant by performing the following Steps: NOTE: Do not measure the tension when the temperature is varying rapidly or with the airplane located in direct sunlight. (a) Allow the temperature of the airframe to stabilize before measuring and adjusting the cable tension. (b) Measure the temperature in the compartment under floor access panel 16C. (c) Refer to Aileron Fuselage Cable Tension Graph, Figure 201, and read the pounds of tension required for the measured temperature. NOTE: Fuselage cable tension tolerance is +10 / -0 pounds of the tension found in Figure 201. (d) Position a cable tensiometer (4, Table 1, 27-00-00) on the aileron fuselage cables at least three inches from turnbuckles and pulleys and measure the tension of both cables. Cable diameter is noted in Figure 201. (e) If no adjustment is required, proceed to Step (20). If adjustment is required, perform the following Steps: 1 Locate the turnbuckles on the control column (T-Column) forward of the pedestal. If not previously removed, remove safety clips (4) from both turnbuckles (3) (Ref. Figure 205). WARNING: If cable tension at any time is below 10 pounds, check all aileron fuselage cable system pulleys for proper cable engagement. 2 Adjust the turnbuckles until both fuselage cables have equal tensions needed at the current temperature found in Figure 201. Cable diameter is noted in Figure 201. 3 Remove rig pin (1) (7, Table 1, 27-00-00) from the aileron quadrant (3) (Ref. Figure 204). 4 Using either control wheel, move the aileron system through three cycles to equalize tension in the aileron fuselage system cables. 5 Install rig pin (1) (7, Table 1, 27-00-00) through the aileron quadrant support bracket (2) and the aileron quadrant (3) (Ref. Figure 204). 6 Verify that the tension of both fuselage cables is within acceptable limits per Figure 201. If cable tension is out of limits repeat Step (19). 7 Verify the pilot’s control wheel is still level within ± 0.5°. If not within limits repeat Steps (18) and (19). (20) Verify safety clips (4) are installed on all turnbuckles (3) in the system (Ref. Figure 205).

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(21) Check the cable tension of the control column interconnect (chain and cable assembly) between the pilot’s and copilot’s control wheel by performing the following Steps: NOTE: The control wheels must align within ± 0.5° of each other. (a) Tension of the control column interconnect must be between 10 to 25 pounds. Cable diameter is 1/8 inch. (b) Position a cable tensiometer (4, Table 1, 27-00-00) on the cable (3) and measure the tension (Ref. Figure 206). (c) If no adjustment is required, proceed to Step (22). If adjustment is required, perform the following Steps: 1 Remove safety clips (4) from both turnbuckles (2). 2 Adjust the turnbuckles (2) to meet the required cable tension in Step (21) (a). 3 Remove rig pin (1) (7, Table 1, 27-00-00) from the aileron quadrant (3) (Ref. Figure 204). 4 Using either control wheel, move the aileron system through three cycles to equalize tension in the chain and cable system. 5 Install rig pin (1) (7, Table 1, 27-00-00) through the aileron quadrant support bracket (2) and the aileron quadrant (3) (Ref. Figure 204). 6 Verify that the tension of the control column interconnect is within acceptable limits per Step (21) (a). If cable tension is out of limits repeat Step (21). (22) Verify the control wheels are aligned within ± 0.5° of each other. If no adjustment is required proceed to Step (23). If adjustment is required perform the following Steps (Ref. Figure 206) (a) If not previously removed, remove safety clips (4) from turnbuckles (2). (b) While maintaining cable tension of 10 to 25 pounds, adjust the turnbuckles (2) until the control wheels are level, then repeat Steps (21) (c) 3 thru (22). (23) Verify safety clips (4) are installed on all turnbuckles (2) in the system. (24) Check the wing cable tension between the aileron quadrant and the wing bellcrank by performing the following Steps: NOTE: Do not measure the tension when the temperature is varying rapidly or with the airplane located in direct sunlight. (a) Allow the temperature of the airframe to stabilize before measuring and adjusting the cable tension. (b) Measure the temperature in the compartment with the aileron cables in the lower wing area. (c) Refer to Aileron Wing Cable Tension Graph, Figure 202, Sheet 2 and read the pounds of tension required for the measured temperature.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL NOTE: Wing cable tension tolerance is ± 8 pounds of the tension found in Figure 202, Sheet 2. Wing cable diameter is 3/16 inch and the fuselage cable diameter is 1/8 inch. (d) Position a cable tensiometer (4, Table 1, 27-00-00) on the aileron wing cable at least three inches from turnbuckles and pulleys and measure the tension of both cables. Cable diameter is noted in Figure 202, Sheet 2. (e) If no adjustment is required, proceed to Step (25). If adjustment is required, perform the following Steps: WARNING: If cable tension at any time is below 10 pounds, check all aileron wing cable system pulleys for proper cable engagement. NOTE: If aileron cable kit 118-5001 is installed the turnbuckles will be accessed through wing access panels 19 inboard of the main landing gear, instead of in the wheel well. 1 Remove safety clips (3) from the turnbuckles (2) (Ref. Figure 207). 2 Adjust the turnbuckles (2) until both cables have equal tensions needed at the current temperature found in Figure 202, Sheet 2. Cable diameter is noted in Figure 202, Sheet 2. 3 Remove rig pin (1) (7, Table 1, 27-00-00) from the aileron quadrant (3) (Ref. Figure 204). 4 Using either control wheel, move the aileron system through three cycles to equalize tension in the aileron wing system cables. 5 Install rig pin (1) (7, Table 1, 27-00-00) through the aileron quadrant support bracket (2) and the aileron quadrant (3) (Ref. Figure 204). 6 Verify that the tension of both cables is within acceptable limits per Figure 202, Sheet 2. If cable tension is out of limits repeat Step (24). (25) Check the position of the aileron using the aileron travel board. Aileron must be at neutral position 0° to 0.5° up. (a) If no adjustment is required, proceed to Step (26). If adjustment is required, perform the following Steps: WARNING: If cable tension at any time is below 10 pounds, check all aileron wing cable system pulleys for proper cable engagement. NOTE: If aileron cable kit 118-5001 is installed the turnbuckles will be accessed through wing access panels 19 inboard of the main landing gear, instead of in the wheel well. 1 Remove safety clips (3) from turnbuckles (2) (Ref. Figure 207). 2 Adjust the aileron wing cable turnbuckles (2) between the quadrant and the wing mounted bellcranks to achieve 0° to 0.5° up at the trailing edge of the aileron. NOTE: Wing cable tension tolerance is ± 8 pounds of the tension found in Figure 202, Sheet 2.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL 3 While adjusting the wing cable turnbuckles, make sure the required cable tension needed at the current temperature found in Figure 202, Sheet 2 is maintained. Cable diameter is noted in Figure 202, Sheet 2. 4 Install safety (3) clips on the turnbuckles (2) (Ref. Figure 207). (26) Verify safety clips (3) are installed on all turnbuckles (2) in the system. (27) Perform Steps (24) thru (26) for the opposite aileron. (28) Remove rig pin (1) (7, Table 1, 27-00-00) from the aileron quadrant (3) (Ref. Figure 204). NOTE: The difference between the full-up and full-down travel of each aileron should be 5° to 8° (i.e. 24° up 17° down = 7°). The difference between the full-up limit of one aileron and the full-down limit of the other (crisscross differential) should be 5° to 8° (i.e. LH 25° up RH 17° down = 8°). (29) Check the maximum travel of the aileron surfaces by performing the following Steps: (a) Using minimum force manually move aileron down and check for surface travel of 16° -0°/+3° down on the aileron travel board. Repeat for opposite aileron. (b) Using minimum force manually move aileron up and check for surface travel of 23° -0°/+3° up on the aileron travel board. Repeat for opposite aileron. (c) Verify the difference between the full-up and full-down travel of each aileron is 5° to 8° and the difference between the full-up limit of one aileron and the full-down limit of the other (crisscross differential) is 5° to 8°. (d) Verify all stops make contact and that opposite (crisscross) stops contact at the same time. (e) If no adjustment is required, proceed to Step (30). If adjustment is required, perform the following Steps: NOTE: Adjusting one stop requires the opposite (crisscross) stop be adjusted also. 1 On the left aileron, loosen jamnut (2) on upper adjustment screw shaft (3) on the adjustable stop and rotate the screw shaft (3) as necessary to obtain maximum downward deflection of 16° -0°/+3° (Ref. Figure 209). 2 On the right aileron, loosen jamnut (8) on lower adjustment screw shaft (6) on the adjustable stop and rotate the screw shaft (6) as necessary to obtain maximum upward deflection of 23° -0°/+3°. 3 On the right aileron, loosen jamnut (2) on upper adjustment screw shaft (3) on the adjustable stop and rotate the screw shaft (3) as necessary to obtain maximum downward deflection of 16° -0°/+3°. 4 On the left aileron, loosen jamnut (8) on lower adjustment screw shaft (6) on the adjustable stop and rotate the screw shaft (6) as necessary to obtain maximum upward deflection of 23° - 0°/+3°.

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5 Verify the difference between the full-up and full-down travel of each aileron is 5° to 8° and the difference between the full-up limit of one aileron and the full-down limit of the other (crisscross differential) is 5° to 8°. If not within limits, repeat Steps (24) (e) 1 thru (24) (e) 5. 6 When adjustment is complete, tighten the jamnut(s) (2 and 8) on the upper or lower adjustment screw shaft(s) (3 and 6) on the adjustable stops. 7 Ensure all stops make contact and that opposite (crisscross) stops contact at the same time. (30) Verify that all turnbuckles are secured with safety clips. (31) If equipped, connect and rig the aileron autopilot servo cables (Ref. Chapter 22-10-00, AILERON SERVO CABLE RIGGING). (32) If the Supplemental Type Certificate (STC) Flight Data Recorder (FDR) sensor bridle clamp was installed, install bridle clamp. Refer to the STC holders instructions. Refer to the STC holders instructions for installation and calibration. (33) Visually check to ensure that aileron travel responds properly to the control wheel movement by performing the following. (a) Move pilot’s control wheel counterclockwise to the left position and make sure that the left aileron moves to the up position and the right aileron moves to the down position smoothly with no unusual noise or binding. (b) Move pilot’s control wheel clockwise to the right position and make sure that the left aileron moves to the down position and the right aileron moves to the up position smoothly with no unusual noise or binding. (c) Repeat Steps (33) (a) and (33) (b) using the copilot’s control wheel. (34) When the system is completely rigged, use a force gauge (2) (6, Table 1, 27-00-00) against the control wheel (1) at a point 5 inches out from the center. Verify that the force required to move the control wheel 10° to the left or right of neutral does not exceed 6 pounds (Ref. Figure 210). (35) Connect external electrical power to the airplane. (36) Select the BATT switch to the ON position. (37) Select the EXT PWR switch to the EXT PWR position. (38) Raise flaps to the full up position. (39) Select the EXT PWR switch to the OFF position. (40) Select the BATT switch to the OFF position. (41) Disconnect external electrical power from the airplane. (42) With the flaps fully retracted and the aileron in the neutral position a maximum step of 0.50 inch is allowable between the outboard flap and the aileron at the trailing edge with the flap trailing edge below the aileron trailing edge.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (43) Remove the aileron travel board and tape from the wing surface. (44) Install left and right wing access panels 19 (Ref. Chapter 6-50-00, WING ACCESS PANELS). (45) Install floor access panels 16C and 16E (Ref. Chapter 6-50-00, FLOOR ACCESS PANELS). (46) Install passenger compartment carpet (Ref. Chapter 25-20-01). (47) Install passenger INSTALLATION).

compartment

seats

(Ref.

Chapter

25-20-00,

PASSENGER

SEAT

(48) Install the control column boot located forward of the pedestal. (49) Install the pilot’s seat (Ref. Chapter 25-10-00, SEAT INSTALLATION).

2. CHECKS A. Operational Check (1) Connect external electrical power to the airplane. (2) Select the BATT switch to the ON position. (3) Select the EXT PWR switch to the EXT PWR position. (4) Lower flaps to the full down position. (5) Select the EXT PWR switch to the OFF position. (6) Select the BATT switch to the OFF position. (7) Disconnect external electrical power from the airplane. (8) Move pilot’s control wheel counterclockwise to the left position and make sure that the left aileron moves to the up position and the right aileron moves to the down position smoothly with no unusual noise or binding. (9) Move pilot’s control wheel clockwise to the right position and make sure that the left aileron moves to the down position and the right aileron moves to the up position smoothly with no unusual noise or binding. (10) Repeat Steps (8) and (9) using the copilot’s control wheel. (11) If requirements are not met, perform the AILERON CONTROL SYSTEM RIGGING procedure in this section.

B. Fuselage Cable Tension Check (1) Move either control wheel to the left and right position three cycles to equalize system tension. (2) Remove passenger seat(s) as required to access floor access panels 16C and 16E (Ref. Chapter 25-20-00, PASSENGER SEAT REMOVAL). (3) Remove passenger compartment carpet as required to access floor access panels 16C and 16E (Ref. Chapter 25-20-01).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (4) Remove floor access panels 16C and 16E (Ref. Chapter 6-50-00, FLOOR ACCESS PANELS). (5) Install rig pin (1) (7, Table 1, 27-00-00) through the aileron quadrant support bracket (2) and the aileron quadrant (3) (Ref. Figure 204). (6) Check the cable tensions by performing the following Steps: NOTE: Do not measure the tension when the temperature is varying rapidly or with the airplane located in direct sunlight. (a) Allow the temperature of the airframe to stabilize before measuring and adjusting the cable tension. (b) Measure the temperature in the fuselage compartment under floor access panel 16C. (c) Refer to Aileron Fuselage Cable Tension Graph Figure 201 and read the pounds of tension required for the measured temperature. NOTE: Fuselage cable tension tolerance is +10/-0 pounds of the tension found in Figure 201. (d) Position a cable tensiometer (4, Table 1, 27-00-00) on the aileron fuselage cables at least three inches from turnbuckles and pulleys and measure the tension of both cables. Cable diameter is noted in Figure 201. (e) If no adjustment is required proceed to Step (7). If adjustment is required, perform the AILERON CONTROL SYSTEM RIGGING procedure in this section. (7) Remove rig pin (1) (7, Table 1, 27-00-00) from the aileron quadrant (3) (Ref. Figure 204). (8) Install floor access panels 16C and 16E (Ref. Chapter 6-50-00, FLOOR ACCESS PANELS). (9) Install passenger compartment carpet (Ref. Chapter 25-20-01). (10) Install passenger seat(s) (Ref. Chapter 25-20-00, PASSENGER SEAT INSTALLATION). (11) Perform the AILERON OPERATIONAL CHECK procedure in this section.

C. Control Column Interconnect Cable Tension Check The control column interconnect is the chain and cable assembly between the pilot’s and copilot’s control wheels. (1) Move either control wheel to the left and right position three cycles to equalize system tension. (2) Remove passenger seat(s) as required to access floor access panel 16E (Ref. Chapter 25-20-00, PASSENGER SEAT REMOVAL). (3) Remove passenger compartment carpet as required to access floor access panel 16E (Ref. Chapter 25-20-01). (4) Remove floor access panel 16E (Ref. Chapter 6-50-00, FLOOR ACCESS PANELS). (5) Install rig pin (1) (7, Table 1, 27-00-00) through the aileron quadrant support bracket (2) and the aileron quadrant (3) (Ref. Figure 204).

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(6) Check the cable tension of the control column interconnect (chain and cable assembly) between the pilot’s and copilot’s control wheel by performing the following Steps: NOTE: The control wheels must align within ± 0.5° of each other. (a) Tension of the control column interconnect must be between 10 to 25 pounds. Cable diameter is 1/8 inch. (b) Verify the control wheels are aligned within ± 0.5° of each other, then position a cable tensiometer (4, Table 1, 27-00-00) on the cable (3) and measure the tension of the cable (Ref. Figure 206). (c) If no adjustment is required, proceed to Step (7). If adjustment is required, perform the AILERON CONTROL SYSTEM RIGGING procedure in this section. (7) Remove rig pin (1) (7, Table 1, 27-00-00) from the aileron quadrant (3) (Ref. Figure 204). (a) Place the right-hand thread cable (5) clevis end over the outboard wing bellcrank forward arm. Install bolt (22), washer (21), nut (20) and cotter pin (19) (Ref. Figure 201). (b) Insert the left-hand thread cable (4) clevis end into the outboard wing bellcrank aft arm. Install bolt (15), washer (16), nut (17) and cotter pin (18). (8) Install floor access panel 16E (Ref. Chapter 6-50-00, FLOOR ACCESS PANELS). (9) Install passenger compartment carpet (Ref. Chapter 25-20-01). (10) Install passenger seat(s) (Ref. Chapter 25-20-00, PASSENGER SEAT INSTALLATION). (11) Perform the AILERON OPERATIONAL CHECK procedure in this section.

D. Wing Cable Tension Check (1) Move either control wheel to the left and right position three cycles to equalize system tension. (2) Remove passenger seat(s) as required to access floor access panel 16E (Ref. Chapter 25-20-00, PASSENGER SEAT REMOVAL). (3) Remove passenger compartment carpet as required to access floor access panel 16E (Ref. Chapter 25-20-01). (4) Remove floor access panels 16E (Ref. Chapter 6-50-00, FLOOR ACCESS PANELS). (5) Install rig pin (1) (7, Table 201, 27-00-00) through the aileron quadrant support bracket (2) and the aileron quadrant (3) (Ref. Figure 204). (6) Remove left and right lower wing access panels 7 (UA-1 and After, UB-1 and After) or 19 (UC-1 and After) as required (Ref. Chapter 6-50-00, WING ACCESS PANELS). (7) Check the wing cable tensions by performing the following Steps: NOTE: Do not measure the tension when the temperature is varying rapidly or with the airplane located in direct sunlight.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (a) Allow the temperature of the airframe to stabilize before measuring and adjusting the cable tension. (b) Measure the temperature in the lower wing compartment with the aileron cables. (c) Refer to Aileron Wing Cable Tension Graph Figure 2, Sheet 1 (UA-1 and After, UB-1 and After) or Figure 202, Sheet 2 (UC-1 and After) and read the pounds of tension required for the measured temperature. NOTE: Wing cable tension tolerance is ± 8 pounds. (d) Position a cable tensiometer (4, Table 1, 27-00-00) on the aileron wing cables at least three inches from turnbuckles and pulleys and measure the cable tension of both cables. Cable diameter is noted in Figure 202, Sheet 1 (UA-1 and After, UB-1 and After) or Figure 202, Sheet 2 (UC-1 and After). (e) Perform Steps (7) (a) thru (7) (d) on opposite aileron wing cables. (f) If no adjustment is required, proceed to Step (8). If adjustment is required, perform the AILERON CONTROL SYSTEM RIGGING procedure in this section. (8) Remove rig pin (1) (7, Table 201, 27-00-00) from the aileron quadrant (3) (Ref. Figure 204). (9) Install floor access panel 16E (Ref. Chapter 6-50-00, FLOOR ACCESS PANELS). (10) Install passenger compartment carpet (Ref. Chapter 25-20-01). (11) Install passenger seat(s) (Ref. Chapter 25-20-00, PASSENGER SEAT INSTALLATION). (12) Install lower wing access panels 7 (UA-1 and After, UB-1 and After) or 19 (UC-1 and After) (Ref. Chapter 6-50-00, WING ACCESS PANELS). (13) Perform the AILERON OPERATIONAL CHECK procedure in this section.

E. Functional Check NOTE: One travel board may be used and moved from one side to the other. (1) Perform the AILERON TRAVEL BOARD INSTALLATION procedure (Ref. 27-00-02). (2) Connect external electrical power to the airplane. (3) Select the BATT switch to the ON position. (4) Select the EXT PWR switch to the EXT PWR position. (5) Lower flaps to the full down position. (6) Select the EXT PWR switch to the OFF position. (7) Select the BATT switch to the OFF position. (8) Disconnect external electrical power from the airplane.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL NOTE: The difference between the full-up and full-down travel of each aileron should be 5° to 8° (6° to 9° for the UA and UB models) (i.e. 24° up 17° down = 7°). The difference between the full-up limit of one aileron and the full-down limit of the other (crisscross differential) should be 5° to 8° (6° to 9° for the UA and UB models) (i.e. LH 25° up RH 17° down = 8°). (9) Move pilot’s control wheel counterclockwise to the full left position and make sure that the left aileron moves to the full up position 23° -0°/+3° and the right aileron moves to the full down position 16° -0°/+3°(15° -0°/+3° for the UA and UB models) smoothly with no unusual noise or binding. (10) Move pilot’s control wheel clockwise to the full right position and make sure that the left aileron moves to the full down position 16° -0°/+3° (15° -0°/+3° for the UA and UB models) and the right aileron moves to the full up position 23° -0°/+3° smoothly with no unusual noise or binding. (11) Verify the difference between the full-up and full-down travel of each aileron is 5° to 8° (6° to 9° for the UA and UB models) and the difference between the full-up limit of one aileron and the full-down limit of the other (crisscross differential) is 5° to 8° (6° to 9° for the UA and UB models). (12) Verify all stops make contact and that opposite (crisscross) stops contact at the same time. (13) If ailerons require adjustment, perform the AILERON CONTROL SYSTEM RIGGING in this section. (14) Remove the travel board.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

POUNDS OF TENSION

1/8" DIAMETER AILERON FUSELAGE CABLE TENSION GRAPH

NOTE: AILERON FUSELAGE CABLE TENSION TOLERANCE +10 / -0 LBS.

UC27B 041761AA.AI

Figure 201 Aileron Fuselage Cable Tension Graph

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POUNDS OF TENSION

3/16" DIAMETER AILERON WING CABLE TENSION GRAPH

UC27B 041951AA

Figure 202 (Sheet 1 of 2) Aileron Wing Cable Tension Graph (UA-1 and After; UB-1 and After)

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POUNDS OF TENSION

3/16" DIAMETER AILERON WING CABLE TENSION GRAPH

UC27B 041767AA

Figure 202 (Sheet 2 of 2) Aileron Wing Cable Tension Graph (UC-1 and After)

27-10-03

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Figure 203 Aileron Clearances

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

1. RIG PIN 2. AILERON QUADRANT SUPPORT BRACKET 3. AILERON QUADRANT

A 1

3

2

AILERON QUADRANT

DETAIL

A UC27B 041762AB.AI

Figure 204 Aileron Quadrant Rig Pin Installation

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1. "T" COLUMN 2. CABLES 3. TURNBUCKLES 4. SAFETY CLIPS

A

1

2

3

4

VIEW LOOKING RIGHT FWD DETAIL

A

Figure 205 Aileron Fuselage Cable Turnbuckles

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UC27B 043845AA.AI

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

A 1. "T" COLUMN 2. TURNBUCKLES 3. CABLES 4. SAFETY CLIPS 5. CHAIN

2 1

5

4

3

VIEW FORWARD LOOKING UP DETAIL

A

UC27B 043786AA.AI

Figure 206 Control Column Interconnect Cable Turnbuckles

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1. LEFT MAIN LANDING GEAR 2. TURNBUCKLES 3. SAFETY CLIPS

A

2

1

3

OUTBD VIEW LOOKING AFT DETAIL

A

Figure 207 Aileron Wing Cable Turnbuckles

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UC27B 043844AA.AI

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

1. RIG PIN HOLE 2. BELLCRANK 3. PUSH-PULL ROD 4. WASHER 5. NUT 6. ROD END 7. BOLT 8. JAMNUT 9. INSPECTION HOLE 10. DOWN STOP ADJUSTMENT BOLT 11. JAMNUT 12. JAMNUT 13. UP STOP ADJUSTMENT BOLT 14. COTTER PIN

A B

3

2 1

ADJUST ROD END FROM BELLCRANK AREA

VIEW LOOKING UP

FWD

11 13

12

10

DETAIL

A 7 8 9

2 6 2 4 5 14 DETAIL

B

UA27B 050118AA.AI

Figure 208 Aileron Position Adjustment (UA-1 and After; UB-1 and After)

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2 1

A B 3 1. UPPER WING SURFACE 2. JAMNUT 3. DOWN STOP ADJUSTMENT SCREW 4. UPPER AILERON SURFACE 5. LOWER WING SURFACE 6. UP STOP ADJUSTMENT SCREW 7. LOWER AILERON SURFACE 8. JAMNUT

4

DETAIL

A 6 7

5 8

DETAIL

B

Figure 209 Aileron Stop Adjustment (UC-1 and After)

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UC27B 043787AA.AI

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

1. CONTROL WHEEL 2. FORCE GAUGE

A

P

NO S

I

ED

N

T C H NO S

T R I EU

1

M P

2 5 INCHES

DETAIL

A UC27B 043700AA.AI

Figure 210 Control Wheel Force Check

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1. BELLCRANK 2. LOWER SUPPORT BRACKET 3. RIG PIN 4. UPPER SUPPORT BRACKET

4 1 2

A

3

VIEW LOOKING AFT AND UP DETAIL

B UA27B 051940AA.AI

Figure 211 Aileron Outboard Wing Bellcrank Rig Pin Installation (UA-1 and After; UB-1 and After)

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

FLIGHT CONTROLS AILERON TRIM TAB MAINTENANCE PRACTICES

27-10-04 200200

1. PROCEDURES A. Removal (UA-1 and After; UB-1 and After) NOTE: The double clevis end on the trim tab actuator push-pull rod is designed to tighten the outer clevis with a binding action against the clevis bolt and the inner clevis, which removes freeplay of the bolt in the hole. Loosen the outer clevis first (the larger nut) before removing the clevis bolt. (1) Remove nut, washer and bolt (9) and disconnect the trim tab actuator push-pull rod (13) from the trim tab horn (8) (Ref. Figure 201). (2) Remove screw (3), washers (2) and nut (1) from the bonding jumpers (4) at each end of the trim tab. Discard the nuts. (3) Remove four screws (5) (2 upper and 2 lower) from the inboard and outboard trim tab hinge halves (7). (4) Remove the trim tab (6) from the aileron (14). NOTE: If the same trim tab is to be installed on the airplane, the trim tab hinge halves may remain attached to the aileron hinge halves. If a new trim tab is to be installed, the trim tab hinge halves should be removed from the aileron and the new hinge halves from the new trim tab should be installed.

B. Installation (UA-1 and After; UB-1 and After) NOTE: Any repair, modification, painting or replacement of the aileron or the aileron trim tab requires balancing (Ref. Chapter 57-50-00). If a new trim tab is being installed, remove both hinge halves from the new aileron and install them in their respective positions on the aileron hinge halves. (1) Align the trim tab hinge halves (7) into position with the aileron trim tab (6) and install the four attaching screws (5) to the inboard and outboard hinge point. (2) Install screws (3), washers (2) and new nuts (1) connecting the bonding jumpers (4) to each end of the aileron trim tab (6) (Ref. Figure 201). NOTE: The double clevis end on the trim tab actuator push-pull rod is designed to tighten the outer clevis with a binding action against the clevis bolt and the inner clevis, which removes freeplay of the bolt in the hole. Tighten the outer clevis first (the larger nut) after installing the clevis bolt. (3) Connect the actuator push-pull rod (13) to the trim tab horn (8) using bolt, washers and nut (9). (4) Perform the AILERON TRIM TAB FUNCTIONAL CHECK procedure (Ref. 27-10-07). (5) Perform the AILERON TRIM TAB FREEPLAY CHECK procedure in this section.

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6

8 9 11 12

9

13

10

A

DETAIL

B

3 5

2 1 4

6

7

1. NUT 2. WASHERS 3. SCREW 4. BONDING JUMPER 5. SCREWS 6. AILERON TAB 7. HINGE HALVES 8. TRIM TAB HORN 9. BOLT, WASHERS, NUT 10. DOUBLE CLEVIS 11. LARGE JAMNUT 12. SMALL JAMNUT 13. PUSH-PULL ROD 14. AILERON

5

B

1 2 4 7

2 3

14

DETAIL

A

Figure 201 Aileron Trim Tab Installation (UA-1 and After; UB-1 and After)

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2

UA27B 044056AA.AI

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

C. Removal (UC-1 and After) NOTE: The double clevis end on the trim tab actuator push-pull rod is designed to tighten the outer clevis with a binding action against the clevis bolt and the inner clevis, which removes freeplay of the bolt in the hole. Tighten the outer clevis first (the larger nut) after installing the clevis bolt. (1) Remove nut, washer and bolt (9) and disconnect the trim tab actuator push-pull rod (13) from the trim tab horn (8) (Ref. Figure 202). (2) Remove screws (3), washers (2) and nuts (1) from the bonding jumpers (4) at each end of the trim tab. Discard the nuts. (3) Remove four screws (5) (2 upper and 2 lower) from the inboard and outboard trim tab hinge halves (7). (4) Remove the trim tab (6) from the aileron (14). NOTE: If the same trim tab is to be reinstalled on the airplane, the trim tab hinge halves may remain attached to the aileron hinge halves. If a new trim tab is to be installed, the trim tab hinge halves should be removed from the aileron and the new hinge halves from the new trim tab should be installed.

D. Installation (UC-1 and After) NOTE: Any repair, modification, painting or replacement of the aileron or the aileron trim tab, requires balancing (Ref. Chapter 57-50-00). If a new trim tab is being installed, remove both hinge halves from the new aileron and install them in their respective positions on the aileron hinge halves. (1) Align the trim tab hinge halves (7) into position on the aileron trim tab (6) and install the four attaching screws (5) to the inboard and outboard hinge point. (2) Install screws (3), washers (2) and new nuts (1) connecting the bonding jumpers (4) to each end of the aileron trim tab (6) (Ref. Figure 202). NOTE: The double clevis end on the trim tab actuator push-pull rod is designed to tighten the outer clevis with a binding action against the clevis bolt and the inner clevis, which removes freeplay of the bolt in the hole. Tighten the outer clevis first (the larger nut) after installing the clevis bolt. (3) Connect the actuator push-pull rod (13) to the trim tab horn (8) using bolt, washer and nut (9). (4) Perform AILERON TRIM TAB FUNCTIONAL CHECK procedure (Ref. 27-10-07). (5) Perform AILERON TRIM TAB FREEPLAY CHECK procedure in this section.

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6 9 9 10 8 9

11 12 13 9

3

DETAIL

2

1. NUT 2. WASHERS 3. SCREW 4. BONDING JUMPER 5. SCREWS 6. AILERON TAB 7. HINGE HALVES 8. TRIM TAB HORN 9. BOLT, WASHERS, NUT 10. DOUBLE CLEVIS 11. LARGE JAMNUT 12. SMALL JAMNUT 13. PUSH-PULL ROD 14. AILERON

B

5

1 4

6 7

B

5

1 2 4 2

7

2 3 14

DETAIL

A UC27B 044054AB.AI

Figure 202 Aileron Trim Tab Installation (UC-1 and After)

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E. Freeplay Check Visually inspect the aileron tab for any damage, security of hinge attach point and for tightness of the actuating system. Inconsistencies should be remedied prior to checking the freeplay of the tab. A trim tab deflection check fixture (1, Table 1, 27-00-00) or equivalent as shown in Figure 1, 27-00-00, a dial indicator (2, Table 1, 27-00-00), a back screw (3, Table 1, 27-00-00) machined as shown in Figure 1, 27-00-00 and a push-pull scale (6, Table 1, 27-00-00) for applying accurate loading to the trim tabs are required for making the inspection for freeplay of the trim tabs. (1) Obtain a copy of Table 201. (2) UA-1 and After; UB-1 and After perform the following Steps: (a) Remove left and right lower wing access panels 44 and 45 (Ref. Chapter 6-50-00, WING ACCESS PANELS). (b) Install rig pin (3) (11, Table 1, 27-00-00) through the aileron outboard wing bellcrank (1) (Ref. Figure 205). (3) UC-1 and After perform the following Steps: (a) Remove the passenger seat(s) as required to gain access to floor access panel 16E (Ref. Chapter 25-20-00, SEAT REMOVAL). (b) Remove the carpet as (Ref. Chapter 25-20-01).

required

to

gain

access

to

floor

access

panel

16E

(c) Remove floor access panel 16E (Ref. Chapter 6-50-00, FLOOR ACCESS PANELS). (d) Install rig pin (1) (7, Table 1, 27-00-00) through the aileron quadrant support bracket (2) and the aileron quadrant (3) (Ref. Figure 204). (4) Apply tape (for paint protection) on the top surface of the trim tab at a point 2.50 inches aft of the trim tab hinge line along the centerline of the trim tab actuator. This will be the point of pressure against the trim tab by the push-pull scale. (5) Apply tape in the corresponding position on the bottom surface of the trim tab for the same purpose. WARNING: Ensure the trim tab freeplay check fixture (1) is securely attached to the aileron (4) before releasing supporting pressure to prevent damage to equipment and injury to personnel (Ref. Figure 203). (6) Secure the trim tab freeplay check fixture (1) to the aileron (4) so that the dial indicator stem (5) tip is positioned on the top surface of the aileron trim tab (3) 2.50-inches aft of the trim tab hinge line on the outboard edge of the aileron trim tab (3). (7) Position the dial indicator (2) so the stem (5) is depressed 0.10-inch when in contact with the aileron trim tab (3) surface initially. Turn the rotating face of the dial indicator (2) to zero. Do not reset during the checking procedure. (8) With the push-pull scale against the top side, apply a full 3 pound downward load. Record the dial reading on Table 201 as A.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (9) Release half the load until a 1.5 pound downward load is obtained. Record the dial reading as B. (10) Apply a full 3 pound upward load on the bottom surface. Record the dial reading as C. (11) Release half the load until a 1.5 pound upward load is obtained. Record the dial reading as D. (12) Perform the calculations to the data on Table 201. Table 201 AILERON TRIM TAB FREE PLAY LIMITS Serial Number: _______________

Date: _______________

(__________) X 2

= (__________)

- (__________)

= (__________)

B

2B

A

X

(__________) X 2

= (__________)

- (__________)

= (__________)

D

2D

C

Y

(__________)

+ (__________)

= (__________)

X

Y

E (E = 0.053 inch maximum)

(a) Record A, B, C and D as positive numbers. (b) Multiply B by 2 and record as 2B. (c) Subtract A from 2B and record as X. (d) Multiply D by 2 and record as 2D. (e) Subtract C from 2D and record as Y. NOTE: The results of X and Y can be a negative number. (f) Add X and Y and record as E. The maximum allowable freeplay is 0.053 inch. (13) If deflection of the trim tab is within the allowable limits (E= 0.053 inch maximum), the trim tab and its linkage are in good condition. (14) If the free play is excessive, disconnect the trim tab actuator rod and visually inspect the bolts and bushing for indications of excessive wear. Replace excessively worn parts. (15) If all associated linkage is in good condition (no excessive wear) the trim tab actuator needs to be checked for excessive play and/or replaced. (16) UC-1 and After perform the following Steps: (a) Remove the rig pin (1) from the aileron quadrant (3) (Ref. Figure 204). (b) Install floor access panel 16E (Ref. Chapter 6-50-00, FLOOR ACCESS PANELS). (c) Install the carpet as required (Ref. Chapter 25-20-01). Page 206 Nov 1/09

27-10-04

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (d) Install the passenger seat(s) as required (Ref. Chapter 25-20-00, SEAT INSTALLATION). (17) UA-1 and After; UB-1 and After perform the following Steps: (a) Remove rig pin (3) (11, Table 1, 27-00-00) from the aileron outboard wing bellcrank (1) (Ref. Figure 205). (b) Install left and right lower wing panels 44 and 45 (Ref. Chapter 6-50-00, WING ACCESS PANELS).

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1. TAB FREEPLAY CHECK FIXTURE 2. DIAL INDICATOR 3. AILERON TRIM TAB 4. AILERON 5. DIAL INDICATOR STEM

A

2

1

10

30

20

5 40

10

50

3

30

20

40

D UNITE

4

DETAIL

A UC27B 042931AA.AI

Figure 203 Aileron Trim Tab Freeplay Fixture Installation

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

1. RIG PIN 2. AILERON QUADRANT SUPPORT BRACKET 3. AILERON QUADRANT

A 1

3

2

AILERON QUADRANT

DETAIL

A UC27B 041762AB.AI

Figure 204 Aileron Quadrant Rig Pin Installation

27-10-04

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

1. BELLCRANK 2. LOWER SUPPORT BRACKET 3. RIG PIN 4. UPPER SUPPORT BRACKET

4 1 2

A

3

VIEW LOOKING AFT AND UP DETAIL

B UA27B 051940AA.AI

Figure 205 Aileron Outboard Wing Bellcrank Rig Pin Installation

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27-10-04

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

FLIGHT CONTROLS AILERON TRIM TAB ACTUATORS AND CABLES MAINTENANCE PRACTICES

27-10-05 200200

1. AILERON TRIM TAB FORWARD CABLE A. Removal CAUTION: If tension in the system is lost during cable maintenance, check the forward cable drum and aileron trim tab actuator cable drum to ensure the cable is not unwound from drums. (1) Attach a red tag to the aileron trim tab control knob with the words “Do Not Operate, Maintenance In Progress”. (2) Remove both flight compartment seats (Ref. Chapter 25-10-00, SEAT REMOVAL). (3) Remove flight compartment carpet as required. (4) Remove the pedestal side panels as required. (5) Remove the flight compartment floor access panels 1 left, 2 left, 4 and 21 (Ref. Chapter 6-50-00, FLOOR ACCESS PANELS). (6) Remove left passenger compartment seats as required to access floor access panels 16A thru 16E (Ref. Chapter 25-20-00, PASSENGER SEAT REMOVAL). (7) Remove left passenger compartment carpet as required to access floor access panels 16A thru 16E (Ref. Chapter 25-20-01). (8) Remove passenger compartment floor access panels 16A thru 16E (Ref. Chapter 6-50-00, FLOOR ACCESS PANELS). (9) Remove belly access panel 3 just aft of the nose landing gear wheel well (Ref. Chapter 6-50-00, FUSELAGE ACCESS PANELS). (10) Remove lower left wing access panels 7, 8 and 9 (UA-1 and After; UB-1 and After) or 19 (UC-1 and After) as required to access aileron trim tab cables, and pressure seals (Ref. Chapter 6-50-00, WING ACCESS PANELS). (11) Remove upper left wing access panel 4 (UA-1 and After; UB-1 and After only) (Ref. Chapter 6-50-00, WING ACCESS PANELS). (12) Adjust the aileron trim tab control knob to allow easy access to the turnbuckles in the left wheel well and lower wing access panel 19 (UC-1 and After) or upper wing panel 4 (UA-1 and After; UB-1 and After) (Ref. Figure 203, Detail A, B and D). (13) Install a cable block (4 or 13) across both aileron trim tab actuator cables (6 or 12) just inboard of WS 135 to prevent loss of aileron trim tab actuator cable tension (Ref. Figure 203, Detail B or D). (14) Remove the trim tab cable turnbuckle safety clips (2) and loosen the turnbuckles (3) (Ref. Figure 203, Detail A and B) just enough to tape the cables together just below the cable drum in the pedestal to prevent backlash of the cables on the drum (Ref. Figure 201, Detail E).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL NOTE: Each turnbuckle (3 and 10) barrel has a groove (5 and 11) at one end to identify the left-hand threaded end (Ref. Figure 203, Detail A, B and D). (15) Attach a tag with the words “left-hand threads terminal end” to the inboard end of the lower turnbuckle (3 and 10) (Ref. Figure 203, Detail B and D). (16) Disconnect the left-hand threads terminal end from the lower turnbuckle (3 and 10) and attach a feed line to the terminal end. Label the feed line with the words “left-hand threads terminal end”. (17) Attach a tag with the words “right-hand threads terminal end” to the inboard end of the upper turnbuckle (3) (Ref. Figure 203, Detail A). (18) Disconnect the right-hand threads terminal end from the upper turnbuckle (3) and attach a feed line to the terminal end. Label the feed line with the words “right-hand threads terminal end”. (19) Remove the cable stop plate (3) and the safety wire (2) from the cable stops (1) (Ref. Figure 204). (20) Remove cable guard pins from pulley brackets. Refer to Figure 205 or 206 for a general location of the pulleys. (21) Remove the fairleads and cable pressure seal from the fuselage. (22) Remove pulleys as required to allow passage of cable stops. NOTE: If cable is to be reused, take care to keep the cable clean and free from damage. Coil cable loosely, no tighter than ten inches in diameter. (23) While pulling the feed lines through the fuselage, withdraw both left and right-hand threads terminal ends through the fuselage and out of the belly access panel 3 (Ref. Chapter 6-50-00, FUSELAGE ACCESS PANELS). CAUTION: Do not drop the nut, washer or taper pin into the fuselage area below the pedestal. (24) Remove nut, washer and tapered pin (4) from the aileron trim control shaft forward universal joint (18) (Ref. Figure 205 or 206). (25) Slide cable drum shaft (1) forward and remove shaft (1), cable drum (2), guard (3) and washers (17) from pedestal bracket. (26) Note position and number of washer(s) (17) for installation of the cable drum (2). (27) Remove the cable guard (3) and drum (2) together through the right side forward pedestal into the cockpit area. (28) Remove the forward cable from the airplane by routing the feed lines through the forward right side of the pedestal, and into the cockpit. (29) Disconnect the feed lines from forward cable left and right-hand threads terminal ends and leave feed lines in place. (30) Unwrap cable from the drum, remove the cable lock pin and remove the cable (Ref. Figure 201).

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B. Installation WARNING: Perform all Steps of this procedure, in the order listed. Do not skip any Steps of this procedure. Failure to do so may result in injury to personnel and damage to equipment. CAUTION: Where the cables pass through structure, the areas of possible contact between the control cables and adjacent structure must be protected with protective elements such as grommets, rub strips, blocks or guide fairings. Where contact of control cables does occur with the protective elements, a force no greater than eight (8) ounces shall be required to move the cable to a position of no contact. At no time should flight control cables contact metal structure with the protective elements removed. If tension in the system is lost during cable maintenance, check the forward cable drum and aileron trim tab actuator cable drum to ensure the cable is not unwound from drums. (1) If a used cable is to be installed, clean with solvent (2, Table 2, 27-00-00) and check for corrosion and damage. Replace cable if necessary. Dip the cable in corrosion preventive compound (4, Table 2, 27-00-00). Excess should be removed by wiping with a clean rag. NOTE: Washers (5) may be added as required to obtain a cable drum end play of 0.031 to 0.063 inch (Ref. Figure 202). (2) Place the cable drum (2) into the guard (4). Position the cable drum (2), guard (4) and washer(s) (5) in the cable drum bracket (3). Insert the drum shaft and check that the cable drum end play is 0.031 to 0.063 inch. (3) When the correct end play is obtained, remove the drum shaft, cable drum (2), guard (4) and washer(s) (5) (set aside the washers for installation). It is permissible to glue the washer(s) (5) in place against the face of the drum (2) for ease of installation. (4) Attach a tag labeled “left-hand threads terminal end” to the forward cable left-hand threads terminal end. (5) Attach a tag labeled “right-hand threads terminal end” to the forward cable right-hand threads terminal end. (6) Wrap the forward cable on the cable drum as follows (Ref. Figure 201): CAUTION: Do not kink the cable while locating the middle of the forward cable. Damage to the cable will occur. (a) Align the terminal ends of the forward cable and carefully mark the midpoint of the cable with ink or paint. With the left-hand threads terminal end side of the cable located on the flat side of the drum, position the mark on the cable in the middle of the cable drum slot and install the cable lock pin (Ref. Detail B). (b) From the lock pin, wrap each cable 2 1/4 turns around the drum beginning with the outside grooves and work toward the middle of the drum (Ref. Detail C). With the drum wound, verify that the terminal ends are still aligned.

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(c) Position the cable guard over the drum and tape the forward cables together just outside of the cable guard to prevent cable backlash at the drum and maintain the drum in the neutral position during installation (Ref. Detail E). When applying tape to the cable, make sure the cables are separated (not crossed) so that it is easy to identify which cable end winds off the left and right side of the drum. (7) In the cockpit attach the right-hand threads terminal end to the feed line labeled “right-hand threads terminal end”. (8) In the cockpit attach the left-hand threads terminal end to the feed line labeled “left-hand threads terminal end”. NOTE: More than one person will be required to route the forward cable. Take precautions to keep the cable clean and free from damage. (9) Pull the feed lines from the belly access panel 3 (Ref. Chapter 6-50-00, FUSELAGE ACCESS PANELS), draw the forward cables through the right side of the pedestal and then out of the belly access panel, until the drum is close to the pedestal. (10) Position the cable drum (2), guard (3) and washer(s) (17) into the drum bracket and install the cable drum shaft (1) (Ref. Figure 205 or 206). (11) On the left side of the upper pedestal, move the aileron trim tab control knob (7) to the 0 position on the dial indicator (6) and align the slot in the forward universal joint (18) with the cable drum shaft (1) and install the tapered pin (4), washer and nut. (12) Identify the forward cable (8) with the left-hand threads terminal end and make sure it winds off the right side of the drum as installed. Identify the forward cable (9) with the right-hand threads terminal end and make sure it winds off the left side of the drum as installed. Route the cable from the drum as follows (Ref. Figure 205 or 206): NOTE: It is permissible to install cable guard pins, fairleads and pulleys as the cable is being routed. (a) Route cable (8) left-hand thread, through the forward pulley at FS 105. (b) Route cable (9) right-hand thread, through the aft pulley at FS 105. (c) Route cable (8) through the bottom pulley at FS 106. (d) Route cable (9) through the top pulley at FS 106. (e) Route cable (8) through the aileron quadrant bottom pulley at FS 319. (f) Route cable (9) through the aileron quadrant top pulley at FS 319. (g) Route cable (8) through the forward pulley at FS 320. (h) Route cable (9) through the aft pulley at FS 320. (i) Route cable (8) through the forward pressure seal hole at BL 27.

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(j) Route cable (9) through the aft pressure seal hole at BL 27. (k) Install any removed pulleys, fairleads and cable guard pins. (13) Lubricate turnbuckles with grease (1, Table 2, 27-00-00) for corrosion protection prior to installation. (14) Remove feed line, and attach the left-hand threads terminal end of cable (8) to the bottom turnbuckle. (15) Remove feed line, and attach the right-hand threads terminal end of cable (9) to the top turnbuckle. (16) Tension the forward cable sufficient to prevent slack. (17) Install the cable stop plate (3) (Ref. Figure 204). (18) Ensure that the forward cable is routed properly by verifying that the cable has been routed exactly as described in Step (12). Ensure cable is engaged in the pulley grooves and all guard pins are installed. (19) Remove cable block (4 or 13) from the trim tab actuator cables (6 or 12) (Ref. Figure 203, Detail B or D). (20) Remove all tape from the cable and turnbuckles. (21) Lubricate the control cables to one inch beyond the length of travel through the pressure seal with grease (1, Table 2, 27-00-00). (22) Fill the cable pressure seal with grease (1, Table 2, 27-00-00), and install the seal in the hole at BL 27 and seal with sealant (12, Table 2, 27-00-00). (23) Perform the AILERON TRIM TAB OPERATIONAL CHECK procedure (Ref. 27-10-07). (24) Perform the AILERON TRIM TAB RIGGING procedure (Ref. 27-10-07). (25) Ensure turnbuckles (3) have been safety clipped (2) (Ref. Figure 203, Detail A and B). (26) Ensure the aileron trim tab cable stops (1) have been properly safety wired (2) (Ref. Figure 204). (27) Install passenger compartment floor access panels 16A thru 16E (Ref. Chapter 6-50-00, FLOOR ACCESS PANELS). (28) Install passenger compartment carpet (Ref. Chapter 25-20-01). (29) Install passenger INSTALLATION).

compartment

seats

(Ref.

Chapter

25-20-00,

PASSENGER

SEAT

(30) Install the flight compartment floor access panels (Ref. Chapter 6-50-00, FLOOR ACCESS PANELS). (31) Install the pedestal side panels. (32) Install flight compartment carpet.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (33) Install both flight compartment seats (Ref. Chapter 25-10-00, SEAT INSTALLATION). (34) Install belly access panel 3 just aft of the nose landing gear wheel well (Ref. Chapter 6-50-00, FUSELAGE ACCESS PANELS). (35) Install lower left wing access panels 7, 8 and 9 (UA-1 and After; UB-1 and After) or 19 (UC-1 and After) (Ref. Chapter 6-50-00, WING ACCESS PANELS). (36) Install upper left wing access panel 4 (UA-1 and After; UB-1 and After Only) (Ref. Chapter 6-50-00, WING ACCESS PANELS). (37) Remove red tag from the aileron trim tab control knob.

2. AILERON TRIM TAB ACTUATOR AND CABLE A. Removal (UA-1 and After; UB-1 and After) CAUTION: If tension in the system is lost during cable maintenance, check the forward cable drum and aileron trim tab actuator cable drum to ensure the cable is not unwound from drums. (1) Attach a red tag to the aileron trim tab control knob with the words “Do Not Operate, Maintenance In Progress”. (2) Perform the left AILERON REMOVAL (UA-1 AND AFTER; UB-1 AND AFTER) procedure (Ref. 27-10-00). (3) Remove lower left wing access panels 7, 35, 37, 43, 44 and 46 as required to access aileron trim tab cables and turnbuckles (Ref. Chapter 6-50-00, WING ACCESS PANELS). (4) Remove the upper left wing access panels 4 and 29 (Ref. Chapter 6-50-00, WING ACCESS PANELS). (5) Adjust the aileron trim tab control knob to allow easy access to the turnbuckles (3 and 10) in the left wheel well and upper wing access panel 4 (Ref. Figure 203, Detail A and D). (6) Install a cable block (7) across both aileron trim tab forward cables (6) just outboard of WS 79.65 to prevent loss of aileron trim tab forward cable tension (Ref. Figure 203, Detail C). (7) Remove safety clips (2 and 9) from the turnbuckles (3 and 10) (Ref. Figure 203, Detail A and D). NOTE: Each turnbuckle (3 and 10) barrel has a groove (5 and 12) at one end to identify the left-hand threaded end (Ref. Figure 203, Detail A, B and D). (8) Attach a tag with the words “left-hand threads terminal end” to the outboard end of the top turnbuckle (3 or 10) (Ref. Detail A and D). (9) Disconnect the left-hand threads terminal end from the top turnbuckle (3 or 10) and attach a feed line to the terminal end. Label the feed line with the words “left-hand threads terminal end”. (10) Attach a tag with the words “right-hand threads terminal end” to the outboard end of the bottom turnbuckle (3 or 10) (Ref. Detail B and D). (11) Disconnect the right-hand threads terminal end from the bottom turnbuckle (3) and attach a feed line to the terminal end. Label the feed line with the words “right-hand threads terminal end”.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (12) Remove the cable guard pins from pulley brackets. NOTE: The double clevis end (4) on the trim tab actuator push-pull rod (8) is designed to tighten the outer clevis with a binding action against the clevis bolt (10) and the inner clevis, which removes freeplay of the bolt in the hole. Loosen the outer clevis first (the larger nut (5)) before removing the clevis bolt (10) (Ref. Figure 211). (13) Loosen the large jam nut (5) on the double clevis (4). (14) Remove nut and washer (3) and bolt (10) and remove the aileron trim tab push-pull rod (8) assembly from the aileron trim tab actuator (12). (15) With assistance, feed the trim tab actuator cables through the wing and remove the cables through the actuator access opening. (16) Detach the feed lines from the trim tab cable terminal ends leaving the feed lines in place. (17) Remove the mount bolts (14) and screw and washer (15) attaching the actuator (13) to the wing and remove actuator (13) (Ref. Figure 205).

B. Drum Cable Replacement (UA-1 and After; UB-1 and After) (1) Perform the AILERON TRIM TAB ACTUATOR AND CABLE REMOVAL procedure in this section. (2) Through the holes in housing half (9), remove nut (17), washer (16) and bolt (14) from the guide (15) (Ref. Figure 209). (3) Remove nut (13), washer (12), bolt (18), washer (27) and separate the actuator end cap (21) from the actuator housing half (9). (4) Remove bearing (22). (5) Remove the safety wire and loosen jam nut (2). (6) Remove rod end (1), jam nut (2) and key washer (3). NOTE: If the cable drum (23) is hard to remove, ensure the threads of the actuator screw (7) are fully engaged and lightly tap on the actuator screw to push the cable drum (23) out. (7) Remove the drum (23) from the actuator. (8) Remove actuator screw (7). NOTE: If cable is to be reused, take care to keep the cable clean and free from damage. Coil cable loosely, no tighter than ten inches in diameter. (9) Unwind the cable (26) from the drum (23). (10) Slide the cable lock pin (25) from the drum (23) and remove cable (26). (11) If used cable is to be installed, clean cable (26) with solvent (2, Table 2, 27-00-00) and check cable for corrosion and damage. Replace cable if necessary. Dip the cable in corrosion preventative compound (4, Table 2, 27-00-00). Remove excess corrosion preventative by wiping with a clean cloth.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (12) Attach a tag labeled “left-hand threads terminal end” to the actuator cable left-hand threads terminal end. (13) Attach a tag labeled “right-hand threads terminal end” to the actuator cable right-hand threads terminal end. (14) Wrap the aileron trim tab actuator cable on the cable drum as follows (Ref. Figure 207): CAUTION: Do not kink the cable while locating the middle of the actuator cable. Damage to the cable will occur. (a) Align the terminal ends of the forward cable and position keeper at the midpoint of the cable. With the right-hand threaded terminal end toward the end of the drum that faces the end cap, center the keeper and aileron trim tab cable in the slot of the drum. Cable ends must be even (Ref. Detail B). (b) From the keeper, wrap each cable end 4 1/4 turns around the drum beginning with the outside grooves and work toward the middle of the drum (Ref. Detail C). With the drum wound, verify that the cable ends are still aligned. (c) Insert drum (23) into the actuator housing half (9). Tape the aileron trim tab actuator cables together just outside of the actuator housing to prevent cable backlash at the drum (Ref. Figure 209). NOTE: Grease (17, Table 2, 27-00-00) may be used. CAUTION: Do not mix greases. Mixing greases reduces lubricant effectiveness. The actuator was originally manufactured with grease (1, Table 2, 27-00-00) and this grease may have been cleaned out and replaced with grease (17, Table 2, 27-00-00). (15) Lightly lubricate the actuator screw (7) threads with grease (1, Table 2, 27-00-00). (16) Install the actuator screw (7) completely into the actuator housing (9) and drum (23) to distribute the grease onto the mating threads. NOTE: Ensure the actuator screw (7) passes through the guide (15). (17) Remove the actuator screw (7) and lubricate with grease again and install into the drum (23). (18) Install bearing (22). NOTE: If there is end play in the drum (23), adjust the bearing adjustment ring (19) until all end play is removed and the drum (23) turns freely. (19) Install the end cap (21) on the actuator housing half (9). (20) Install bolt (18), washers (12 and 27) and nut (13) (three places). (21) With cable ends even, adjust the actuator screw (11) to attain a measurement of 0.98 ± 0.06 inch from the top of the actuator screw (11) to the actuator housing half (12) (Ref. Figure 211). (22) Through the holes in housing half (9), install nut (17), washer (16) and bolt (14) through the guide (15) and actuator screw (7) (Ref. Figure 209). (23) Install jam nut (2) and key washer (3) on rod end (1).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (24) Install rod end (1) on the actuator screw (7). (25) Adjust the rod end (9) as required to attain a measurement of 1.17 ± 0.03 inches between the center of the bearing in the rod end (9) and the top of the actuator screw (11) (Ref. Figure 211). (26) Engage the key washer (1) in the alignment slots in the top of the actuator screw (11). (27) Tighten nut (2) and install safety wire on the nut (2) and key washer (1).

C. Installation (UA-1 and After; UB-1 and After) WARNING: Perform all Steps of this procedure, in the order listed. Do not skip any Steps of this procedure. Failure to do so may result in injury to personnel and damage to equipment. CAUTION: Where the cables pass through structure, the areas of possible contact between the control cables and adjacent structure must be protected with protective elements such as grommets, rub strips, blocks or guide fairings. Where contact of control cables does occur with the protective elements, a force no greater than eight (8) ounces shall be required to move the cable to a position of no contact. At no time should flight control cables contact metal structure with the protective elements removed. If tension in the system is lost during cable maintenance, check the forward cable drum and aileron trim tab actuator cable drum to ensure the cable is not unwound from drums. (1) Install the trim tab actuator by performing the following Steps (Ref. Figure 205): (a) Mount the actuator on the left wing aft spar with the cables facing inboard. Install screw and washer (15) and bolts (14). NOTE: Grease (17, Table 2, 27-00-00) may be used. CAUTION: Do not mix greases. Mixing greases reduces lubricant effectiveness. The actuator was originally manufactured with grease (1, Table 2, 27-00-00) and this grease may have been cleaned out and replaced with grease (17, Table 2, 27-00-00). (2) Service the aileron trim tab actuator with grease (1, Table 2, 27-00-00). (3) Attach the actuator cable right-hand threads terminal end (8) to the feed line labeled “right-hand threads terminal end” (Ref. Figure 205, Detail F). (4) Attach the actuator cable left-hand threads terminal end (9) to the feed line labeled “left-hand threads terminal end”. NOTE: More than one person will be required to route the actuator cable. Take precautions to keep the cable clean and free from damage. (5) Route cable (9) from the trim tab actuator through the top pulley at WS 167, then to the turnbuckle in the left wheel well (Ref. Figure 205). (6) Route cable (8) from the trim tab actuator through the bottom pulley at WS 167, then to the turnbuckle in the left wheel well. (7) Install cable guard pins in the pulley brackets. Install grommets on wheel well ribs and cove ribs.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (8) Lubricate turnbuckles with grease (1, Table 2, 27-00-00) for corrosion protection prior to installation. NOTE: Each turnbuckle (3 and 10) barrel has a groove (5 and 12) at one end to identify the left-hand threaded end (Ref. Figure 203, Detail A, B and D). (9) Remove feed line, and attach the left-hand threads terminal end of cable (9) to the top turnbuckle (Ref. Figure 205). (10) Remove feed line, and attach the right-hand threads terminal end of cable (8) to the bottom turnbuckle. (11) Tension the actuator cable sufficient to prevent slack. (12) Ensure that the actuator cable is routed properly by verifying that the cable has been routed exactly as described in Steps (5) and (6). Ensure cable is engaged in the pulley grooves and all guard pins are installed. (13) Remove cable block (7) from both aileron trim tab forward cables (6) (Ref. Figure 203, Detail C). (14) Remove all tape from the cables (1 and 8) and turnbuckles (3 and 10) (Ref. Figure 203, Detail A and D). (15) Perform the following operational check: (a) Rotate the aileron trim tab control knob counterclockwise and verify the aileron trim tab actuator rod end retracts smoothly with no unusual noise or binding. (b) Rotate the aileron trim tab control knob clockwise and verify the aileron trim tab actuator rod end extends smoothly with no unusual noise or binding. (c) If these requirement are not met, verify all Steps of this procedure have been properly accomplished. NOTE: The double clevis end (4) on the trim tab actuator push-pull rod is designed to tighten the outer clevis with a binding action against the clevis bolt and the inner clevis, which removes freeplay of the bolt in the hole. Tighten the outer clevis last (the large jam nut (5)) after installing the clevis bolt (Ref. Figure 211). (16) Install the aileron trim tab push-pull rod assembly (8) on the aileron trim tab actuator (12) and install the bolt (10) washer and nut (3). (17) Tighten the large jam nut (5) on the double clevis (4). (18) Check the measurement of the push-pull rod assembly (8) for 8.58 ± 0.06 inches in length. (19) If no adjustment is required, proceed to Step (20). If adjustment is required, perform the following Steps: (a) Loosen small jam nut (13) on the double clevis (7). (b) Adjust double clevis (7) until measurement is obtained. (c) Tighten small jam nut (13) on the double clevis (7) and check measurement.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (20) Perform the left AILERON INSTALLATION (UA-1 AND AFTER; UB-1 AND AFTER) procedure (Ref. 27-10-00). (21) Perform the AILERON TRIM TAB OPERATIONAL CHECK procedure (Ref. 27-10-07). (22) Perform the AILERON TRIM TAB RIGGING procedure (Ref. 27-10-07). (23) Ensure turnbuckles (3) have been safety clipped (2) (Ref. Figure 203, Detail A and D). (24) Install lower left wing access panels 7, 35, 37, 43, 44 and 46 (Ref. Chapter 6-50-00, WING ACCESS PANELS). (25) Install upper left wing access panels 4 and 29 (Ref. Chapter 6-50-00, WING ACCESS PANELS). (26) Remove red tag from the aileron trim tab control knob.

D. Removal (UC-1 and After) CAUTION: If tension in the system is lost during cable maintenance, check the forward cable drum and aileron trim tab actuator cable drum to ensure the cable is not unwound from drums. (1) Connect external electrical power to the airplane. (2) Select the BATT switch to the ON position. (3) Select the EXT PWR switch to the EXT PWR position. (4) Lower flaps to the full down position. (5) Select the EXT PWR switch to the OFF position. (6) Select the BATT switch to the OFF position. (7) Disconnect external electrical power from the airplane. (8) Attach a red tag to the aileron trim tab control knob with the words “Do Not Operate, Maintenance In Progress”. (9) Perform the left AILERON REMOVAL (UC-1 AND AFTER) procedure (Ref. 27-10-00). (10) Remove the cove panel from the left wing trailing edge at the outboard flap. (11) Remove the trim tab actuator cover panel. (12) Remove lower left wing access panels 19 as required to access aileron trim tab cables, turnbuckles and stops (Ref. Chapter 6-50-00, WING ACCESS PANELS). (13) Adjust the aileron trim tab control knob to allow easy access to the turnbuckles (3 and 10) in the left wheel well and lower wing access panel 19 (Ref. Figure 203, Detail A and B). (14) Install a cable block (7) across both aileron trim tab forward cables (6) just outboard of WS 79.65 to prevent loss of trim tab forward cable tension (Ref. Figure 203, Detail C). (15) Remove safety clips (2 and 9) from the turnbuckles (3 and 10) (Ref. Figure 203, Detail A, B and D).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL NOTE: Each turnbuckle (3) barrel has a groove (5) at one end to identify the left-hand threaded end. (16) Attach a tag with the words “left-hand threads terminal end” to the outboard end of the top turnbuckle (3 or 10) (Ref. Detail A). (17) Disconnect the left-hand threads terminal end from the top turnbuckle (3 or 10) and attach a feed line to the terminal end. Label the feed line with the words “left-hand threads terminal end”. (18) Attach a tag with the words “right-hand threads terminal end” to the outboard end of the bottom turnbuckle (3) (Ref. Detail B). (19) Disconnect the right-hand threads terminal end from the bottom turnbuckle (3) and attach a feed line to the terminal end. Label the feed line with the words “right-hand threads terminal end”. NOTE: The cable guard pins for the pulleys at WS 184.4 and WS 172.4 are retained by cotter pins. (20) Remove the cable guard pins from pulley brackets. NOTE: The double clevis end (4) on the trim tab actuator push-pull rod (8) is designed to tighten the outer clevis with a binding action against the clevis bolt (10) and the inner clevis, which removes freeplay of the bolt in the hole. Loosen the outer clevis first (the larger nut (5)) before removing the clevis bolt (10) (Ref. Figure 212). (21) Loosen the large jam nut (5) on the double clevis. (22) Remove nut and washer (3) and bolt (10) and remove the aileron trim tab push-pull rod (8) assembly from the aileron trim tab actuator (1). (23) With assistance, feed the trim tab actuator cables through the wing and remove the cables through the actuator access opening. (24) Detach the feed lines from the trim tab cable terminal ends leaving the feed lines in place. (25) Remove the safety wire (15) from the trim tab actuator mount bolts (14) (Ref. Figure 206). (26) Remove the mount bolts (14) attaching the actuator (13) to the wing and remove actuator (13).

E. Drum Cable Replacement (UC-1 and After) (1) Perform the AILERON TRIM TAB ACTUATOR AND CABLE REMOVAL procedure in this section. NOTE: If the drum screw alignment pin cover (8) is damaged during removal a new one can be made from 0.032 inch thick 2024-T3 Alclad, 0.98 ± 0.03 by 1.5 ± 0.1 inches (Ref. Figure 210). (2) Remove the drum screw alignment pin cover (8) by removing the sealant and carefully prying it off the actuator housing (5). (3) Remove the drum screw alignment pin (7). (4) Remove the drum screw (9) from the actuator drum (4).

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NOTE: Note the length and position of bolts (3 and 15) for installation. (5) Remove the bolts (3 and 15), washer and nut (6) (three places) from the base plate (1) and actuator housing (5). (6) Separate the actuator base plate (1) from the actuator housing (5). (7) Remove shim (2), bearing (14) and drum (4). NOTE: If cable is to be reused, take care to keep the cable clean and free from damage. Coil cable loosely no tighter than ten inches in diameter. (8) Unwind the cable (16) from the drum (4). (9) Slide the cable keeper (17) from the drum (4) and remove cable (16). (10) If a used cable is to be installed, clean with solvent (2, Table 2, 27-00-00) and check for corrosion and damage. Replace cable if necessary. Dip the cable in corrosion preventive compound (4, Table 2, 27-00-00). Excess should be removed by wiping with a clean rag. (11) Attach a tag labeled “left-hand threads terminal end” to the actuator cable left-hand threads terminal end. (12) Attach a tag labeled “right-hand threads terminal end” to the actuator cable right-hand threads terminal end. (13) Wrap the aileron trim tab actuator cable on the cable drum as follows (Ref. Figure 208): CAUTION: Do not kink the cable while locating the middle of the actuator cable. Damage to the cable will occur. (a) Align the terminal ends of the forward cable and position keeper at the midpoint of the cable. With the right-hand threaded terminal end toward the end of the drum that faces the base plate, center the keeper and aileron trim tab cable in the slot of the drum. Cable ends must be even (Ref. Detail B). (b) From the lock pin, wrap each cable end 4 1/4 turns around the drum beginning with the outside grooves and work toward the middle of the drum (Ref. Detail C). With the drum wound, verify that the cable ends are still aligned. (c) Insert drum (4) into the actuator housing (5). Tape the aileron trim tab actuator cables together just outside of the actuator housing to prevent cable backlash at the drum (Ref. Figure 210). NOTE: Grease (17, Table 2, 27-00-00) may be used. CAUTION: Do not mix greases. Mixing greases reduces lubricant effectiveness. The actuator was originally manufactured with grease (1, Table 2, 27-00-00) and this grease may have been cleaned out and replaced with grease (17, Table 2, 27-00-00). (14) Lightly lubricate the drum screw threads (9) with grease (1, Table 2, 27-00-00). (15) Install the drum screw (9) completely into the actuator housing (5) and drum (4) to distribute the grease onto the mating threads.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (16) Remove the drum screw (9) and lubricate with grease again. Install into the drum (4). (17) Install bearing (14) and shim (2) on the drum (4). (18) Install the base plate (1) on the actuator housing (5). (19) Check for end play of the drum screw (9). If no adjustment is required, proceed to Step (20). If adjustment is required, perform the following Steps: (a) Replace shim (2). (b) Remove the new shims (2) laminations until all end play is removed and the drum (4) turns freely. (20) Install bolts (3 and 15), washer and nut (6) (three places). (21) With cable ends even, adjust the drum screw (2) to attain a measurement of 0.71 ± 0.06 inch from the top of the drum screw to the actuator housing (Ref. Figure 212). Install drum screw alignment pin (7) (Ref. Figure 210). (22) Verify the measurement from the top of the drum screw (2) to the center of the bearing in the rod end (9) is 1.17 ± 0.03 inches. If no adjustment is required, proceed to Step (23). If adjustment is required, perform the following Steps (Ref. Figure 12): (a) Remove safety wire, loosen nut (11) and disengage key washer (12) from the alignment slots in the drum screw (2). (b) Adjust the rod end (9) as required to attain the 1.17 ± 0.03 inches. (c) Engage the key washer (12) in the alignment slots in the top of the drum screw (2). (d) Tighten nut (11) and install safety wire. NOTE: If the drum screw alignment pin cover (8) was damaged during removal a new one can be made from 0.032 inch thick 2024-T3 Alclad sheet metal, 0.98 ± 0.03 by 1.5 ± 0.1 inches (Ref. Figure 210). (23) Install drum screw alignment pin cover (8) and seal with adhesive (18, Table 2, 27-00-00).

F. Installation (UC-1 and After) WARNING: Perform all Steps of this procedure, in the order listed. Do not skip any Steps of this procedure. Failure to do so may result in injury to personnel and damage to equipment. CAUTION: Where the cables pass through structure, the areas of possible contact between the control cables and adjacent structure must be protected with protective elements such as grommets, rub strips, blocks or guide fairings. Where contact of control cables does occur with the protective elements, a force no greater than eight (8) ounces shall be required to move the cable to a position of no contact. At no time should flight control cables contact metal structure with the protective elements removed. If tension in the system is lost during cable maintenance, check the forward cable drum and aileron trim tab actuator cable drum to ensure the cable is not unwound from drums.

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(1) Install the trim tab actuator by performing the following Steps (Ref. Figure 206): (a) Mount the actuator on the left wing aft spar with the cables facing inboard. (b) Install bolts (14) and safety wire (15). NOTE: Grease (17, Table 2, 27-00-00) may be used. CAUTION: Do not mix greases. Mixing greases reduces lubricant effectiveness. The actuator was originally manufactured with grease (1, Table 2, 27-00-00) and this grease may have been cleaned out and replaced with grease (17, Table 2, 27-00-00). (2) Service the aileron trim tab actuator with grease (1, Table 2, 27-00-00). (3) Attach the actuator cable right-hand threads terminal end (8) to the feed line labeled “right-hand threads terminal end” (Ref. Figure 206, Detail E). (4) Attach the actuator cable left-hand threads terminal end (9) to the feed line labeled “left-hand threads terminal end”. NOTE: More than one person will be required to route the actuator cable. Take precautions to keep the cable clean and free from damage. (5) With assistance route the cables through the wing as follows: NOTE: It is permissible to install cable guard pins as the cable is being routed. The guard pins for the pulleys at WS 184.4 and WS 172.4 are retained by cotter pins. It is permissible to install these guard pins with heads facing down for clearance purposes. (a) Route cable (8) through pulleys at WS 184.4 and WS 172.4 (Ref. Figure 206). (b) Route cable (9) through aileron cove area through the pulley at WS 165.7. (c) Route cable (9) through the top pulley at WS 134.5. (d) Route cable (8) through the bottom pulley at WS 134.5. (e) Install cable guard pins in the pulley brackets. Install grommets on wheel well ribs and cove ribs. (6) Lubricate turnbuckles with grease (1, Table 2, 27-00-00) for corrosion protection prior to installation. NOTE: Each turnbuckle (3) barrel has a groove (5) at one end to identify the left-hand threaded end (Ref. Figure 3, Detail A and B). (7) Remove feed line, and attach the left-hand threads terminal end of cable (9) to the top turnbuckle (Ref. Figure 206). (8) Remove feed line, and attach the right-hand threads terminal end of cable (8) to the bottom turnbuckle. (9) Tension the actuator cable sufficient to prevent slack.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (10) Ensure that the actuator cable is routed properly by verifying that the cable has been routed exactly as described in Step (5). Ensure cable is engaged in the pulley grooves and all guard pins are installed. (11) Remove cable block (7) from both aileron trim tab forward cables (6) (Ref. Figure 203, Detail C). (12) Remove all tape from the cables (6 or 11) and turnbuckles (3 or 10) (Ref. Figure 203, Detail A, B and D). (13) Perform the following operational check: (a) Rotate the aileron trim tab control knob counterclockwise and verify the aileron trim tab actuator rod end retracts smoothly with no unusual noise or binding. (b) Rotate the aileron trim tab control knob clockwise and verify the aileron trim tab actuator rod end extends smoothly with no unusual noise or binding. (c) If these requirement are not met, verify all Steps of this procedure have been properly accomplished. (14) Install the cove panel on the left wing trailing edge at the outboard flap. (15) Install the trim tab actuator cover panel. NOTE: The double clevis end (4) on the trim tab actuator push-pull rod is designed to tighten the outer clevis with a binding action against the clevis bolt and the inner clevis, which removes freeplay of the bolt in the hole. Tighten the outer clevis last (the large jam nut (5)) after installing the clevis bolt (10) (Ref. Figure 212). (16) Install the aileron trim tab push-pull rod (8) assembly on the aileron trim tab actuator (1) and install the bolt (10). (17) Install washer and nut (3). (18) Tighten the large jam nut (5) on the double clevis (4). (19) Check the measurement of the aileron trim tab push-pull rod (8) assembly for 8.92 ± 0.06 inches in length. (20) If no adjustment is required, proceed to the Step (20). If adjustment is required, perform the following Steps: (a) Loosen small jam nut (13) on the double clevis (7). (b) Adjust the double clevis (7) until measurement is obtained. (c) Tighten small jam nut (13) on double clevis (7) and check measurement. (21) Perform the left AILERON INSTALLATION (UC-1 AND AFTER) procedure (Ref. 27-10-00). (22) Perform the AILERON TRIM TAB OPERATIONAL CHECK procedure (Ref. 27-10-07). (23) Perform the AILERON TRIM TAB RIGGING procedure (Ref. 27-10-07).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (24) Ensure turnbuckles (3 and 10) have been safety clipped (2 and 9) (Ref. Figure 203, Detail A and B). (25) Install lower left wing access panels 19 as required to access aileron trim tab cables, turnbuckles and stops (Ref. Chapter 6-50-00, WING ACCESS PANELS). (26) Remove red tag from the aileron trim tab control knob.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

TAPER PIN DRUM SHAFT

B C D E

CABLE DRUM

A CABLE GUARD DETAIL

A FORWARD CABLE WITH LEFT - HAND THREADS TERMINAL END

MIDDLE OF CABLE AND MIDDLE OF CABLE DRUM SLOT

SIDE WITH RIGHT-HAND THREADS TERMINAL END

CABLE LOCK PIN CABLE LOCK PIN INSTALLED FLAT SIDE OF DRUM SIDE WITH LEFT - HAND THREADS TERMINAL END

FORWARD CABLE WITH RIGHT - HAND THREADS TERMINAL END

(CABLE GUARD NOT SHOWN) DETAIL

B

(DRUM NOT FULLY WRAPPED) (CABLE GUARD NOT SHOWN) DETAIL

C

CABLE WITH RIGHT-HAND THREADS TERMINAL END BEGINNING AT THE DRUM SHAFT SIDE OF THE DRUM

CABLE WITH LEFT-HAND THREADS TERMINAL END BEGINNING AT THE FLAT SIDE OF THE DRUM

CABLE GUARD FORWARD SIDE OF DRUM FORWARD CABLE RIGHT - HAND THREADS TERMINAL END

FORWARD CABLE WITH LEFT - HAND THREADS TERMINAL END

TAPE (TEMPORARY)

DETAIL

(DRUM FULLY WRAPPED)

E

DETAIL

D

THIS INFORMATION IS FOR THE MODEL 1900/1900C ONLY. DO NOT USE ON OTHER MODEL 1900 SERIES AIRCRAFT, BECAUSE THERE ARE CRITICAL DIFFERENCES.

Figure 201 Aileron Trim Tab Cable Winding

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UC27B 044491AD.AI

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

1 1. PEDESTAL 2. DRUM 3. BRACKET 4. CABLE GUARD 5. WASHERS

2

3

5

FWD

END PLAY 0.031 TO 0.063

4

VIEW LOOKING UP AT BOTTOM OF FWD AILERON TRIM TAB CABLE DRUM. UC27B 044833AB.AI

Figure 202 Aileron Trim Tab Cable Drum Installation

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1. AILERON TRIM TAB FUSELAGE CABLES 2. SAFETY CLIPS 3. TURNBUCKLES 4. CABLE BLOCKS 5. LEFT HAND THREAD GROOVE 6. AILERON TRIM TAB ACTUATOR CABLES

1 2

3

C

5 6

D

A B OUTBD

VIEW LOOKING UP INTO AFT WHEEL WELL DETAIL

A 2

3

6

4

5

OUTBD

VIEW LOOKING UP AT LEFT WING LOWER SURFACE (UC-1 AND AFTER) DETAIL

B

Figure 203 (Sheet 1 of 3) Aileron Trim Tab Cable Blocks and Turnbuckles

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UC27B 044608AB.AI

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

6. CABLES 7. CABLE BLOCK

7

6

OUTBD

VIEW LOOKING UP

DETAIL

C UC27B 045107AA.AI

Figure 203 (Sheet 2 of 3) Aileron Trim Tab Cable Blocks and Turnbuckles (UC-1 and After)

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8. AILERON TRIM TAB FUSELAGE CABLE 9. SAFETY CLIPS 10. TURNBUCKLES 11. LEFT HAND THREAD GROOVE 12. AILERON TRIM TAB ACTUATOR CABLES 13. CABLE BLOCK

10

13

11

FORWARD

8

9 12

INBOARD

VIEW LOOKING DOWN AT LEFT WING UPPER SURFACE DETAIL

D

Figure 203 (Sheet 3 of 3) Aileron Trim Tab Cable Blocks and Turnbuckles (UA-1 and After; UB-1 and After) Page 222 Nov 1/09

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UB27B 052173AA.AI

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

1. CABLE STOPS 2. SAFETY WIRE 3. CABLE STOP PLATE

A

1 3 1

2

2

OUTBD

VIEW LOOKING UP DETAIL

A

UC27B 043891AA.AI

Figure 204 Aileron Trim Tab Cable Stop Adjustment

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

1. DRUM SHAFT 2. CABLE DRUM 3. CABLE GUARD 4. TAPER PIN 5. AILERON TRIM TAB CONTROL SHAFT 6. DIAL INDICATOR 7. TRIM TAB CONTROL KNOB 8. CABLE 9. CABLE 10. AILERON

4 18 1 6

17

11. AILERON TRIM TAB 12. DOUBLE CLEVIS 13. ACTUATOR 14. MOUNT BOLT 15. SCREW AND WASHER 16. LARGE JAM NUT 17. WASHERS 18. UNIVERSAL JOINT 19. CABLE STOPS

2 7

5

3

AILERON TRIM TAB CONTROL ASSEMBLY DETAIL

CABLE STOPS INBD

A

DETAIL

C

CABLE PRESSURE SEALS DETAIL

B

9

A

19

9 FS 319

8

B

FS 320

TURNBUCKLES

R

BL 27

L

9

8 L

C 8

D

9

9

DETAIL

8 FS 105

12

16

8 WS 167 14

10 11

LEFT-HAND THREAD

D

R

F

15

9

WS 213

E

FS 106 13 13 AILERON TRIM TAB CONTROL DETAIL

E

RIGHT-HAND THREAD

8

14

14

VIEW LOOKING AFT DETAIL

UA27B 050079AC.AI

F

Figure 205 Aileron Control Trim Tab Control System (UA-1 and After; UB-1 and After)

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4 18 1

6 17

CABLE STOPS 2

7 3

DETAIL

5

C

INBD AILERON TRIM TAB CONTROL ASSEMBLY CABLE PRESSURE SEALS DETAIL

A

DETAIL

B TURNBUCKLES

9 DETAIL

D

19 FS 319

8 FS 320 BL 27

A

B

9

8

R

14

15

L

C

14

9

L R 8

D

E

WS 134.5 WS 165.7

9 9 FS 105

FS 106

WS 172.4

LEFT HAND THREAD

8

WS 184.4

F

WS 203

10 16

13

11

1. DRUM SHAFT 2. CABLE DRUM 3. CABLE GUARD 4. TAPER PIN 5. AILERON TRIM TAB CONTROL SHAFT 6. DIAL INDICATOR 7. TRIM TAB CONTROL KNOB 8. CABLE 9. CABLE

10. AILERON 11. AILERON TRIM TAB 12. DOUBLE CLEVIS 13. ACTUATOR 14. MOUNT BOLT 15. SAFETY WIRE 16. LARGE JAM NUT 17. WASHERS 18. UNIVERSAL JOINT 19. CABLE STOPS

8

RIGHT HAND THREAD 12

14

15

13

14 VIEW LOOKING FORWARD DETAIL

AILERON TRIM TAB CONTROL

DETAIL

F

UC27B 044490AC.AI

E

Figure 206 Aileron Trim Tab Control System (UC-1 and After)

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A B C D

FWD DETAIL

CABLE KEEPER (PART OF CABLE ASSEMBLY)

A

SIDE TOWARD ACTUATOR END CAP

CENTER OF CABLE SIDE TOWARD ROD END

SIDE WITH RIGHT-HAND THREADS TERMINAL END.

SIDE WITH LEFT-HAND THREADS TERMINAL END.

DETAIL

START WRAPPING OUTSIDE GROOVES MOVING INWARD 4 1/4 TURNS BOTH SIDES

B

SIDE WITH LEFT-HAND THREADS TERMINAL END.

SIDE WITH RIGHT-HAND THREADS TERMINAL END.

DETAIL

C

DRUM FULLY WRAPPED DETAIL

D

THIS INFORMATION IS FOR THE MODEL 1900/1900C ONLY. DO NOT USE ON OTHER MODEL 1900 SERIES AIRCRAFT, BECAUSE THERE ARE CRITICAL DIFFERENCES.

UA27B 050066AB.AI

Figure 207 Aileron Trim Tab Actuator Drum Winding (UA-1 and After; UB-1 and After)

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B C D

FWD DETAIL

CABLE KEEPER (PART OF CABLE ASSEMBLY)

A

A

SIDE TOWARD ACTUATOR BASE PLATE

CENTER OF CABLE SIDE TOWARD ACTUATOR CAP

SIDE WITH RIGHT-HAND THREADS TERMINAL END.

SIDE WITH LEFT-HAND THREADS TERMINAL END.

DETAIL

START WRAPPING OUTSIDE GROOVES MOVING INWARD 4 1/4 TURNS BOTH SIDES

B

SIDE WITH LEFT-HAND THREADS TERMINAL END.

SIDE WITH RIGHT-HAND THREADS TERMINAL END.

DETAIL

C

DRUM FULLY WRAPPED DETAIL

D

THIS INFORMATION IS FOR THE MODEL 1900/1900C ONLY. DO NOT USE ON OTHER MODEL 1900 SERIES AIRCRAFT, BECAUSE THERE ARE CRITICAL DIFFERENCES.

Figure 208 Aileron Trim Tab Actuator Drum Winding (UC-1 and After) Page 230 Nov 1/09

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UC27B 044704AB.AI

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

19. BEARING ADJUSTMENT RING 20. COTTER PIN 21. END CAP 22. BEARING 23. CABLE DRUM 24. BEARING 25. CABLE KEEPER (PART OF CABLE ASSEMBLY) 26. CABLE 27. WASHER (3 PLACES)

1. ROD END 2. JAM NUT 3. KEY WASHER 4. BUSHING 5. HOUSING HALF 6. SCREW 7. ACTUATOR SCREW 8. LUBRICATION FITTING 9. HOUSING HALF 10. WASHER 11. NUT 12. WASHER (3 PLACES) 13. NUT (3 PLACES) 14. BOLT 15. GUIDE 16. WASHER 17. NUT 18. BOLT (3 PLACES)

9

10

11

8

17

16

7 12

14 13 6 15

5

2

3

4

1

6

18

27

20 19

21 22

A

23

23

24 25

26

DETAIL

A UA27B 050065AA.AI

Figure 209 Aileron Trim Tab Actuator (UA-1 and After; UB-1 and After)

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1. BASE PLATE 2. SHIM 3. BOLTS (TWO PLACES) 4. DRUM 5. ACTUATOR HOUSING 6. WASHER, NUT (THREE PLACES) 7. DRUM SCREW ALIGNMENT PIN 8. DRUM SCREW ALIGNMENT PIN COVER 9. DRUM SCREW 10. KEY WASHER 11. NUT 12. ROD END 13. BEARING 14. BEARING 15. BOLT 16. CABLE 17. CABLE KEEPER (PART OF CABLE ASSEMBLY)

UC27B 044705AB.AI

Figure 210 Aileron Trim Tab Actuator (UC-1 and After) Page 232 Nov 1/09

27-10-05

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

1. KEY WASHER 2. NUT 3. NUT AND WASHER 4. DOUBLE CLEVIS 5. LARGE JAM NUT 6. LARGE JAM NUT 7. DOUBLE CLEVIS 8. PUSH-PULL ROD ASSEMBLY 9. ROD END 10. BOLT 11. ACTUATOR SCREW 12. ACTUATOR 13. SMALL JAM NUT

UA27B 050067AB.AI

Figure 211 Aileron Trim Tab Actuator and Push-Pull Rod Assembly (UA-1 and After; UB-1 and After)

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1. ACTUATOR 2. DRUM SCREW 3. WASHER AND NUT 4. DOUBLE CLEVIS 5. LARGE JAM NUT 6. LARGE JAM NUT 7. DOUBLE CLEVIS 8. PUSH-PULL ROD ASSEMBLY 9. ROD END 10. BOLT 11. JAM NUT 12. KEY WASHER 13. SMALL JAM NUT

UC27B 045249AB.AI

Figure 212 Aileron Trim Tab Actuator and Push-Pull Rod Assembly (UC-1 and After) Page 234 Nov 1/09

27-10-05

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

FLIGHT CONTROLS AILERON TRIM TAB INDICATOR MAINTENANCE PRACTICES

27-10-06 200200

1. AILERON TRIM TAB CONTROL A. Removal (1) Remove the right flight compartment pedestal side panel. (2) Position the aileron trim control dial (1) to 0 (Ref. Figure 201). (3) Loosen the setscrew (3) in the aileron trim control knob (2). (4) Remove the aileron trim control knob (2). (5) Remove the snap ring (9). (6) Pull up and remove the aileron trim control dial (1). (7) Remove the gear (4). (8) Working through the right side of the pedestal, remove the safety wire from the upper portion of the universal joint (7) securing the pin (6). (9) Remove the pin (6) from the universal joint (7). (10) Remove the shaft (5) from the pedestal.

B. Installation (1) Install the shaft (5) into the pedestal (Ref. Figure 201). (2) Working through the right side of the pedestal, install the upper pin (6) into the universal joint (7) and safety wire. (3) Install the gear (4). (4) Install the aileron trim control dial (1). (5) Install the snap ring (9). (6) Install the aileron trim control knob (2). Apply Loctite 242 to the setscrew (3) threads and tighten. (7) Perform the AILERON TRIM TAB INDICATOR ADJUSTMENT procedure in this section. (8) Install the right flight compartment pedestal side panel.

C. Inspection (1) Loosen the setscrew (3) in the aileron trim control knob (2). (2) Position the aileron trim tab control dial (1) to 0 (Ref. Figure 201). (3) Remove the aileron trim control knob (2).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (4) Remove the snap ring (9). (5) Pull up and remove the aileron trim control dial (1). (6) Inspect the gears for distortion and missing teeth. (7) Install the aileron trim control dial (1). (8) Install the snap ring (9). (9) Install the aileron trim control knob (2). Apply Loctite 242 to the setscrew (3) threads and tighten.

D. Indicator Adjustment (1) Remove the passenger seat(s) as required (Ref. Chapter 25-20-00, PASSENGER SEAT REMOVAL). (2) Remove the passenger compartment carpet as required (Ref. Chapter 25-20-01). (3) Remove the floor access panel 16 at FS 322 (Ref. Chapter 6-50-00, FLOOR ACCESS PANELS). (4) Install rig pin (1) (7, Table 1, Chapter 27-00-00) in the aileron quadrant support bracket (2) and the aileron quadrant (3) at FS 322 (Ref. Figure 202). (5) Install the aileron travel board (Ref. 27-00-02). (6) Verify that the left aileron is reading 0° on the travel board and that the trailing edge of the aileron and aileron trim tab are aligned. (7) Loosen the setscrew (3) in the aileron trim control knob (2) (Ref. Figure 201). (8) Remove the aileron trim control knob (2). (9) Remove the snap ring (9). (10) Pull the aileron trim control dial (1) up until it turns freely. (11) Turn the dial (1) to the desired position and lower to engage the gear (4). (12) Install the snap ring (9). (13) Install the aileron trim control knob (2). Apply Loctite 242 to the setscrew (3) threads and tighten. (14) Remove the rig pin (1) from the aileron quadrant (3) (Ref. Figure 202). (15) Install the floor access panel 16 at FS 322 (Ref. Chapter 6-50-00, FLOOR ACCESS PANELS). (16) Install the passenger compartment carpet (Ref. Chapter 25-20-01). (17) Install the passenger seat(s) (Ref. Chapter 25-20-00, PASSENGER SEAT INSTALLATION). (18) Remove the travel board (Ref. 27-00-02).

Page 202 Nov 1/09

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

1. AILERON TRIM CONTROL DIAL 2. AILERON TRIM CONTROL KNOB 3. SETSCREW 4. GEAR 5. GEAR SHAFT 6. PIN 7. UNIVERSAL JOINT 8. SHAFT 9. SNAP RING

A

1

9

2

8 7

6 5

DETAIL

4

A 3

UC27B 041807AA.AI

Figure 201 Aileron Trim Tab Control

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

1. RIG PIN 2. AILERON QUADRANT SUPPORT BRACKET 3. AILERON QUADRANT

A 1

3

2

AILERON QUADRANT

DETAIL

A UC27B 041762AB.AI

Figure 202 Aileron Quadrant Rig Pin Installation

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

FLIGHT CONTROLS AILERON TRIM TAB CONTROL SYSTEM MAINTENANCE PRACTICES

27-10-07 200200

1. AILERON TRIM TAB A. Rigging WARNING: Perform all Steps of this procedure, in the order listed. Do not skip any Steps of this procedure. Failure to do so may result in injury to personnel and damage to equipment. Before rigging the aileron trim tab, the ailerons must be properly rigged as described in AILERON CONTROL SYSTEM RIGGING. Refer to Chapter 27-10-03. Failure to do so may result in injury to personnel and damage to equipment. CAUTION: Where the cables pass through structure, the areas of possible contact between the control cables and adjacent structure must be protected with protective elements such as grommets, rub strips, blocks or guide fairings. Where contact of control cables does occur with the protective elements, a force no greater than eight (8) ounces shall be required to move the cable to a position of no contact. At no time should flight control cables contact metal structure with the protective elements removed. (1) UA-1 and After; UB-1 and After perform the following Steps: (a) Remove left lower wing access panels 7 and 44 (Ref. Chapter 6-50-00, WING ACCESS PANELS). (b) Remove the left side upper wing access panel 4 (Ref. Chapter 6-50-00, WING ACCESS PANELS). (c) Install rig pin (3) (11, Table 1, 27-00-00) through the aileron outboard wing bellcrank (1) (Ref. Figure 205). (2) UC-1 and After perform the following Steps: (a) Remove the passenger seat(s) as required to gain access to floor access panel 16E (Ref. Chapter 25-20-00, SEAT REMOVAL). (b) Remove the carpet as required to gain access to floor access panel 16E (Ref. Chapter 25-20-01). (c) Remove floor access panel 16E (Ref. Chapter 6-50-00, FLOOR ACCESS PANELS). (d) Remove left side lower wing access panel 19 (Ref. Chapter 6-50-00, WING ACCESS PANELS). (e) Install rig pin (1) (7, Table 1, 27-00-00) through the aileron quadrant support bracket (2) and the aileron quadrant (3) (Ref. Figure 202). (3) Perform the AILERON TRAVEL BOARD INSTALLATION procedure (Ref. 27-00-02). (4) Perform the AILERON TRIM TAB TRAVEL BOARD INSTALLATION procedure (Ref. 27-00-02).

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(5) Check the cable tension by performing the following Steps: NOTE: Do not measure the tension when the temperature is varying rapidly or with the airplane located in direct sunlight. (a) Allow the temperature of the airframe to stabilize before measuring and adjusting the cable tension. (b) Measure the temperature in the compartment with the aileron trim tab cables in the lower wing area. (c) Refer to Aileron Trim Tab Cable Tension Graph Figure 201, and read the pounds of tension required for the measured temperature. (d) Rotate the aileron trim tab control knob to the left and right three cycles to equalize system tension and return to the 0 position on the dial indicator. NOTE: Cable tension tolerance is + 3 / - 2 pounds of the tension found in Figure 201. (e) Position a cable tensiometer (4, Table 1, 27-00-00) on the aileron trim tab cables at least three inches from the turnbuckles and pulleys and measure the cable tension of both cables. Cable diameter is noted in Figure 201. (f) If no adjustment is required, proceed to Step (6). If adjustment is required, perform the following Steps: 1 Remove the safety clips (2) from the aileron trim tab cable turnbuckles (3) (Ref. Figure 203). WARNING: If cable tension at any time is below 5 pounds, check all aileron trim system cable drums and pulleys for proper cable engagement. NOTE: Each turnbuckle (3) barrel has a groove (5) at one end to identify the left-hand threaded end. 2 Adjust the turnbuckles (3) until both cables (1) have equal tensions needed at the current temperature found in Figure 201. Cable diameter is noted in Figure 201. 3 Rotate the aileron trim tab control knob to the left and right three cycles to equalize system tension and return to the 0 position on the dial indicator and check the cable tension. If tension is out of limits, repeat Step (5) (f) 2. 4 Install the safety clips (2) on the aileron trim tab cable turnbuckles (3). (6) Using the aileron trim tab travel board, check the aileron trim tab for a deflection of 15° ± 1.5° up and 15° ± 1.5° down from the 0° position on the travel board and ensure that the cable stops (1) contact the stop plate (3) at full deflection (Ref. Figure 204). (a) If no adjustment is required, proceed to Step (7). If adjustment is required, perform the following Steps: 1 Remove safety wire (2) from the cable stop(s) (1) located in the aft left main landing gear area.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL 2 Loosen the cable stop(s) (1) and adjust as needed to achieve proper deflection. NOTE: Wrap five turns on each end of safety wire opposite of the cable stop contact area. 3 Tighten cable stop(s) (1) and torque to 40-45 inch-pounds and install safety wire (2). (7) Perform a visual check to make sure the aileron trim tab movement corresponds to the movement indicated on the trim tab indicator: (a) Rotate the aileron trim tab control knob counterclockwise and make sure that the aileron trim tab moves down smoothly with no unusual noise or binding. (b) Rotate the aileron trim tab control knob clockwise and make sure that the aileron trim tab moves up smoothly with no unusual noise or binding. (8) Rotate the trim tab control knob so that the trim tab reads 0° on the aileron trim tab travel board and verify that the zero on the trim tab Indicator aligns with the arrow on the pedestal. If no adjustment is required proceed to Step (9). If adjustment is required, perform the AILERON TRIM TAB INDICATOR ADJUSTMENT procedure (Ref. 27-10-06). NOTE: With the aileron trim tab set at neutral, ± 0.5° of servo travel is permissible with full up and down travel of the aileron. Servo effect shall not cause the trim tab to exceed its maximum travel setting at full-up or full-down trim tab and full-up or full-down aileron. (9) Check aileron trim tab servo travel as follows: (a) Rotate the trim tab control knob so that the trim tab reads 0° on the aileron trim tab travel board. (b) Remove rig pin (1) (7, Table 1, 27-00-00) from the aileron quadrant (3) (UC-1 and After) (Ref. Figure 202). (c) Remove rig pin (3) (11, Table 1, 27-00-00) from the aileron outboard wing bellcrank (UA-1 and After; UB-1 and After) (Ref. Figure 205). (d) Rotate either control wheel counterclockwise to the full left position. Ensure the trim tab servo travel is no more than ± 0.5°. (e) Rotate either control wheel clockwise to the full right position. Ensure the trim tab servo travel is no more than ± 0.5°. (f) Rotate the trim tab control knob counterclockwise to the full left position. (g) Rotate either control wheel counterclockwise to the full left position and verify the aileron trim tab does not exceed 16.5° on the aileron trim tab travel board. (h) Rotate either control wheel clockwise to the full right position and verify that the aileron trim tab does not exceed 16.5° on the aileron trim tab travel board. (i) Rotate the trim tab control knob clockwise to the full right position.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (j) Rotate either control wheel counterclockwise to the full left position and verify the aileron trim tab does not exceed 16.5° on the aileron trim tab travel board. (k) Rotate either control wheel clockwise to the full right position and verify that the aileron trim tab does not exceed 16.5° on the aileron trim tab travel board. (l) If any of the above requirements are not met, Install the aileron quadrant rig pin and repeat Steps (6) thru (9). If further adjustments are required, perform the AILERON TRIM TAB ACTUATOR AND CABLE REMOVAL/ INSTALLATION procedures (Ref. Chapter 27-10-05). (10) Remove the aileron travel board. (11) Remove the aileron trim tab travel board. (12) UC-1 and After perform the following Steps: (a) Install floor access panel 16E (Ref. Chapter 6-50-00, FLOOR ACCESS PANELS). (b) Install the carpet as required (Ref. Chapter 25-20-01). (c) Install the passenger seat(s) as required (Ref. Chapter 25-20-00, SEAT INSTALLATION). (d) Install left side lower wing access panel 19 (Ref. Chapter 6-50-00, WING ACCESS PANELS). (13) UA-1 and After; UB-1 and After perform the following Steps: (a) Install left and right lower wing panels 44 and 45 (Ref. Chapter 6-50-00, WING ACCESS PANELS). (b) Install left side upper wing access panel 4 (Ref. Chapter 6-50-00, WING ACCESS PANELS).

B. Operational Check (1) Rotate the aileron trim tab control knob counterclockwise and make sure that the aileron trim tab moves down smoothly with no unusual noise or binding. (2) Rotate the aileron trim tab control knob clockwise and make sure that the aileron trim tab moves up smoothly with no unusual noise or binding. (3) If requirements are not met, perform the AILERON TRIM TAB RIGGING procedure in this section.

C. Cable Tension Check (1) Remove lower left wing access panels 7 (UA-1 and After, UB-1 and After), or 17 (UC-1 and After) outboard of the left main landing gear (Ref. Chapter 6-50-00, WING ACCESS PANELS). (2) Rotate the aileron trim tab control knob to the left and right position three cycles to equalize system tension and return the indicator dial to 0. (3) Check the cable tension by performing the following Steps: NOTE: Do not measure the tension when the temperature is varying rapidly or with the airplane located in direct sunlight.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (a) Allow the temperature of the airframe to stabilize before measuring and adjusting the cable tension. (b) Measure the temperature in the compartment with the aileron trim tab control cables in the lower wing area. (c) Refer to Aileron Trim Tab Cable Tension Graph Figure 201 and read pounds of tension required for the measured temperature. NOTE: Trim tab cable tension tolerance is + 3 / - 2 pounds. (d) Position a cable tensiometer (4, Table 1, 27-00-00) on the aileron trim tab cables at least three inches from turnbuckles and pulleys and measure the cable tension. Cable diameter is noted in Figure 201. (e) If no adjustment is required, proceed to Step (4). If adjustment is required, perform the AILERON TRIM TAB RIGGING procedure in this section. (4) Install lower left wing access panels 7 (UA-1 and After, UB-1 and After) or 17 (UC-1 and After) (Ref. Chapter 6-50-00, WING ACCESS PANELS). (5) Perform the AILERON TRIM TAB OPERATIONAL CHECK procedure in this section.

D. Functional Check (1) UA-1 and After; UB-1 and After perform the following Steps: (a) Remove left lower wing access panel 44 (Ref. Chapter 6-50-00, WING ACCESS PANELS). (b) Install rig pin (3) (11, Table 1, 27-00-00) through the aileron outboard wing bellcrank (1) (Ref. Figure 205). (2) UC-1 and After perform the following Steps: (a) Remove the passenger seat(s) as required to gain access to floor access panel 16E (Ref. Chapter 25-20-00, SEAT REMOVAL). (b) Remove the carpet as required to gain access to floor access panel 16E (Ref. Chapter 25-20-01). (c) Remove floor access panel 16E (Ref. Chapter 6-50-00, FLOOR ACCESS PANELS). (d) Install rig pin (1) (7, Table 1, 27-00-00) through the aileron quadrant support bracket (2) and the aileron quadrant (3) (Ref. Figure 202). (3) Perform the AILERON TRAVEL BOARD INSTALLATION procedure (Ref. 27-00-02). Verify that the aileron is at 0°. If the aileron is not at 0°, perform the AILERON CONTROL SYSTEM RIGGING procedure (Ref. 27-10-03). (4) Perform the AILERON TRIM TAB TRAVEL BOARD INSTALLATION procedure (Ref. 27-00-02). (5) Rotate the aileron trim tab control knob counterclockwise to the full left position and make sure that the trim tab moves to the full down position 13.5° to 16.5° smoothly with no unusual noise or binding.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (6) Rotate the trim tab control knob clockwise to the full right position and make sure that the trim tab moves to the full up position 13.5° to 16.5° smoothly with no unusual noise or binding. (7) If aileron trim tab requires adjustment, perform the AILERON TRIM TAB RIGGING procedure in this section. (8) UC-1 and After perform the following Steps: (a) Remove the rig pin (1) from the aileron quadrant (3) (Ref. Figure 202). (b) Install floor access panel 16E (Ref. Chapter 6-50-00, FLOOR ACCESS PANELS). (c) Install the carpet as required (Ref. Chapter 25-20-01). (d) Install the passenger seat(s) as required (Ref. Chapter 25-20-00, SEAT INSTALLATION). (9) UA-1 and After; UB-1 and After perform the following Steps: (a) Remove rig pin (3) (11, Table 1, 27-00-00) from the aileron outboard wing bellcrank (1) (Ref. Figure 205). (b) Install left lower wing panel 44 (Ref. Chapter 6-50-00, WING ACCESS PANELS). (10) Remove the aileron trim tab travel board. (11) Remove the aileron travel board.

TOLERANCE: +3, -2 POUNDS OF TENSION

1/16" DIAMETER AILERON TAB CABLE TENSION GRAPH

POUNDS OF TENSION

40

30

20

10

0 UC27B 043433AA.AI

Figure 201 Aileron Trim Tab Cable Tension Graph

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1. RIG PIN 2. AILERON QUADRANT SUPPORT BRACKET 3. AILERON QUADRANT

A 1

3

2

AILERON QUADRANT

DETAIL

A UC27B 041762AB.AI

Figure 202 Aileron Quadrant Rig Pin Installation

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1. AILERON TRIM TAB FUSELAGE CABLES 2. SAFETY CLIPS 3. TURNBUCKLES 4. CABLE BLOCKS 5. LEFT HAND THREAD GROOVE 6. AILERON TRIM TAB ACTUATOR CABLES

1 2

3

C

5 6

D

A B OUTBD

VIEW LOOKING UP INTO AFT WHEEL WELL DETAIL

A 2

3

6

4

5

OUTBD

VIEW LOOKING UP AT LEFT WING LOWER SURFACE (UC-1 AND AFTER) DETAIL

B

Figure 203 Aileron Trim Tab Turnbuckle Adjustment

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UC27B 044608AB.AI

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

1. CABLE STOPS 2. SAFETY WIRE 3. CABLE STOP PLATE

A

1 3 1

2

2

OUTBD

VIEW LOOKING UP DETAIL

A

UC27B 043891AA.AI

Figure 204 Aileron Trim Tab Cable Stop Adjustment

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1. BELLCRANK 2. LOWER SUPPORT BRACKET 3. RIG PIN 4. UPPER SUPPORT BRACKET

4 1 2

A

3

VIEW LOOKING AFT AND UP DETAIL

B UA27B 051940AA.AI

Figure 205 Rig Pin Installation (UA-1 and After; UB-1 and After)

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FLIGHT CONTROLS AILERON BALANCE WEIGHTS MAINTENANCE PRACTICES

27-10-08 200200

1. PROCEDURES (UC-1 AND AFTER) WARNING: Whenever any part of this system is dismantled, adjusted, repaired or renewed, detailed investigation must be made on completion to make sure that distortion, tools, rags or any other loose articles or foreign matter that could impede the free movement and safe operation of the system are not present, and that the systems and installations in the work area are clean.

A. Clip Inspection (1) Perform REMOVING GROUND POWER procedures (Ref. Chapter 24-40-00). (2) Perform BATTERY DISCONNECTION procedure (Ref. Chapter 24-31-00). (3) Perform AILERON REMOVAL procedure on both the right and left ailerons (Ref. 27-10-00). (4) Gain access to the tooling hole in Rib Station 15.06 (P/N 118-130000-67) by removing and retaining the existing P/N MS27039-1-XX screws (4 places), the P/N 118-130000-127 cover, and the P/N 118-130000-129 balance weight(s) (Ref. Figure 201). For airplanes in compliance with MSB 27-3928 proceed to Step (8). For airplanes not in compliance with MSB 27-3928 complete the following Steps. CAUTION: Tooling holes at Rib Stations 67.91 and 121.53 need to be enlarged a minimum 0.13-inch aft of center to prevent damage to aileron balance weight clips (Ref. Figures 203 and 204 for dimensions). (5) Enlarge one tooling hole (for borescope inspection) in Rib Stations 15.06 (P/N 118-130000-67), 65.05 (P/N 118-130000-75), 67.91 (P/N 118-130000-77), and 121.53 (P/N 118-130000-95) to 0.75 inch diameter (Ref. Figures 201 through 204 for locations of tooling holes). WARNING: Solvents, primers, and paints are flammable and toxic to skin, eyes, and respiratory tract. Skin and eye protection is required. Avoid repeated or prolonged contact. Keep away from flames or sources of heat. Use in a well ventilated area or respiratory protection equipment may be required. (6) Deburr edges of holes and clean up drilling debris using a vacuum cleaner or cheesecloth (191, Table 1, Chapter 91-00-00) and solvent (8 or 14, Table 2, 27-00-00 or 24, Table 1, Chapter 91-00-00) to ensure work areas are clean. WARNING: Chemical conversion coating (11, Table 2, 27-00-00) may affect skin, eyes, and respiratory tract. Chemical goggles and neoprene gloves will be worn. Use in a well ventilated area. (7) Protect all bare metal surfaces using appropriate chemical conversion coating (11, Table 2, 27-00-00) and Epoxy Polyamide Primer (16, Table 2, 27-00-00).

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(8) Fabricate a guide tube from locally procured material to prevent the fiber-optic probe from flexing when gaining access to the middle rib(s). WARNING: Be sure guide tube is removed when fiber-optic probe is removed. (9) Borescope both ailerons, paying particular attention to P/Ns 101-130001-191, -193 and P/Ns 118-130000-123, -124 aileron balance weight clips (Ref. Figure 205 for locations of clips) (Ref. Figure 206 for views of clips with and without cracks). (a) If no discrepancy exists, apply aluminum foil tape (187, Table 1, Chapter 91-00-00) over the oversized holes and proceed to Step (10). (b) If a discrepancy exists, contact Hawker Beechcraft Technical Support at 1-800-429-5372 or 316-676-3140. NOTE: Information Only Field Service Kit 129-4036-0001 provides information to fabricate clips for aileron balance weight installation, and can be used to replace one, any, or all clips used to attach the balance weight to the ribs. Any repair, modification, painting or replacement of the aileron or aileron trim tab requires balancing. (10) Perform the AILERON CHECKING BALANCE procedure (Ref. Chapter 57-50-00). NOTE: Length of P/N MS27039-1-XX screws used to attach P/N 118-130000-127 cover and P/N 118-130000-129 balance weights at Rib Station 15.06 is determined by number of balance weights. (11) Perform AILERON INSTALLATION procedure on both the right and left ailerons (Ref. 27-10-00). (12) Perform AILERON FUNCTIONAL CHECK procedure (Ref. 27-10-03). (13) Perform AILERON TRIM TAB FUNCTIONAL CHECK procedure (Ref. 27-10-07). (14) Perform AILERON FREEPLAY CHECK procedure (Ref. 27-10-00). (15) Perform BATTERY CONNECTION procedure (Ref. Chapter 24-31-00). (16) Ensure all work areas are clean and clear of tools and miscellaneous items of equipment. (17) Return airplane to service.

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Hawker Beechcraft Corporation MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

FORWARD

•

RIB STATION 15.06

PIN 118-130000-67 RIB

I

PIN MS27039-1-XX SCREWS (4 PLACES) NOTE: SCREW LENGTH DETERMINED BY NUMBER OF BALANCE WEIGHTS

FORWARD

--t---LOCATION OF ENLARGED HOLE TO 0.75 INCH DIAMETER

•

--CL

PIN 118-130000-127 COVER AND PIN 118-130000-129 BALANCE WEIGHT(S)

LEFT SIDE SHOWN, RIGHT SIDE OPPOSITE DETAIL

A

UC27B

094425M.AI

Figure 201 Location of Tooling Hole in Rib Station 15.06

27-10-08

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Hawker Beechcraft Corporation MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

~

FORWARD

PIN 118-130000-75 RIB RIVETS

----=

FORWARD

---1--

CL - -

---~--=\ LOCATION OF ENLARGED HOLE TO 0.75 INCH DIAMETER

RIB STATION 65.05

LEFT SIDE SHOWN, RIGHT SIDE OPPOSITE DETAIL

A

Figure 202 Location of Tooling Hole in Rib Station 65.05

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UC27B 094427M.AI

Hawker Beechcraft Corporation MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

FORWARD

~

PIN 118-130000-77 RIB RIB STATION 67.91

RIVETS

-1 1--- ---I

0.13 INCH MINIMUM AFT OF CENTER

1

I

~

--

CL

I

1

I

FORWARD

-=LOCATION OF ENLARGED HOLE TO 0.75 INCH DIAMETER LEFT SIDE SHOWN, RIGHT SIDE OPPOSITE DETAIL

A

UC27B

09442BAA.AI

Figure 203 Location of Tooling Hole in Rib Station 67.91

27-10-08

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Hawker Beechcraft Corporation MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

. _ FORWARD

LOCATION OF ENLARGED HOLE TO 0.75 INCH DIAMETER TOOLING HOLE

PIN 118-130000-95 RIB

----~-

CL--

II

II ~

0.13 INCH MINIMUM AFT OF CENTER

----~

FORWARD RIVETS

LEFT SIDE SHOWN, RIGHT SIDE OPPOSITE DETAIL

A

Figure 204 Location of Tooling Hole in Rib Station 121.53

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UC27B

094429M.AI

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

P/N 118-130000-123 CLIP (LEFT) P/N 118-130000-124 CLIP (RIGHT)

RIB STATION 15.06

RIB STATION 27.03 P/N 101-130001-191 CLIP (4 PLACES)

RIB STATION 60.33

AILERON BALANCE WEIGHTS RIB STATION 42.53

RIB STATION 67.91

VIEW OF AILERON, INBOARD SIDE

P/N 101-130001-193 CLIP RIB STATION 79.95

FORWARD RIB STATION 65.05

P/N 101-130001-191 CLIP (3 PLACES)

RIB STATION 121.53 AILERON BALANCE WEIGHT

VIEW OF AILERON, OUTBOARD SIDE

RIB STATION 101.50

LEFT AILERON SHOWN, RIGHT AILERON OPPOSITE

UC27B 094430AA.AI

Figure 205 Location of Aileron Balance Weight Clips

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Hawker Beechcraft Corporation MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

UNCRACKED CLIP

CRACKED CLIP

UC27B

094431M.AI

Figure 206 View of Aileron Balance Weight Clip With and Without Cracks

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

FLIGHT CONTROLS RUDDER MAINTENANCE PRACTICES

27-20-00 200200

1. PROCEDURES A. Removal (1) Attach a red tag to the rudder pedals with the words “Do Not Operate, Maintenance In Progress”. (2) Remove the rudder static wicks. (3) Remove the tail cone. (4) Remove nut (18), washer (17) and bolt (19) securing push-pull rod (20) to the rudder control horn (14) (Ref. Figure 201, Detail D). (5) Loosen nut (8) securing tightener (2) of the rudder trim actuator (1) push-pull rod (9). Remove cotter pin (7), nut (6), washer (5) and bolt (3) securing clevis (4) to the rudder trim tab horn (Ref. Figure 201, Detail B). (6) Remove the screw, washers and nuts securing the two bonding jumpers to the vertical stabilizer (26) structure in the area of the rudder clevis (25) (Ref. Figure 201, Detail F). (7) Remove three nuts (16), washers (10 and 15) and bolts (11) which attach rudder torque tube (12) to the rudder control horn (14) (Ref. Figure 201, Detail D). NOTE: Observe the number and position of the shims (13) for installation. CAUTION: Do not force the rudder surface to the left or right to access the mounting bolts. Damage to the rudder will occur. (8) With assistance support the rudder (21) and remove four bolts (22) and eight washers (23 and 24) from the left upper and left lower rudder clevis (25). Carefully move rudder to access the right side and remove four bolts (22) and eight washers (23 and 24) from the right upper and right lower rudder clevis (25) (Ref. Figure 201, Detail E and F). (9) Remove the rudder from the airplane.

B. Installation NOTE: Any repair, modification, painting or replacement of the rudder or the rudder trim tab requires balancing (Ref. Chapter 55-40-00). If a new rudder surface is being installed, remove the clevises from the new rudder and replace the old clevises on the airplane with the new clevises. (1) Prepare the area where the bonding cables attach for electrical bonding (Ref. Chapter 20-00-01, PREPARATION OF SURFACE). (2) Position the laminated shims (13) on the rudder control horn (14) (Ref. Figure 201, Detail D). (3) Route a feed line through the rudder and attach the feed line to the rudder trim actuator push-pull rod (9) (Ref. Figure 201, Detail B).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (2) Prepare the area where the bonding cables attach for electrical bonding (Ref. Chapter 20-00-01, PREPARATION OF SURFACE). (3) Position the laminated shims (13) on the rudder control horn (14) (Ref. Figure 201, Detail D). (4) Route a feed line through the rudder and attach the feed line to the rudder trim actuator push-pull rod (9) (Ref. Figure 201, Detail B). CAUTION: Do not force the rudder surface to the left or right to access the mounting bolts. Damage to the rudder will occur. (5) With the feed line, carefully guide the rudder trim actuator (1) push-pull rod (9) through the rudder while carefully placing the rudder torque tube (12) in position on the rudder control horn (14) ensuring that the laminated shims (13) are in place. (6) Loosely install bolts (22) and washers (23 and 24) that secure upper and lower rudder clevis (25) to the rudder structure (Ref. Figure 201, Detail G). (7) Loosely install bolts (11) and washers (10) to maintain alignment of torque tube (12), laminated shim (13) and rudder control horn (14). (8) Tighten bolts (22) and washers (23 and 24) that secure the upper and lower rudder clevis (25) to the rudder structure. Visually check for alignment between clevises (25) and hinges (27) (Ref. Figure 201). (9) Check for a snug fit between the rudder control horn (14) and the rudder torque tube (12) (Ref. Figure 201, Detail D). (10) If no adjustment is required, proceed to Step (10). If adjustment is required, perform the following Steps: (a) Remove bolts (11) and washers (10). (b) Remove bolts (22) and washers (23 and 24) that secure the upper and lower rudder clevis (25) to the rudder (21) structure. NOTE: Any combination of shims (13) with 0.003 or 0.062 inch laminations may be used. The maximum total thickness for the shim material should not exceed 0.190 inch. (c) Add or remove shims (13) as required to obtain a snug fit between the rudder control horn (14) and the rudder torque tube (12) and maintain alignment between clevises (25) and hinges (27). (d) Repeat Steps (4) thru (9). (11) Install the bolts (11), washers (10 and 15) and nuts (16) which attach the rudder torque tube (12) to the rudder control horn (14). Torque the nuts 50 to 70 inch-pounds. Apply torque paint to the threads and nuts (Ref. Figure 201, Detail D). NOTE: After torqueing, the bolts (11) shall not rotate in their holes under 24 to 35 inch-pounds of torque applied to the bolt. (12) Torque the bolts (22) that secure the upper and lower rudder clevis (25) to the rudder (21) structure 50 to 70 inch-pounds (Ref. Figure 201, Detail F).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (13) Position push-pull rod (20) into rudder control horn (14). Install bolt (19), washer (17) and nut (18) (Ref. Figure 201, Detail D). (14) Install the tail cone. (15) Remove the red tag from the rudder pedals. (16) Perform the RUDDER SYSTEM FUNCTIONAL CHECK procedure (Ref. 27-20-02). (17) Perform the RUDDER TAB SYSTEM FUNCTIONAL CHECK procedure (Ref. 27-20-07). (18) Verify that the installation is within the following limitations: (a) A gap of 0.10 ± 0.06 inch should exist between the trailing edge skin of the vertical stabilizer and the rudder nose with the rudder deflected to full travel towards that edge. (b) The minimum gap between the top edge of the rudder and the aft fairing bullet should be 0.12 inch. (c) The gap between the lower edge of the rudder and the upper edge of the tail cone should be from 0.12 to 0.38 inch. NOTE: It is permissible to trim the edge of the rudder or trim tab 0.06 inch in order to achieve the distances noted in Step (c). If trimming is required, the rudder should be checked for balance. (d) With full elevator down and elevator trim tab at trailing edge full down, check that there is clearance between the elevator trim tabs and the top of the rudder at full left and full right rudder. Permissible to trim a notch from the top trailing edge of the rudder (3) not to exceed 0.50 X 0.25 inch. (Ref. Figure 204, Detail A). (19) Install rudder static wicks.

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D A C

DETAIL

15. WASHER 16. NUT 17. WASHER 18. NUT 19. BOLT 20. PUSH-PULL ROD 21. RUDDER 22. BOLT 23. WASHER 24. WASHER 25. CLEVIS 26. VERTICAL STABILIZER 27. HINGE ASSEMBLY

1. RUDDER TRIM ACTUATOR 2. TIGHTENER 3. BOLT 4. CLEVIS 5. WASHER 6. NUT 7. COTTER PIN 8. NUT 9. PUSH-PULL ROD 10. WASHER 11. BOLT 12. RUDDER TORQUE TUBE 13. SHIM 14. RUDDER CONTROL HORN

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UC27B 044849AA.AI

Figure 201 Rudder Installation

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C. Freeplay Check NOTE: Movement or jarring of the airplane will invalidate the rudder freeplay readings. The airplane should be placed in a hangar and no personnel in or on the airplane during the freeplay check. (1) Visually inspect the rudder for any damage, security of the hinge attach points and for tightness of the actuating system. (2) Perform the RUDDER TRAVEL BOARD INSTALLATION procedure (Ref. 27-00-02). (3) Remove the aft fuselage access panel 8 (Ref. Chapter 6-50-00, FUSELAGE ACCESS PANELS). (4) Install rig pin (2) (7, Table 1, 27-00-00) in the rudder aft torque shaft (1) (Ref. Figure 202). NOTE: The double clevis ends on the tab actuator push-pull rod is designed to tighten the outer clevis with a binding action against the clevis bolt and the inner clevis, which removes freeplay of the bolt in the hole. Loosen the outer clevis first (the larger nut) before removing the clevis bolt. (5) Disconnect the rudder tab actuator push-pull rod (do not change adjustment of the rod) and align the tab trailing edge with the rudder trailing edge. Secure the rudder tab surface to the rudder surface. (6) Attach a scale (8) or dial indicator (1) to the travel board so the left and right movement can be measured at the trailing edge (Ref. Figure 203). (7) Apply a small piece of masking tape (for paint protection) on the right hand side of the rudder 1 inch forward of the rudder trailing edge just above the rudder trim tab (9). This will be the point of pressure against the rudder (3) by the push-pull scale. (8) Apply another piece of masking tape in the corresponding position on the left side of the rudder for the same purpose. (9) Position the dial indicator (1) so the stem (2) is 0.50 inch forward of the trailing edge of the rudder and is depressed 0.10 inch when in contact with the rudder (3) surface initially. Turn the rotating face of the dial indicator (1) to zero. Do not reset during the checking procedure. (10) With a push-pull scale (6, Table 1, 27-00-00) against the right side of the rudder (3), apply four pounds of load. Record the dial reading. (11) With a push-pull scale (6, Table 1, 27-00-00) against the left side of the rudder (3), apply four pounds of load. Record the dial reading. NOTE: The maximum freeplay travel limit is the total difference between the dial reading of Steps (10) and (11). (12) The maximum freeplay travel limit is 0.12 inch. Excess movement must be corrected. (13) If freeplay limits are exceeded, inspect all components for cracks and wear, repair or replace as required. (14) Connect the rudder tab actuator push-pull rod (do not change adjustment of the rod) to the rudder trim tab (9) using bolt, washer and nut.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL NOTE: The double clevis ends on the tab actuator push-pull rod is designed to tighten the outer clevis with a binding action against the clevis bolt and the inner clevis, which removes freeplay of the bolt in the hole. Loosen the outer clevis first (the larger nut) before removing the clevis bolt. (15) Tighten the large outer nut of the double clevis. (16) Perform the RUDDER TRAVEL BOARD REMOVAL procedure (Ref. 27-00-02).

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A

1. RUDDER AFT TORQUE SHAFT 2. RIG PIN

1 2

VIEW

NOTE: EARLIER VERSIONS OF THE TORQUE SHAFT SECTOR MAY NOT HAVE LIGHTENING HOLES INSTALLED.

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UC27B 041711AC.AI

Figure 202 Rudder Aft Rig Pin Installation

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Figure 203 Rudder Freeplay Check

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

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1. ELEVATOR 2. AFT FAIRING BULLET 3. RUDDER

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NOTCH TRIMMED FROM RUDDER TIP NOT TO EXCEED 0.50 X 0.25

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UC27B 045113AB.AI

Figure 204 Rudder Surface Trimming

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FLIGHT CONTROLS RUDDER CABLES MAINTENANCE PRACTICES

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1. RUDDER FORWARD CONTROL CABLE A. Removal (1) Attach a red tag to the control wheel with the words “Do Not Operate the Rudder System, Maintenance In Progress”. (2) Remove the pilot seat (Ref. Chapter 25-10-00, SEAT REMOVAL). (3) Remove all left flight compartment carpet. (4) Remove flight compartment floorboards 2, 4 and 21 (Ref. Chapter 6-50-00, FLOOR ACCESS PANELS). (5) Remove all left side passenger compartment seats (Ref. Chapter 25-20-00, SEAT REMOVAL). (6) Remove the passenger compartment left side carpet (Ref. Chapter 25-20-01). (7) Remove the left side cabin floorboards 16A thru 16I, 15, 24 and 25 (Ref. Chapter 6-50-00, FLOOR ACCESS PANELS). (8) Remove the aft fuselage access panels 7 and 8 (Ref. Chapter 6-50-00, FUSELAGE ACCESS PANELS). (9) Disconnect the autopilot servo cables (if equipped) (Ref. Chapter 22-10-00). (10) Airplanes equipped with Mechanical Steering (Ref. Chapter 32-50-00, Mechanical Steering Maintenance Practices) must have the steering system disconnected. This will permit rudder forces to be measured properly. The mechanical steering system is disconnected by performing the following (a) Perform THREE-POINT JACKING procedure (Ref. Chapter 7-10-00). (b) Apply electrical power to the airplane. (c) Verify that the Mechanical Steering Disconnect Actuator located on the left side of the nose landing gear wheel well is in the extended position. (d) Remove electrical power from the airplane. (11) Remove safety clips (9) from the rudder control cable turnbuckles (5 and 11) located in the aft fuselage area. (Ref. Figure 203). (12) Attach a tag with the words “forward rudder left control cable” to the forward end of the left side turnbuckle (5). (13) Attach a tag with the words “forward rudder right control cable” to the forward end of the right side turnbuckle (11).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (14) Disconnect the forward rudder left control cable (7) from the turnbuckle (5) and attach a feed line to the terminal end (6). Label the feed line with the words “forward rudder left control cable terminal end”. (15) Disconnect the forward rudder right control cable (15) from the turnbuckle (11) and attach a feed line to the terminal end (10). Label the feed line with the words “forward rudder right control cable terminal end”. (16) Remove the pressure seals located at the aft pressure bulkhead at FS 557.50. (17) Remove the cable retaining pins from the pulley brackets. Refer to Figure 201 for general location. NOTE: It may be necessary to remove cable pulleys and any fairleads if cable passage is restricted. It is permissible to remove the rudder gust lock brace installed directly above the rudder bellcrank for ease of rudder cable end removal. (18) Remove safety wire, screw and disconnect the forward cables from the forward rudder bellcrank. (19) With assistance, feed the forward rudder control cables through the pulleys and any fairleads, bringing the feed lines to the forward rudder bellcrank area. NOTE: If cable is to be reused, take care to keep the cable clean and free from damage. Coil cable loosely, no tighter than 29 inches in diameter. (20) Disconnect the feed lines from the left and right rudder forward control cables and leave the feed lines in place. (21) Remove the cables from the airplane.

B. Installation WARNING: Perform all Steps of this procedure, in the order listed. Do not skip any Steps of this procedure. Failure to do so may result in injury to personnel and damage to equipment. CAUTION: Where the cables pass through structure, the areas of possible contact between the control cables and adjacent structure must be protected with protective elements such as grommets, rub strips, blocks or guide fairings. Where contact of control cables does occur with the protective elements, a force no greater than eight (8) ounces shall be required to move the cable to a position of no contact. At no time should flight control cables contact metal structure with the protective elements removed. NOTE: The forward rudder right cable is slightly longer than the forward rudder left cable. (1) Check cable for damage and replace cable if necessary. If a used cable is installed, cable should be cleaned with solvent (2, Table 2, 27-00-00) and then dipped in corrosion preventive compound (4, Table 2, 27-00-00). Excess should be removed by wiping with a clean rag. (2) Attach the feed line labeled “forward rudder left control cable terminal end” to the forward rudder left control cable terminal end. (3) Attach the feed line labeled “forward rudder right control cable terminal end” to the forward rudder right control cable terminal end.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL NOTE: More than one person will be required to route the rudder control cables. (4) Route both forward rudder cables from the forward elevator bellcrank through the fuselage as follows (Ref. Figure 201): NOTE: It is permissible to install cable guard pins and any removed pulleys as the cable is being routed. (a) Route forward rudder right cable through the right side pulley at FS 118. (b) Route forward rudder right cable through the right side pulley at FS 210. (c) Route forward rudder right cable through the right side pulley at FS 331, if installed. This pulley was installed on UB-62 and after. (d) Route forward rudder right cable through the right side pulley at FS 495. (e) Route forward rudder right cable through the right side pulley at FS 546. (f) Route forward rudder right cable through the pressure seal retaining plate and through the right side pressure seal hole. (g) Route forward rudder right cable through the right side pulley at FS 575. (h) Route forward rudder left cable through the left side pulley at FS 118. (i) Route forward rudder left cable through the left side pulley at FS 210. (j) Route forward rudder left cable through the left side pulley at FS 331. (k) Route forward rudder left cable through the left side pulley at FS 495. (l) Route forward rudder left cable through the left side pulley at FS 534. (m) Route forward rudder left cable through the pressure seal retaining plate and through the left side pressure seal hole. (n) Route forward rudder left cable through the left side pulley at FS 575. (o) Ensure all pulleys are installed and that all cable guard pins are installed in the pulley brackets. CAUTION: Do not over torque the cable locking plate attachment screws or damage to the bellcranks will occur. Maximum torque will not exceed 15 inch-pounds. (5) Connect the forward rudder right control cable to the forward rudder bellcrank upper groove by installing the cable locking plate and attaching screw. Safety wire the screw to the bellcrank. (6) Connect the forward rudder left control cable to the forward rudder bellcrank lower groove by installing the cable locking plate and attaching screw. Safety wire the screw to the bellcrank. (7) Lubricate the turnbuckles (5 and 11) with grease (1, Table 2, 27-00-00) for corrosion protection prior to installation (Ref. Figure 203). (8) Remove the feed line and connect the forward rudder left control cable (7) to the left rudder turnbuckle (5) in the aft fuselage area.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (9) Remove the feed line and connect the forward rudder right control cable (15) to the right rudder turnbuckle (11) in the aft fuselage area. (10) Tension the cables to prevent slack. (11) Ensure that the fuselage cables are routed properly by verifying that the cables have been routed exactly as described in Step (4). Ensure the cables are engaged in the pulley grooves and all pulley guard pins are installed. (12) Remove all tape from the cables and turnbuckles. (13) Fill the pressure bulkhead cable seals with grease (1, Table 2, 27-00-00) and install the pressure seals (Ref. Figure 201). (14) Lubricate the rudder forward control cables to one inch beyond the length of travel through the pressure seal with grease (1, Table 2, 27-00-00). (15) If removed, install the rudder gust lock brace located directly above the rudder bellcrank. (16) Remove the red tag from the control wheel. (17) Perform the RUDDER OPERATIONAL CHECK procedure (Ref. 27-20-02). (18) Perform the RUDDER CONTROL SYSTEM RIGGING procedures (Ref. 27-20-02). (19) Ensure that safety clips are installed on both turnbuckles (5 and 11) (Ref. Figure 203). (20) Connect and rig the rudder autopilot servo cables (if equipped) (Ref. Chapter 22-10-00). (21) Perform LOWERING AIRPLANE FROM THREE-POINT JACKING procedure (Ref. Chapter 7-10-00). (22) Install the aft fuselage access panels 7 and 8 (Ref. Chapter 6-50-00, FUSELAGE ACCESS PANELS). (23) Install the left side cabin floorboards 16A thru 16I, 15, 24 and 25 (Ref. Chapter 6-50-00, FLOOR ACCESS PANELS). (24) Install the passenger compartment left side carpet (Ref. Chapter 25-20-01). (25) Install all left side passenger compartment seats (Ref. Chapter 25-20-00, SEAT INSTALLATION). (26) Install flight compartment floorboards 2, 4 and 21 (Ref. Chapter 6-50-00, FLOOR ACCESS PANELS). (27) Install all left side flight compartment carpet. (28) Install the pilot seat (Ref. Chapter 25-10-00, SEAT INSTALLATION).

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2. RUDDER AFT CONTROL CABLES A. Removal CAUTION: On airplanes with manual nose wheel steering, disconnect the nose wheel steering system prior to towing the airplane. The steering system is directly connected to the rudder cable system. Towing the airplane with cable blocks installed or with a bellcrank pin installed will result in damage to the airplane. (1) Attach a red tag to the control wheel with the words “Do Not Operate the Rudder System, Maintenance In Progress”. (2) Remove the aft fuselage access panels 7 and 8 (Ref. Chapter 6-50-00, FUSELAGE ACCESS PANELS). (3) Install gust locks (Ref. 27-70-00) or rudder forward bellcrank rig pin (9, Table 1, 27-00-00). (4) Disconnect the rudder autopilot servo cables (if equipped) (Ref. Chapter 22-10-00). (5) Install cable blocks (5) to the rudder left and right control cables (4) aft of the pulleys (3) in the aft fuselage area (Ref. Figure 202). (6) Remove safety clips (9) from the rudder control cable turnbuckles (5 and 11) in the aft fuselage area (Ref. Figure 203). (7) Attach a tag labeled “aft rudder right control cable” to the aft end of the right side turnbuckle (11). Disconnect cable (13) from the turnbuckle (11). (8) Attach a tag labeled “aft rudder left control cable” to the aft end of the left side turnbuckle (5). Disconnect cable (3) from the turnbuckle (5). (9) Remove safety wire from screws and disconnect the left and right aft rudder control cables (3 and 13) from the bellcranks on the aft rudder torque tube (1). (10) Remove the cables from the airplane.

B. Installation WARNING: Perform all Steps of this procedure, in the order listed. Do not skip any Steps of this procedure. Failure to do so may result in injury to personnel and damage to equipment. CAUTION: On airplanes with manual nose wheel steering, disconnect the nose wheel steering system prior to towing the airplane. The steering system is directly connected to the rudder cable system. Towing the airplane with cable blocks installed or with a bellcrank pin installed will result in damage to the airplane. (1) Check cable for damage and replace cable if necessary. If a used cable is installed, cable should be cleaned with solvent (2, Table 2, 27-00-00) and then dipped in corrosion preventive compound (4, Table 2, 27-00-00). Excess should be removed by wiping with a clean rag. CAUTION: Do not over torque the cable locking plate attachment screws or damage to the bellcranks will occur. Maximum torque will not exceed 15 inch-pounds.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (2) Identify the rudder right aft control cable (13) and connect it to the lower side of the right bellcrank on the aft rudder torque tube (1) by installing the cable locking plate and attaching screw. Safety wire the screw to the bellcrank (Ref. Figure 203). (3) Identify the rudder left aft control cable (3) and connect it to the upper side of the left bellcrank on the aft rudder torque tube (1) by installing the cable locking plate and attaching screw. Safety wire the screw to the bellcrank. (4) Lubricate the turnbuckles (5 and 11) with grease (1, Table 2, 27-00-00) for corrosion protection prior to installation. (5) Connect the aft rudder left control cable (3) to the left rudder turnbuckle (5) in the aft fuselage area. (6) Connect the aft rudder right control cable (13) to the right rudder turnbuckle (11) in the aft fuselage area. (7) Tension the cables to prevent slack. (8) Remove the cable blocks (5) from the left and right forward rudder control cables (4) in the aft fuselage area aft of the pulleys (3) (Ref. Figure 202). (9) Remove the red tag from the control wheel. (10) Perform RUDDER OPERATIONAL CHECK procedure (Ref. 27-20-02). (11) Perform RUDDER CONTROL SYSTEM RIGGING procedures (Ref. 27-20-02). (12) Ensure that safety clips on both turnbuckles are installed. (13) Connect and rig the rudder autopilot servo cables (if equipped) (Ref. Chapter 22-10-00). (14) Install the aft fuselage access panels 7 and 8 (Ref. Chapter 6-50-00, FUSELAGE ACCESS PANELS).

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Figure 201 Rudder Trim Tab Control System

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1. AFT PRESSURE BULKHEAD 2. ELEVATOR CONTROL PULLEYS 3. RUDDER CONTROL PULLEYS (LEFT) 4. RUDDER CONTROL LEFT CABLE 5. CABLE BLOCK

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UC27B 051642AA.AI

Figure 202 Rudder Control Cable Block Installation

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1. RUDDER TORQUE TUBE 2. RUDDER AUTOPILOT CABLE 3. AFT RUDDER LEFT CONTROL CABLE 4. AFT RUDDER LEFT CONTROL CABLE, LH-THREADS TERMINAL END 5. RUDDER LEFT CABLE TURNBUCKLE 6. FORWARD RUDDER LEFT CONTROL CABLE, RH-THREADS TERMINAL END 7. FORWARD RUDDER LEFT CONTROL CABLE 8. RUDDER AUTOPILOT CABLE TURNBUCKLE 9. TURNBUCKLE SAFETY CLIPS 10. FORWARD RUDDER RIGHT CONTROL CABLE, RH-THREADS TERMINAL END 11. RUDDER RIGHT CABLE TURNBUCKLE 12. AFT RUDDER RIGHT CONTROL CABLE, LH-THREADS TERMINAL END 13. AFT RUDDER RIGHT CONTROL CABLE 14. RUDDER AUTOPILOT CABLE BRIDLE CLAMP 15. FORWARD RUDDER RIGHT CONTROL CABLE 16. RUDDER AUTOPILOT TURNBUCKLE SAFETY CLIPS

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Figure 203 Rudder Control Cable Turnbuckles

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FLIGHT CONTROLS RUDDER CONTROL SYSTEM MAINTENANCE PRACTICES

27-20-02 200200

1. PROCEDURES A. Rigging WARNING: Perform all Steps of this procedure, in the order listed. Do not skip any Steps of this procedure. Failure to do so may result in injury to personnel and damage to equipment. CAUTION: Where the cables pass through structure, the areas of possible contact between the control cables and adjacent structure must be protected with protective elements such as grommets, rub strips, blocks or guide fairings. Where contact of the control cables does occur with the protective elements, a force no greater than eight (8) ounces shall be required to move the cables to a position of no contact. At no time should flight control cables contact metal structure with the protective elements removed. NOTE: Do not connect the actuating components of the autopilot until all flight control systems have been properly rigged. (1) Remove the left flight compartment seat (Ref. Chapter 25-10-00, SEAT REMOVAL). (2) Remove the left flight compartment carpet. (3) Remove the left flight compartment floorboard 2 (Ref. Chapter 6-50-00, FLOOR ACCESS PANELS). (4) Remove left and right aft fuselage panels 7 and 8 (Ref. Chapter 6-50-00, FUSELAGE ACCESS PANELS). (5) Remove the tailcone 17 (Ref. Chapter 6-50-00, FUSELAGE ACCESS PANELS). (6) Airplanes equipped with Mechanical Steering (Ref. Chapter 32-50-00, Mechanical Steering Maintenance Practices) must have the steering system as if the airplane was in-flight. This will permit rudder forces to be measured properly. The airplane must be placed on jacks and the steering disconnect actuator must be in the extended position. This is accomplished by performing the following: (a) Perform THREE-POINT JACKING (PREFERRED PROCEDURE) procedure (Ref. Chapter 7-10-00). (b) Apply electrical power to the airplane. The mechanical steering disconnect actuator should extend. (c) Verify that the mechanical steering disconnect actuator located on the left side of the nose landing gear wheel well is in the extended position. (d) Remove electrical power from the airplane. (7) Remove safety clips (7), completely loosen turnbuckle (6) and disconnect the rudder servo lower bridle cable (3) (Ref. Figure 209, Detail B).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (8) Cut safety wire (10), remove screw (11), attaching plate (12) and disconnect the rudder servo upper bridle cable (4) from the rudder aft bellcrank (1) (Ref. Figure 209, Detail C). (9) Perform the RUDDER TRAVEL BOARD INSTALLATION procedure (Ref. 27-00-02). (10) Install rig pin (2) (7, Table 1, 27-00-00) in the aft rudder torque tube (1) assembly. Using minimum force try to manually move the rudder to verify proper rig pin installation (Ref. Figure 204). (11) Install rig pin (6) (9, Table 1, 27-00-00) in the forward rudder bellcrank (Ref. Figure 205). (12) Check the cable tension by performing the following Steps: NOTE: Do not measure the tension when the temperature is varying rapidly or with the airplane located in direct sunlight. (a) Allow the temperature of the airframe to stabilize before measuring and adjusting the cable tension. (b) Measure the temperature in the compartment next to the rudder cables (2 and 4) in the aft fuselage area through panels 7 and 8 (Ref. Figure 206). (c) Refer to Rudder Cable Tension Graph Figure 203 and read the pounds of tension for the measured temperature. NOTE: Cable tension tolerance is ± 8 pounds of the tension found in Figure 203. (d) Position a cable tensiometer (4, Table 1, 27-00-00) at least three inches from the turnbuckles (1 and 3) and measure and record the tension in both the left and right cables (2 and 4) (Ref. Figure 206). It is permissible to measure tension at any point of the rudder cable system. Cable diameter is noted in Figure 203. (e) If no adjustments are required, proceed to Step (13). If adjustment is required, perform the following Steps: 1 Remove the safety clips from both turnbuckles (1 and 3) (Ref. Figure 206). WARNING: If cable tension at any time falls below 10 pounds, check all rudder system pulleys for proper cable engagement. 2 Adjust the turnbuckles until both cables have equal tension for the measured temperature found in Figure 203. Cable diameter is noted in Figure 203. 3 Remove both rig pins. Using either set of rudder pedals, move the rudder system through three full cycles to equalize tension in the rudder system cables. 4 Verify that the tension of both cables is within acceptable limits per Figure 203. If cable tension is out of limits, install rig pins and repeat Step (12). 5 Install safety clips on the turnbuckles (1 and 3) (Ref. Figure 206).

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(13) Install rig pin (2) (7, Table 1, 27-00-00) in the aft rudder torque tube (1) assembly. Using minimum force try to manually move the rudder to verify proper rig pin installation. (Ref. Figure 204). NOTE: The double clevis ends on the trim tab actuator push-pull rod are designed to tighten the outer clevis with a binding action against the clevis bolt and the inner clevis, which removes free play of the bolt in the hole. Loosen the outer clevis first (the larger nut) before removing the clevis bolt. (14) Disconnect the rudder trim tab actuator push-pull rod (do not change adjustment of the rod) and align the trim tab trailing edge with the rudder trailing edge. (15) Remove nut, washer and bolt (6) and disconnect the rudder push-pull rod (7) from the rudder control horn (4) (Ref. Figure 201). (16) Using minimum force manually move rudder to the left and check for a deflection of 25° +1°/ -0° on the rudder travel board (Ref. 27-00-02, READING A TRAVEL BOARD). Make sure that the rudder trim tab surface is aligned with the rudder surface. (17) Using minimum force manually move rudder to the right and check for a deflection of 25° +1/ -0° on the rudder travel board (Ref. 27-00-02, READING A TRAVEL BOARD). Make sure that the rudder trim tab surface is aligned with the rudder surface. (18) If no adjustment is required, proceed to Step (19). If adjustment is required, perform the following Steps: (a) Remove safety wire from stop bolt(s) (2 and 11) (Ref. Figure 201). (b) Loosen jam nut(s) (1 and 10) on stop bolt(s) (2 and 11). (c) Adjust stop bolt(s) (2 and 11) and perform Steps (16), (17) and (18) to check travel. (d) When adjustment is complete, tighten jam nut(s) (1 and 10) on stop bolt(s) (2 and 11). Perform Steps (16), (17) and (18) to check travel. (e) Safety wire stop bolt(s) (2 and 11) (178, Table 1, Chapter 91-00-00). (19) Connect the rudder push-pull rod (7) to the rudder control horn (4) and install bolt, washer and nut (6). (20) Check the position of the rudder using the rudder travel board. With aft rig pin installed, the rudder must be at neutral (0° deflection). (21) If no adjustment is required, proceed to Step (22). If adjustment is required, perform the following Steps: (a) Remove nut, washer and bolt (6) attaching push-pull rod (7) to rudder control horn (4) (Ref. Figure 201) and discard nut. (b) Loosen jam nut (8) on rudder push-pull rod (7). (c) Rotate rod end (5) as necessary until rudder is at neutral (0° deflection). (d) Verify that the threads of the rod end (5) are visible through the inspection hole (9) at the end of the push-pull rod (7) after adjustment is completed.

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(e) Tighten jam nut (8) on rudder push-pull rod (7). NOTE: When installing bolt (6), the head of the bolt must be in the up position. (f) Install bolt, washer and new nut (6) attaching push-pull rod (7) to rudder control horn (4). (22) Install new nut (6) if not already done in Step (21) (f). (23) Remove both rig pins. (24) With assistance, move the rudder until rudder control horn (4) contacts the left stop bolt (11) (Ref. Figure 201). (25) Use a go/no-go scale (8, Table 1, 27-00-00) to verify that the right stop bolt (3) on the left side forward rudder bellcrank (7) has a 0.38 +0.06/ -0 inch clearance from the stop bolt (2) located in the structure (Ref. Figure 202). (26) If no adjustment is required, proceed to Step (27). If adjustment is required, perform the following Steps: (a) Loosen jam nut (9) on right stop bolt (3). (b) Adjust stop bolt (3) as necessary to obtain proper gap clearance. (c) When adjustment is complete, tighten jam nut (9) on stop bolt (3). (27) Move the rudder until rudder control horn (4) contacts the right stop bolt (2) (Ref. Figure 201). (28) Use a go/no-go scale (8, Table 1, 27-00-00) to verify that the left stop bolt (8) on the left forward rudder bellcrank (7) has a 0.38 +0.06/ -0 inch clearance from the stop bolt (1) located in the structure (Ref. Figure 202). (29) If no adjustment is required, proceed to Step (30). If adjustment is required, perform the following Steps: (a) Loosen jam nut (10) on left stop bolt (8). (b) Adjust stop bolt (8) as necessary to obtain proper gap clearance. (c) When adjustment is complete, tighten jam nut (10) on stop bolt (8). NOTE: Pedal forces must be measured at the pivot point (2) of rudder pedal (1) (8 inch radius arm). Limits must be met with all rudder pedals (1) in both full forward and full aft adjustment positions. There should be no sudden changes in force due to friction or interference in the system. (30) Using a hand force gage, check for a maximum of 14 pounds of pedal force during movement to the full left and full right positions from the rudder pedal (1) pivot line (2) of each set of rudder pedals (Ref. Figure 207). (31) Temporarily install the tailcone to check clearances. Do not tighten all the screws until clearance checks have been completed.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (32) Remove inspection panel 16 from the left side of the tailcone (Ref. Chapter 6-50-00, FUSELAGE ACCESS PANELS). (33) Verify that the rudder installation is within the following limitations: (a) The gap at the leading edge of the rudder shall be 0.19 +0.06/ -0 inch measured between the vertical stabilizer skin trailing edge and the nose of the rudder when the rudder is deflected fully left and right. (b) The minimum gap between the top edge of the rudder and the aft fairing bullet should be 0.12 inch. (c) The gap between the trailing edge of the rudder trim tab and the tailcone should be 0.12 to 0.38 inch. (d) With full elevator down and elevator trim tab at trailing edge full down, check that there is clearance between the elevator trim tabs and the top of the rudder at full left and full right rudder. Permissible to trim a notch from the top trailing edge of the rudder (3) not to exceed 0.50 X 0.25 inch (Ref. Figure 208, Detail A). (34) Attach the rudder servo lower bridle cable (3) to turnbuckle (6) (Ref. Figure 209, Detail B). CAUTION: Do not over torque the cable locking plate attachment screw (11) or damage to the bellcrank (1) may occur. Maximum torque must not exceed 15 inch-pounds. (35) Attach rudder servo upper bridle cable (4) to the aft rudder bellcrank (1) with the attaching plate (12), install screw (11) and safety wire (10) (178, Table 1, Chapter 91-00-00) (Ref. Figure 209, Detail C). (36) Perform the RUDDER SERVO CABLE TENSIONING procedure (Ref. Chapter 22-10-00). (37) Final Travel Check. NOTE: Make sure that the rudder trim tab surface is aligned with the rudder surface. (a) Move the rudder pedals to the full left position and verify that the rudder surface moves left 25° +1°/ -0° (Ref. 27-00-02, READING A TRAVEL BOARD). (b) Through the tailcone inspection panel verify that the left rudder stop bolt (11) contacts the aft rudder control horn (4) (Ref. Figure 201). (c) Move the rudder pedals to the full right position and verify that the rudder surface moves right 25° +1°/ -0° and that the right rudder stop bolt (2) contacts the aft rudder control horn (4). (d) Through the tailcone inspection panel, verify that the right rudder stop bolt (11) contacts the aft rudder control horn (4). (e) If these requirements are not met, repeat this rigging procedure in its entirety. (38) Connect the rudder trim tab actuator push-pull rod (do not change adjustment of the rod) to the rudder trim tab using bolt, washer and nut. NOTE: The double clevis ends on the trim tab actuator push-pull rod are designed to tighten the outer clevis with a binding action against the clevis bolt and the inner clevis, which removes free play of the bolt in the hole.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (39) Tighten the large outer nut on the clevis. (40) Perform the RUDDER TRIM TAB SYSTEM RIGGING procedure (Ref. 27-20-07). (41) Install inspection panel 16 to the left side of the tailcone (Ref. Chapter 6-50-00, FUSELAGE ACCESS PANELS). (42) Complete the tailcone installation. Tighten the temporarily installed tailcone screws from Step (31). (43) Perform LOWERING AIRPLANE AFTER THREE-POINT JACKING procedure (Ref. Chapter 7-10-00). (44) Perform RUDDER TRAVEL BOARD REMOVAL procedure (Ref. 27-00-02). (45) Install the left and right aft fuselage panels 7 and 8 (Ref. Chapter 6-50-00, FUSELAGE ACCESS PANELS). (46) Install the left flight compartment floorboard 2 (Ref. Chapter 6-50-00, FLOOR ACCESS PANELS). (47) Install the left flight compartment carpet. (48) Install the left flight compartment seat (Ref. Chapter 25-10-00, SEAT INSTALLATION).

2. RUDDER A. Operational Check (1) Push forward on the pilot’s left rudder pedal and make sure that the rudder moves to the full left position smoothly with no unusual noise or binding. (2) Push forward on the pilot’s right rudder pedal and make sure that the rudder moves to the full right position smoothly with no unusual noise or binding. (3) Repeat Steps 1 and 2 using the copilot’s rudder pedals. (4) If requirements are not met, perform the RUDDER CONTROL SYSTEM RIGGING procedure in this section.

B. Cable Tension Check (1) Remove left and right aft fuselage panels 7 and 8 (Ref. Chapter 6-50-00, FUSELAGE ACCESS PANELS). (2) Remove passenger compartment seats as required (Ref. Chapter 25-20-00, PASSENGER SEATS - MAINTENANCE PRACTICES). (3) Remove passenger compartment carpet as required (Ref. Chapter 25-20-01, CARPET REMOVAL AND INSTALLATION). (4) Remove cabin floorboard panels 16A through 16J as required (Ref. Chapter 6-50-00, FLOOR ACCESS PANELS). (5) Remove safety clips (7), completely loosen turnbuckle (6) and disconnect the rudder servo lower bridle cable (3) (Ref. Figure 209, Detail B).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (6) Cut safety wire (10), remove screw (11), attaching plate (12) and disconnect the rudder servo upper bridle cable (4) from the rudder torque shaft bellcrank (1) (Ref. Figure 209, Detail C). (7) Using either set of rudder pedals move the rudder system to the left and right three cycles to equalize the rudder system tension. (8) Install a rig pin (2) (7, Table 1, 27-00-00) in the aft rudder torque tube (1) (Ref. Figure 204). (9) Check the cable tension by performing the following Steps: NOTE: Do not measure the tension when the temperature is varying rapidly or with the airplane located in direct sunlight. (a) Allow the temperature of the airframe to stabilize before measuring and adjusting the cable tension. (b) Through aft fuselage access panels 7 and 8 measure the temperature in the aft fuselage compartment next to the rudder cables (2 and 4) near the turnbuckles (1 and 3) (Ref. Figure 206). (c) Refer to the Rudder Cable Tension Graph Figure 203 and read the pounds of tension for the measured temperature. (d) Position a cable tensiometer (4, Table 1, 27-00-00) at least three inches from the turnbuckles (1 and 3) and measure the cable tension of both cables (2 and 4) (Ref. Figure 206). Cable diameter is noted in Figure 203. NOTE: Cable tension tolerance is ± 8 pounds tension. (e) If no adjustment is required, proceed to Step (10). If adjustment is required, perform the RUDDER CONTROL SYSTEM RIGGING procedure in this section. (10) Remove the rig pin (2) from the aft rudder torque tube (1) (Ref. Figure 204). (11) Attach the rudder servo lower bridle cable (3) to turnbuckle (6) (Ref. Figure 209, Detail B). CAUTION: Do not over torque the cable locking plate attachment screw (11) or damage to the bellcrank (1) may occur. Maximum torque must not exceed 15 inch-pounds. (12) Attach rudder servo upper bridle cable (4) to the aft rudder bellcrank (1) with the attaching plate (12), install screw (11) and safety wire (10) (178, Table 1, Chapter 91-00-00) (Ref. Figure 209, Detail C). (13) Perform the RUDDER SERVO CABLE TENSIONING procedures (Ref. Chapter 22-10-00). (14) Install cabin floorboard panels 16A through 16J (Ref. Chapter 6-50-00, FLOOR ACCESS PANELS). (15) Install passenger compartment carpet as required (Ref. Chapter 25-20-01, CARPET REMOVAL AND INSTALLATION). (16) Install passenger compartment seats as required (Ref. Chapter 25-20-00, PASSENGER SEAT INSTALLATION).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (17) Install left and right aft fuselage panels 7 and 8 (Ref. Chapter 6-50-00, FUSELAGE ACCESS PANELS). (18) Perform the RUDDER OPERATIONAL CHECK procedure in this section.

C. Functional Check (1) Perform the RUDDER TRAVEL BOARD INSTALLATION procedure (Ref. 27-00-02). NOTE: The double clevis ends on the trim tab actuator push-pull rods are designed to tighten the outer clevis with a binding action against the clevis bolt and the inner clevis, which removes free play of the bolt in the hole. Loosen the outer clevis first (the larger nut) before removing the clevis bolt. (2) Disconnect the rudder trim tab actuator push-pull rod (3) (do not change adjustment of the rod end) and align the trim tab trailing edge with the rudder trailing edge. (3) Push forward on the pilot’s left rudder pedal and make sure that the rudder moves to the full left position and check for a deflection of 25° +1°/ -0° (Ref. 27-00-02, READING A TRAVEL BOARD). Make sure the rudder system moves smoothly with no unusual noises or binding. (4) Push forward on the pilot’s right rudder pedal and make sure that the rudder moves to the full right position and check for a deflection of 25° +1°/ -0°. Make sure the rudder system moves smoothly with no unusual noises or binding. (5) If desired travel is achieved proceed to Step (6). If the rudder surface does not achieve desired full left and right travel, perform the RUDDER CONTROL SYSTEM RIGGING procedure in this section. (6) Connect the rudder trim tab actuator push-pull rod (3) (do not change adjustment of the rod) to the rudder trim tab. (7) Perform the RUDDER TRAVEL BOARD REMOVAL procedure (Ref. 27-00-02).

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A 1. RIGHT STOP BOLT JAM NUT 2. RIGHT STOP BOLT 3. RUDDER TORQUE TUBE 4. RUDDER CONTROL HORN 5. ROD END 6. BOLT, WASHER AND NUT 7. PUSH-PULL ROD 8. PUSH-PULL ROD JAM NUT 9. INSPECTION HOLE 10. LEFT STOP BOLT JAM NUT 11. LEFT STOP BOLT

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CL

4 10

11 5

9

8

6

7 FWD

DETAIL

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Figure 201 Rudder Stops

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1. LEFT STOP IN STRUCTURE 2. RIGHT STOP IN STRUCTURE 3. RIGHT STOP BOLT 4. RUDDER INTERCONNECT ROD 5. RIG PIN HOLE 6. RIGHT SIDE RUDDER FORWARD BELLCRANK 7. LEFT SIDE RUDDER FORWARD BELLCRANK 8. LEFT STOP BOLT 9. RIGHT STOP BOLT JAM NUT 10. LEFT STOP BOLT JAM NUT

0.38 + 0.06/-0 INCH 2

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8 3 9

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0.38 + 0.06/-0 INCH

4

7

5

4 6

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Figure 202 Rudder Forward Bellcrank Stop Assembly

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POUNDS OF TENSION

3/16" DIAMETER RUDDER CABLE TENSION GRAPH

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Figure 203 Rudder Cable Tension Graph

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A

1. RUDDER AFT TORQUE SHAFT 2. RIG PIN

1 2

VIEW

NOTE: EARLIER VERSIONS OF THE TORQUE SHAFT SECTOR MAY NOT HAVE LIGHTENING HOLES INSTALLED.

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Figure 204 Rudder Aft Rig Pin Installation

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1. LEFT STOP IN STRUCTURE 2. RIGHT STOP IN STRUCTURE 3. FORWARD PEDESTAL 4. RIGHT STOP BOLT 5. RUDDER FORWARD BELLCRANK 6. RIG PIN 7. LEFT STOP BOLT

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3

4

5 7

6

DETAIL

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Figure 205 Rudder Forward Bellcrank Rig Pin Installation

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B

1

A

1. LEFT RUDDER TURNBUCKLE 2. LEFT RUDDER CABLE 3. RIGHT RUDDER TURNBUCKLE 4. RIGHT RUDDER CABLE 5. RUDDER SERVO CABLE

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FWD VIEW LOOKING INBOARD FROM LEFT HAND SIDE DETAIL

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3 5

4

FWD VIEW LOOKING INBOARD FROM RIGHT HAND SIDE DETAIL

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Figure 206 Rudder Cable Locations

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A 1. RUDDER PEDAL 2. PIVOT POINT

1 1 2

2

COPILOT'S SHOWN, PILOT'S OPPOSITE

DETAIL

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Figure 207 Rudder Pedal Force Check

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1

2

A 3

1. ELEVATOR 2. AFT FAIRING BULLET 3. RUDDER

1

2

0.50 MAX

0.25 MAX

NOTCH TRIMMED FROM RUDDER TIP NOT TO EXCEED 0.50 X 0.25

3

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Figure 208 Rudder Surface Trimming

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C

1. RUDDER AFT BELLCRANK 2. BRIDLE CLAMP 3. RUDDER SERVO LOWER CABLE 4. RUDDER SERVO UPPER CABLE 5. RUDDER SERVO ASSEMBLY 6. TURNBUCKLE 7. SAFETY CLIPS 8. LEFT-HAND THREAD GROOVE 9. CABLE END 10. SAFETY WIRE 11. SCREW 12. PLATE

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A

B 2

9

3 4

10 11 12 DETAIL

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7

8

6

5

DETAIL

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DETAIL

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7 UC27B 061589AA.AI

Figure 209 Rudder Servo Installation

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FLIGHT CONTROLS RUDDER PEDALS MAINTENANCE PRACTICES

27-20-03 200200

1. PILOT RUDDER PEDAL A. Removal (1) Perform the SEAT REMOVAL procedure for the pilot’s seat (Ref. Chapter 25-10-00). (2) Disconnect the brake master cylinder from the right pedal (1) or left pedal (24) by removing the cotter pin, washer and pin (Ref. Figure 201). (3) Remove the two cotter pins (17), nuts (6), washers (11) and bolts (12) securing the pedal (1 or 24) to the pedal arm (2 or 23).

B. Installation (1) Connect the pedal (1 or 24) to the pedal arm (2 or 23) with the two bolts (12), washers (11), nuts (6) and cotter pins (17). Lubricate the attaching hardware with a light coat of grease (106, Table 1, Chapter 91-00-00) (Ref. Figure 201). (2) Connect the brake master cylinder to the right pedal (1) or left pedal (24) by installing the cotter pin, washer and pin. (3) Perform the SEAT INSTALLATION procedure for the pilot’s seat (Ref. Chapter 25-10-00).

C. Assembly Removal (1) Perform the SEAT REMOVAL procedure for the pilot’s seat (Ref. Chapter 25-10-00). (2) Remove the carpet as needed on the pilot’s side (Ref. Chapter 25-20-01). (3) Remove the flight compartment floorboard panel 1 (Ref. Chapter 6-50-00). (4) Disconnect the brake master cylinder from the right pedal (1) and left pedal (24) by removing the cotter pin, washer and pin (Ref. Figure 201). (5) Remove the nut (6), taper pin washer (5) and pin (3) securing the steering arm (7) to the torque tube (4). Slide the arm off of the tube. (6) Remove the cotter pin (17), nut (6), washer (11) and bolt (12) securing the right tube (14) to the right arm (22). (7) Remove the cotter pin (17), nut (6), washer (11) and bolt (12) securing the left tube (19) to the left arm (21). NOTE: Note the proper position of the arms (21 and 22), pedals (1 and 24) and bushing (10) on the torque tube (4) before removing the rudder pedal assembly. These positions must be maintained to assure proper operation when the assembly is installed. (8) Remove the nut (6), washer (11) and bolt (12) securing the bushing (10) on the torque tube (4). (9) Slide the torque tube (4) inboard until it is clear of the fitting (20).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (10) Move the torque tube (4) until it no longer aligns with the outboard fitting (20) and slide it outboard until the tube clears the inboard fitting (9). (11) Lift the rudder pedal assembly up and out of the work area.

D. Assembly Installation (1) Position the rudder pedal assembly in the work area. (2) Slide the torque tube (4) into the inboard fitting (9) (Ref. Figure 201). (3) Align the torque tube (4) with the outboard fitting (20). Slide the tube outboard until the tube is positioned in the fitting in the previous noted position. (4) Position the bushing (10) in place on the torque tube (4). Secure the bushing in place with the bolt (12), washer (11) and nut (6). (5) Position the left tube (19) in the left arm (21). Install the bolt (12), washer (11), nut (6) and cotter pin (17). (6) Position the right tube (14) in the right arm (22). Install the bolt (12), washer (11), nut (6) and cotter pin (17). CAUTION: Excessive driving of the taper pin can cause cracks in the tubing and/or casing. A lightweight rawhide or nylon mallet should be used to set the taper pin when it is installed. (7) Slide the steering arm (7) on the torque tube (4). Install the taper pin (3), taper pin washer (5) and nut (6) to secure the steering arm to the torque tube. Ensure that the small end of the taper pin is at least flush with, but not more than 0.06 inch above the surface of the arm. Torque nut 15 to 20 inch-pounds. (8) Position the brake master cylinder on the right pedal (1) and left pedal (24). Install the pin, washer and cotter pin. (9) Install the flight compartment floorboard panel and carpet as needed (Ref. Chapter 25-20-01). (10) Perform the SEAT INSTALLATION procedure for the pilot’s seat (Ref. Chapter 25-10-00).

2. PILOT RUDDER PEDAL BELLCRANK A. Removal (1) Perform THREE-POINT JACKING (PREFERRED PROCEDURE) procedure (Ref. Chapter 7-10-00). (2) Perform the SEAT REMOVAL procedure for the pilot’s seat (Ref. Chapter 25-10-00). (3) Remove the carpet as needed (Ref. Chapter 25-20-01). (4) Remove flight compartment floorboard panels 1 and 2 (Ref. Chapter 6-50-00). (5) Gain access to the bellcrank (18) (Ref. Figure 201) by removing the screws, washers and nuts securing the gust lock block channel to the airplane structure (Ref. Figure 203).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL NOTE: Release the tension on the rudder cables by loosening the turnbuckles. Disconnect the cables from the bellcrank. (6) Remove the cotter pin (17), nut (6), washer (11) and bolt (12) securing the left tube (19) to the left bellcrank (18) (Ref. Figure 201). (7) Perform the actions in Step (6) on the right tube (14). NOTE: If needed, loosen the lock nut on the end of the connecting tube (15) to aid in the removal of the tube. Note the position of the end of the tube for future use. (8) Remove the cotter pin (17), nut (6), washer (11) and bolt (12) securing the connecting tube (15) to the left bellcrank (18). (9) Remove the cotter pin (17), nut (6), washer (11) and bolt (12) securing the bellcrank (18) to the bracket (13).

B. Installation NOTE: Connect the rudder cables to the bellcrank (18) before performing the next Step (Ref. Figure 201). After the bellcrank is installed, apply slight tension to the cables by tightening the turnbuckles. (1) Position the left bellcrank (18) in the bracket (13). Install the bolt (12), washer (11), nut (6) and cotter pin (17) (Ref. Figure 201). (2) Position the connecting tube (15) in the bellcrank (18). Install the bolt (12), washer (11), nut (6) and cotter pin (17). If the lock nut was loosened, restore the tube to the original position and tighten the lock nut. (3) Position the left tube (19) in the bellcrank (18). Install the bolt (12), washer (11), nut (6) and cotter pin (17). (4) Perform the actions in Step (3) on the right tube (14). (5) Position the gust lock block channel on the airplane structure and secure with the screws, washers and nuts (Ref. Figure 203). (6) Perform the RUDDER CONTROL SYSTEM RIGGING procedure (Ref. 27-20-02). (7) Install the flight compartment floorboard panels and carpet as needed. (8) Perform the SEAT INSTALLATION procedure for the pilot’s seat (Ref. Chapter 25-10-00). (9) Perform LOWERING THE AIRPLANE AFTER THREE-POINT JACKING procedure (Ref. Chapter 7-10-00).

3. PILOT RUDDER PEDAL TUBE A. Removal (1) Perform THREE-POINT JACKING (PREFERRED PROCEDURE) procedure (Ref. Chapter 7-10-00). (2) Perform the SEAT REMOVAL procedure for the pilot’s seat (Ref. Chapter 25-10-00).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (3) Remove the carpet as needed. (4) Remove flight compartment floorboard panels 1 and 2 (Ref. Chapter 6-50-00). NOTE: Perform Steps (5) and (6) for the left tube (19). Perform Steps (7) and (8) for the right tube (14) (Ref. Figure 201). (5) Remove the cotter pin (17), nut (6), washer (11) and bolt (12) securing the left tube (19) to the left arm (21). (6) Remove the cotter pin (17), nut (6), washer (11) and bolt (12) securing the left tube (19) to the bellcrank (18). (7) Remove the cotter pin (17), nut (6), washer (11) and bolt (12) securing the right tube (14) to the right arm (22). (8) Perform the actions in Step (6) on the right tube (14).

B. Installation NOTE: Perform Steps (1) and (2) for the left tube (19). Perform Steps (3) and (4) for the right tube (14) (Ref. Figure 201). (1) Position the left tube (19) in the left arm (21). Install the bolt (12), washer (11), nut (6) and cotter pin (17). (2) Position the left tube (19) in the bellcrank (18). Install the bolt (12), washer (11), nut (6) and cotter pin (17). (3) Position the right tube (14) in the right arm (22). Install the bolt (12), washer (11), nut (6) and cotter pin (17). (4) Perform the actions in Step (2) on the right tube (14). (5) Install the flight compartment floorboard panels and carpet as needed. (6) Perform the SEAT INSTALLATION procedure for the pilot’s seat (Ref. Chapter 25-10-00). (7) Perform LOWERING THE AIRPLANE AFTER THREE-POINT JACKING procedure (Ref. Chapter 7-10-00).

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Figure 201 Pilot Rudder Pedal Bellcrank and Tube Installation

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B B

A DETAIL

3

A

PILOT SIDE 4 2

5

1

1. PIN 2. ROD END 3. RUDDER PEDAL 4. COTTER PIN 5. WASHER 6. BRAKE MASTER CYLINDER

6

DETAIL

B

Figure 202 Brake Master Cylinder Installation

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A

1. PILOTS BELLCRANK 2. GUST LOCK CHANNEL 3. SCREWS

1 2

3

3

DETAIL

A

UC27B 074053AA.AI

Figure 203 Gust Lock Channel Installation

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1. NOSE LANDING GEAR STEERING PUSH ROD 2. COTTER PIN 3. NUT 4. WASHER 5. BOLT

A 5 4

1

2 4 3

DETAIL

A UC27B 074051AA.AI

Figure 204 Nose Landing Gear Steering Rod Installation

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4. COPILOT RUDDER PEDAL A. Removal (1) Perform the SEAT REMOVAL procedure for the copilot’s seat (Ref. Chapter 25-10-00). (2) Disconnect the brake master cylinder from the right pedal (1) or left pedal (20) by removing the cotter pin, washer and pin (Ref. Figure 202). (3) Remove the two cotter pins (17), nuts (16), washers (6) and bolts (13) securing the pedal (1 or 20) to the pedal arm (2 or 19).

B. Installation (1) Connect the rudder pedal (1 or 20) to the rudder pedal arm (19 or 2) with the two bolts (13), washers (6), nuts (16) and cotter pins (17). Lubricate the attaching hardware with a light coat of grease (106, Table 1, Chapter 91-00-00) (Ref. Figure 202). (2) Connect the brake master cylinder to the right pedal (1) or left pedal (20) by installing the cotter pin, washer and pin. (3) Perform the SEAT INSTALLATION procedure for the copilot’s seat (Ref. Chapter 25-10-00).

C. Assembly Removal (1) Perform the SEAT REMOVAL procedure for the copilot’s seat (Ref. Chapter 25-10-00). (2) Remove the carpet as needed on the copilot’s side. (3) Remove flight compartment floorboard panel 1 (Ref. Chapter 06-50-00). (4) Disconnect the brake master cylinder from the right pedal (1) and left pedal (20) by removing the cotter pin, washer and pin (Ref. Figure 202). (5) Remove the cotter pin (17), nut (16), washer (6) and bolt (13) securing the right tube (7) to the right arm (18). (6) Remove the cotter pin (17), nut (16), washer (6) and bolt (13) securing the left tube (12) to the left arm (15). NOTE: Note the proper position of the arms (15 and 18), pedals (20 and 1) and bushings (5) on the torque tube (3) before removing the rudder pedal assembly. These positions must be maintained to assure proper operation when the assembly is installed. (7) Remove the nut (16), washer (6) and bolt (13) securing one of the bushings (5) on the torque tube (3). (8) Slide the torque tube (3) inboard or outboard until it is clear of the fitting (4 or 14). (9) Move the torque tube (3) until it no longer aligns with the selected fitting (4 or 14) and slide it until the tube clears the remaining fitting that still retains the tube. (10) Lift the rudder pedal assembly up and out of the work area.

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D. Assembly Installation (1) Position the rudder pedal assembly in the work area. (2) Slide the torque tube (3) into one of the fittings (4 or 14) (Ref. Figure 202). (3) Align the torque tube (3) with the other fitting (4 or 14). Slide the tube as required until the tube is positioned in the fittings in the previous noted position. (4) Position the bushings (5) in place on the torque tube (3). Secure the bushing in place with the bolt (13), washer (6) and nut (16). (5) Position the left tube (12) in the left arm (15). Install the bolt (13), washer (6), nut (16) and cotter pin (17). (6) Position the right tube (7) in the right arm (18). Install the bolt (13), washer (6), nut (16) and cotter pin (17). (7) Position the brake master cylinder on the right pedal (1) and left pedal (20). Install the pin, washer and cotter pin. (8) Install the flight compartment floorboard panels and carpet as needed. (9) Perform the SEAT INSTALLATION procedure for the copilot’s seat (Ref. Chapter 25-10-00).

5. COPILOT RUDDER PEDAL BELLCRANK A. Removal (1) Perform THREE-POINT JACKING (PREFERRED PROCEDURE) procedure (Ref. Chapter 7-10-00). (2) Perform the SEAT REMOVAL procedure for the copilot’s seat (Ref. Chapter 25-10-00). (3) Remove the carpet as needed on the copilot’s side. (4) Remove the flight compartment floorboard panels 1 and 2 (Ref. Chapter 06-50-00). (5) Remove the cotter pin (17), nut (16), washer (6) and bolt (13) securing the left tube (12) to the right bellcrank (8) (Ref. Figure 202). (6) Perform the actions in Step (5) on the right tube (7). (7) Remove the cotter pin (17), nut (16), washer (6) and bolt (13) securing the connecting tube (10) to the bellcrank (8). NOTE: If needed, loosen the lock nut (9) on the end of the connecting tube (10) to aid in the removal of the tube. Note the position of the end of the tube for future use. Remove seats and carpet as required. (8) Remove the cotter pin (17), nut (16), washer (6) and bolt (13) securing the bellcrank (8) to the bracket (11).

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B. Installation (1) Position the bellcrank (8) in the bracket (11). Install the bolt (13), washer (6), nut (16) and cotter pin (17). If the lock nut was loosened, restore the tube to the original position and tighten the lock nut (Ref. Figure 202). (2) Position the connecting tube (10) in the bellcrank (8). Install the bolt (13), washer (6), nut (16) and cotter pin (17). (3) Position the left tube (12) in the bellcrank (8). Install the bolt (13), washer (6), nut (16) and cotter pin (17). (4) Perform the actions in Step (3) on the right tube (7). (5) Perform the RUDDER CONTROL SYSTEM RIGGING procedure (Ref. 27-20-02). (6) Install the flight compartment floorboard panels and carpet as needed. (7) Perform the SEAT INSTALLATION procedure for the copilot’s seat (Ref. Chapter 25-10-00). (8) Perform LOWERING THE AIRPLANE AFTER THREE-POINT JACKING procedure (Ref. Chapter 7-10-00).

6. COPILOT RUDDER PEDAL TUBE A. Removal (1) Perform THREE-POINT JACKING (PREFERRED PROCEDURE) procedure (Ref. Chapter 7-10-00). (2) Perform the SEAT REMOVAL procedure for the copilot’s seat (Ref. Chapter 25-10-00). (3) Remove the carpet as needed on the copilot’s side. (4) Remove flight compartment floorboard panels 1 and 2 (Ref. Chapter 6-50-00). NOTE: Perform Steps (5) and (6) for the left tube (12). Perform Steps (7) and (8) for the right tube (7) (Ref. Figure 202). (5) Remove the cotter pin (17), nut (16), washer (6) and bolt (13) securing the left tube (12) to the left arm (15). (6) Remove the cotter pin (17), nut (16), washer (6) and bolt (13) securing the left tube (12) to the bellcrank (8). (7) Remove the cotter pin (17), nut (16), washer (6) and bolt (13) securing the right tube (7) to the right arm (18). (8) Perform the actions in Step (6) on the right tube (7).

B. Installation (1) Position the left tube (12) in the left arm (15). Install the bolt (13), washer (6), nut (16) and cotter pin (17) (Ref. Figure 202).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (2) Position the left tube (12) in the bellcrank (8). Install the bolt, washer, nut and cotter pin. (3) Position the right tube (7) in the right arm (18). Install the bolt (13), washer (6), nut (16) and cotter pin (17). (4) Position the right tube (7) in the bellcrank (8). Install the bolt, washer, nut and cotter pin. (5) Install the flight compartment floorboard panels and carpet as needed. (6) Perform the SEAT INSTALLATION procedure for the copilot’s seat (Ref. Chapter 25-10-00). (7) Perform LOWERING THE AIRPLANE AFTER THREE-POINT JACKING procedure (Ref. Chapter 7-10-00).

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Figure 205 Copilot Rudder Pedal Bellcrank and Tube Installation

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B B

A DETAIL

3

A

COPILOT SIDE 4 2

5

1

1. PIN 2. ROD END 3. RUDDER PEDAL 4. COTTER PIN 5. WASHER 6. BRAKE MASTER CYLINDER

6

DETAIL

B

Figure 206 Brake Master Cylinder Installation

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

7. RUDDER PEDAL ARM INSPECTION Perform the following inspection on all four rudder pedal arms at the intervals designated in Chapter 5-20-00.

A. Aluminum Rudder Pedal Arms (Unbushed) (1) Remove the desired rudder pedal from the rudder pedal arm. Perform the PILOT or COPILOT RUDDER PEDAL REMOVAL procedure, in this section. (2) Inspect the attach holes for elongation or excessive wear. Measure the wall thickness of the arm casting. If the thickness at any point is less than 0.130 inch, continue with this procedure. If the wall thickness is in tolerance, proceed to Step (7) (Ref. Figure 207, Detail A). (3) Perform the PILOT or COPILOT RUDDER PEDAL ASSEMBLY REMOVAL procedure, in this section. (4) Ream the hole out to 0.249 to 0.250 inch diameter. (5)

Install a 105740X-ZF-0250 bushing with retaining compound (91, Table 1, Chapter 91-00-00). Hold the bushing for 5 minutes after assembly. Allow the retaining compound to cure before attaching the rudder pedal. Minimum cure time is 24 hours at 70°F or 30 minutes at 250°F.

(6) Perform the PILOT or COPILOT RUDDER PEDAL ASSEMBLY INSTALLATION procedure, in this section. (7) Perform the PILOT or COPILOT RUDDER PEDAL INSTALLATION procedure, in this section.

B. Aluminum Rudder Pedal Arms With Bushings Installed in the Rudder Attach Holes (1) Remove the desired rudder pedal from the rudder pedal arm. Perform the PILOT or COPILOT RUDDER PEDAL REMOVAL procedure, in this section. (2) Inspect the bushing for any movement. If the bushing is loose, measure the distance between the outside of the bushing and the inside of the rudder pedal arm. If the measurement exceeds 0.015 inch, the rudder pedal arm should be replaced (Ref. Figure 207, Detail B). (3) Measure the wall thickness of the arm casting. If the thickness at any point is less than 0.080 inch, replace the rudder pedal arm (Ref. Figure 207, Detail A). NOTE: If the inspection does not require arm replacement or drilling, proceed to Step (7). If the rudder pedal arm must be replaced or drilled, perform the PILOT or COPILOT RUDDER PEDAL ASSEMBLY REMOVAL procedure, in this section. (4) If the measurements in Steps (2) and (3) do not exceed the limits indicated above but the hole is elongated causing the bushing to be loose, remove the bushing and ream out the holes to 0.280 to 0.281 inch. (a) Should the elongated holes not be cleaned up with the reamer, the rudder pedal arm must be replaced. (b) The 0.080 inch wall thickness minimum must not be exceeded.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (5) Install a 105739X-XD-250 bushing in the hole with retaining compound (91, Table 1, Chapter 91-00-00). Hold the bushing for 5 minutes after assembly. Allow the retaining compound to cure before attaching the rudder pedal. The minimum cure time is 24 hours at 70°F or 30 minutes at 250°F. (6) Perform the PILOT or COPILOT RUDDER PEDAL ASSEMBLY INSTALLATION procedure, in this section. (7) Perform the PILOT or COPILOT RUDDER PEDAL INSTALLATION procedure, in this section.

C. Magnesium Rudder Pedal Arms With Bushings Installed in the Rudder Attach Holes (1) Remove the two bolts that attach the brake pedal to the rudder pedal arm. (2) Inspect the bushings for any movement. If the bushing is loose, measure the distance between the outside of the bushing and the inside of the rudder pedal arm (Ref. Figure 207, Detail B). If the measurement exceeds 0.015 inch, the rudder pedal arm should be replaced. (3) Measure the wall thickness of the arm casting. If the thickness at any point is less then 0.080 inch (Ref. Figure 207, Detail A), replace the pedal arm. (4) If the measurements in Step (2) and (3) do not exceed the limits indicated above but the hole is elongated causing the bushing to be loose, remove the bushing and drill out the hole with a K (0.280 - 0.281 inch) drill bit. (5) Should the elongated holes not be cleaned up with the K drill bit, the rudder pedal arm must be replaced. (6) The 0.080 inch wall thickness minimum must not be exceeded. (7) If the above conditions are satisfactory, treat the hole surfaces with Dow 19 or equivalent corrosion prevention for magnesium parts (Ref. Chapter 20-09-00). (8) Install a 105739X-XD-250 bushing in the hole with retaining compound (91, Table 1, Chapter 91-00-00). Hold the bushing for 5 minutes after assembly. Allow the compound to cure before attaching the brake pedal. The minimum cure time is 24 hours at 70°F or 30 minutes at 250°F. (9) Install the rudder pedal and lubricate the attaching bolt with a light coat of grease (79, Table 1, Chapter 91-00-00).

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Figure 207 Rudder Pedal Arm Inspection

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FLIGHT CONTROLS RUDDER TRIM TAB MAINTENANCE PRACTICES

27-20-04 200200

1. PROCEDURES A. Removal (1) Perform the RUDDER REMOVAL procedure (Ref. 27-20-00). (2) Disconnect the bond jumper (4) from the bottom of the rudder tab (3). Remove the screw, nut and washers (2) and discard the nut (Ref. Figure 201). (3) Remove the hinge pin retention screw (5). NOTE: Lubricating the hinge and hinge pin with lubricant (106, Table 1, Chapter 91-00-00) will facilitate hinge pin removal. (4) While supporting the rudder tab (3), remove the hinge pin (1) from the hinge. Remove the tab (3) from the rudder (6).

B. Installation NOTE: Repair, modification, painting or replacement of the rudder or the rudder tab requires balancing (Ref. Chapter 55-40-00). (1) Lubricate the rudder tab hinge and the hinge pin (1) with lubricant (106, Table 1, Chapter 91-00-00) (Ref. Figure 201). (2) Position the rudder tab (3) on the rudder (6) and install the hinge pin (1). (3) Secure the end of the hinge pin (1) to the tab (3) with the retention screw (5). (4) Connect the bonding jumper (4) to the tab (3) using the screw, washer and new nut (2). (5) Perform the RUDDER INSTALLATION procedure (Ref. 27-20-00). (6) Perform the ELECTRICAL BONDING CHECK procedure (Ref. Chapter 20-00-01, ELECTRICAL BONDING-MAINTENANCE PRACTICES). (7) Perform the RUDDER TRIM TAB FUNCTIONAL CHECK procedure (Ref. 27-20-07). If the trim tab was replaced perform the RUDDER TRIM TAB RIGGING procedure (Ref. 27-20-07). (8) Perform the RUDDER TRIM TAB FREEPLAY CHECK procedure in this section.

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1. HINGE PIN 2. SCREW, WASHER, AND NUT 3. RUDDER TAB 4. BONDING JUMPER 5. RETENTION SCREW 6. RUDDER

2

3

A

4 1

5

6 FWD

VIEW LOOKING UP DETAIL

A UC27B 042933AA.AI

Figure 201 Rudder Trim Tab Removal and Installation

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C. Freeplay Check NOTE: Visually inspect the rudder tab for any damage, security of hinge attach point and for tightness of the actuating system. Inconsistencies should be remedied prior to checking the freeplay of the tab. (1) Obtain a copy of Table 1. (2) Remove the aft fuselage access panel 8 (Ref. Chapter 6-50-00, FUSELAGE ACCESS PANELS). (3) Install rig pin (2) (19, Table 7, Chapter 91-00-00) in the rudder aft torque tube (1) (Ref. Figure 202). (4) Apply tape (for paint protection) on the right surface of the rudder tab along the centerline of the tab actuator 6.50 inches aft of the tab hinge line. This will be the point of pressure against the tab by the push-pull scale. (5) Apply tape in the corresponding position on the left surface of the tab. WARNING: Ensure the tab freeplay fixture is securely attached to the rudder before releasing supporting pressure to prevent damage to equipment or injury to personnel (6) Secure the tab freeplay fixture (1) (1, Table 1, 27-00-00) to the rudder (3) so that the dial indicator stem (5) tip is positioned on the left side surface of the rudder tab (4) 6.00 inches aft of the tab hinge line at the upper edge of the rudder tab (4) (Ref. Figure 203). (7) Position the dial indicator (2) so the stem (5) is depressed 0.10 inch when in contact with the tab (4) surface initially. Turn the rotating face of the dial indicator (2) to zero. Do not reset the dial indicator (2) during this procedure. (8) With the push-pull scale (6, Table 1, 27-00-00) perform Steps (1) thru (4) and record the dial readings on Table 1. (a) Apply a 3 pound load against the left rudder tab surface. Record the dial reading as A. (b) Release half the load until a 1.5 pound load is obtained. Record the dial reading as B. (c) Apply a 3 pound load against the right rudder tab surface. Record the dial reading as C. (d) Release half the load until a 1.5 pound load is obtained. Record the dial reading as D. (9) Perform the calculations to the data on Table 1 as follows: (a) Record A, B, C and D as positive numbers. (b) Multiply B by 2 and record as 2B. (c) Subtract A from 2B and record as X. (d) Multiply D by 2 and record as 2D.

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(e) Subtract C from 2D and record as Y. NOTE: The results of X and Y can be negative numbers. (f) Add X and Y and record as E. Table 201 Rudder Trim Tab Freeplay Limits Serial Number: _______________

Date: _______________

(__________) X 2

-

=

B

(__________) 2B

(__________) X 2

=

D

(__________)

+

(__________)

X

Y

=

A -

2D

(__________)

(__________)

(__________)

X =

C =

(__________)

(__________) Y

(__________) E

(E = 0.026-inch maximum) (10) If freeplay of the tab is within the allowable limit, the tab and its linkage are in good condition. (11) If the freeplay is excessive, disconnect the tab actuator rod and visually inspect the bolts and bushings for indications of excessive wear. Replace excessively worn parts. (12) If all associated linkage is in good condition (no excessive wear), the actuator needs to be checked for excessive play and/or replaced. (13) Remove the freeplay fixture (1) and tape from the rudder (Ref. Figure 203). (14) Remove rig pin (2) from the rudder aft torque tube (1) (Ref. Figure 202). (15) Install the aft fuselage access panel 8 (Ref. Chapter 6-50-00, FUSELAGE ACCESS PANELS).

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A

1. RUDDER AFT TORQUE SHAFT 2. RIG PIN

1 2

VIEW

NOTE: EARLIER VERSIONS OF THE TORQUE SHAFT SECTOR MAY NOT HAVE LIGHTENING HOLES INSTALLED.

A

UC27B 041711AC.AI

Figure 202 Rudder Aft Rig Pin Installation

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1. TAB FREEPLAY FIXURE 2. DIAL INDICATOR 3. RUDDER 4. RUDDER TAB 5. DIAL INDICATOR STEM

A

UNITED

1

30 20

40 50

10

40 0

20

10

30

2

5

3 4

VIEW LOOKING UP AND FORWARD DETAIL

A Figure 203 Tab Freeplay Fixture Installation

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UC27B 043484AA.AI

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

FLIGHT CONTROLS RUDDER TRIM TAB CABLES AND ACTUATORS MAINTENANCE PRACTICES

27-20-05 200200

1. RUDDER TRIM TAB FORWARD CABLE A. Removal CAUTION: If tension in the system is lost during cable maintenance, check the forward cable drum and rudder trim tab actuator cable drums to ensure the cable is not unwound from drums. (1) Attach a red tag to the rudder trim tab control knob with the words “Do Not Operate, Maintenance In Progress”. (2) Remove both flight compartment carpet and seats (Ref. Chapter 25-10-00). (3) Remove the left and right pedestal side access panels. (4) Remove the left and right forward pedestal side access panels. (5) Remove the floor access panels 1 left, 2 left, 4 and 23 to gain access to the trim tab cables being removed (Ref. Chapter 6-50-00, FLOOR ACCESS PANELS). (6) Remove the environmental outlet duct assembly floor access panel 21 located under the pilot seat (Ref. Chapter 6-50-00, FLOOR ACCESS PANELS). (7) Remove the left side passenger seats and carpets as required to gain access to the trim tab cables being removed (Ref. Chapter 25-20-00). (8) Remove the left side passenger compartment floor access panels 16A through 16H to gain access to the trim tab cables being removed (Ref. Chapter 6-50-00, FLOOR ACCESS PANELS). (9) Remove belly access panel 3 just aft of the nose landing gear wheel well (Ref. Chapter 6-50-00, FUSELAGE ACCESS PANELS). (10) Remove the tie straps (1) securing the conduit tubes (6) (Ref. Figure 202). (11) Slide the conduit tubes (6) aft as required to gain access to the turnbuckles (2 and 9). (12) On the right side of the upper pedestal, move the rudder trim tab control knob to approximately align the forward cable terminal ends (3 and 4). (13) Attach cable block (5) to the left and right aft cable, forward of the bracket (8) (FS 408.25) to prevent all of the trim tab cables aft of the forward cable from moving. NOTE: Each turnbuckle barrel has a groove at one end to identify the left-hand threaded end. (14) Attach a tag with the words “forward cable left-hand threads terminal end” to the forward end of the lower inboard turnbuckle (9). (15) Remove clips and disconnect the left-hand threads terminal end (4) from the lower inboard turnbuckle (9) and attach a feed line to the terminal end. Label the feed line with the words “left-hand threads terminal end”.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (16) Attach a tag with the words “forward cable right-hand threads terminal end” to the forward end of the upper outboard turnbuckle (2). (17) Remove clips and disconnect the right-hand threads terminal end (3) from the upper outboard turnbuckle (2) and attach a feed line to the terminal end. Label the feed line with the words “right-hand threads terminal end”. (18) Remove the fairleads from the frames, if necessary. NOTE: Some of the pulleys cannot be cleared by the terminal ends even with cable guard pins removed. Remove those pulleys, if necessary to provide for adequate clearance. (19) Remove cable guard pins from pulley brackets. Refer to Figure 201 for a general location of the pulleys. NOTE: If cable is to be reused, take care to keep the cable clean and free from damage. Coil cable loosely, no tighter than ten inches in diameter. (20) While pulling the feed lines through the fuselage, withdraw both left and right-hand threads terminal ends through the fuselage and out of the belly access panel 3 (Ref. Chapter 6-50-00, FUSELAGE ACCESS PANELS). NOTE: To prevent “backlash” of the cable around the drum, tape the two cables together before removal or installation of the drum in the cable guard housing (Ref. Figure 203, Detail E). (21) Remove nut, washer and tapered pin (11) from the rudder trim tab control shaft forward universal joint (12) (Ref. Figure 204). (22) Slide forward and remove shaft (13) from the cable drum (7) and pedestal bracket. (23) Note position and number of washers (15) for installation of the cable drum for a 0.031 to 0.063 inch end play. (24) Remove the cable guard and drum (7) together through the right side forward pedestal into the cockpit area. (25) Remove the forward cable from the airplane by routing the feed lines through the forward right side of the pedestal, and into the cockpit. (26) Disconnect the feed lines from forward cable left and right-hand threads terminal ends and leave feed lines in place. (27) Unwrap cable from the drum and remove the cable lock pin (Ref. Figure 203).

B. Installation WARNING: Perform all Steps of this procedure, in the order listed. Do not skip any Steps of this procedure. Failure to do so may result in injury to personnel and damage to equipment. CAUTION: If tension in the system is lost during cable maintenance, check the forward cable drum and rudder trim tab actuator cable drums to ensure the cable is not unwound from drums.

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27-20-05

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Where the cables pass through structure, the areas of possible contact between the control cables and adjacent structure must be protected with protective elements such as grommets, rub strips, block or guide fairings. Where contact of control cables does occur with the protective elements, a force no greater than eight (8) ounces shall be required to move the cable to a position of no contact. At no time should flight control cables contact metal structure with the protective elements removed. (1) Check cable for cleanliness and damage. Dip the cable in corrosion preventative compound (4, Table 2, 27-00-00). Remove excess corrosion preventative by wiping with a clean cloth. (2) Place the cable drum (3) into the guard (5). Position the cable drum (3) in the cable drum bracket (2). Insert the drum shaft and check the end play of the cable drum of 0.031 to 0.063 inch (Ref. Figure 205). NOTE: Add washers (1) as required to obtain a cable drum end play of 0.031 to 0.063 inch. (3) When the correct end play is obtained, remove the drum shaft, cable drum (3), guard (5) and washers (1) (set aside the washers for installation). It is permissible to glue the washers (1) in place against the forward side of the drum (3) for ease of installation. (4) Attach a tag labeled “left-hand threads terminal end” to the forward cable left-hand threads terminal end. (5) Attach a tag labeled “right-hand threads terminal end” to the forward cable right-hand threads terminal end. (6) Wrap the forward cable on the cable drum as follows (Ref. Figure 203): CAUTION: Do not kink the cable while locating the middle of the offset forward cable. Damage to the cable will occur. (a) Offset the terminal ends of the forward cable by 1.50 inches with the right-hand threaded terminal end side of cable longer. With the cable ends offset, carefully mark the midpoint of the offset cable with ink or paint. With the right-handed threads terminal end side of the cable located on the flat side of the drum, position the mark on the cable in the middle of the cable drum slot and install the cable lock pin. With the cable lock pin installed, and with the drum unwound, verify that the terminal ends are still offset by 1.50 inches (Ref. Detail B). (b) From the lock pin, wrap each cable 2 1/4 turns around the drum beginning with the outside grooves and work toward the middle of the drum (Ref. Detail C). With the drum wound, verify that the terminal ends are still offset by 1.50 inches (Ref. Detail B). (c) Position the cable guard over the drum and tape the forward cables together just outside of the cable guard to prevent cable backlash at the drum (Ref. Detail E). When applying tape to the cable, make sure the cables are separated (not crossed) so that it is easy to identify which cable end winds off the left and right side of the drum. (7) In the cockpit attach the right-hand threads terminal end to the feed line labeled “right-hand threads terminal end”. (8) In the cockpit attach the left-hand threads terminal end to the feed line labeled “left-hand threads terminal end”.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL NOTE: More than one person will be required to route the forward cable. Take precautions to keep the cable clean and free from damage. (9) Pull the feed lines from the belly access panel 3 (Ref. Chapter 6-50-00, FUSELAGE ACCESS PANELS), draw the forward cables through the right side of the pedestal and then out of the belly access panel, until the drum is close to the pedestal. (10) Position washers (15) that were set aside in Step (3) on the forward side of the cable drum (Ref. Figure 204). (11) Position the cable drum (7), guard and washers (15) into the drum bracket forward of the pedestal and install the cable drum shaft (13). (12) On the right side of the upper pedestal, move the rudder trim tab control knob (14) to align the slot in the forward universal joint (12) with the cable drum shaft (13) and install the tapered pin, washer and nut (11). (13) Identify the forward cable with the left-hand threads terminal end (6) and make sure it winds off the left side of the drum as installed. Identify the forward cable with the right-hand threads terminal end (5) and make sure it winds off the right side of the drum as installed. Route the cable from the drum as follows (Ref. Figure 204, Detail B): NOTE: It is permissible to install cable guard pins as the cable is being routed. (a) Route the right-hand threaded terminal end cable (4 and 5) through the right pulley (1 and 8) (right side of the belly access panel). (b) Route the left-hand threaded terminal end cable (2 and 6) through the aft left pulley (3 and 9) (left side of the belly access panel area). (c) Using the feed lines, pull the cable into the fuselage and continue with Step (d). (d) Route left-hand threaded terminal end cable (6) over the bottom pulley (10) (lower inboard set of pulleys at FS 103 under the pilot). This cable continues as the bottom cable under the pilot and into the forward cabin through the fairleads. (e) Route right-hand threaded terminal end cable (5) over the top pulley (10) (lower inboard set of pulleys at FS 103 under the pilot). This cable continues as the top cable under the pilot and into the forward cabin through the fairleads. (14) Using the feed lines pull the cable ends through the fuselage to the turnbuckle connections. (15) Lubricate turnbuckles (2 and 9) with grease (1, Table 2, 27-00-00) for corrosion protection prior to installation (Ref. Figure 202). (16) Remove feed line, and attach the left-hand threads terminal end (4) to the turnbuckle (9). (17) Remove feed line, and attach the right-hand threads terminal end (3) to the turnbuckle (2). (18) Tension the forward cable sufficient to prevent slack. (19) Make sure that the forward cable is routed properly and is engaged in the pulleys. (20) Remove cable block (5).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (21) Remove all tape from the cable, turnbuckles and pulleys. (22) Install all cable guard pins and pulleys removed during the RUDDER TRIM TAB FORWARD CABLE REMOVAL procedure. (23) Perform the RUDDER TRIM TAB SYSTEM RIGGING procedure (Ref. 27-20-07). (24) Ensure turnbuckles (2 and 9) have been safety clipped and install conduit tubes (6) over the turnbuckles (2 and 9) between the conduit brackets (8) FS 408.25 and (7) FS 378.25. Secure the conduit tubes (6) with tie straps (1), forward of bracket (7) and aft of bracket (8) (Ref. Figure 202). (25) Install the left side passenger compartment floor access panels 16A through 16H (Ref. Chapter 6-50-00, FLOOR ACCESS PANELS). (26) Install left side passenger compartment carpet (Ref. Chapter 25-20-01) and seats (Ref. Chapter 25-20-00). (27) Install the environmental outlet duct assembly access panel 21 located under the pilot seat (Ref. Chapter 6-50-00, FLOOR ACCESS PANELS). (28) Install floor access panels 1 left, 2 left, 4 and 23 (Ref. Chapter 6-50-00, FLOOR ACCESS PANELS). (29) Install left and right pedestal side access panels. (30) Install left and right forward pedestal side access panels. (31) Install flight compartment carpet and seats (Ref. Chapter 25-10-00). (32) Install belly access panel 3 just aft of the nose landing gear wheel well (Ref. Chapter 6-50-00, FUSELAGE ACCESS PANELS). (33) Remove the red tag from the rudder trim tab control knob.

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RUDDER TRIM TAB ACTUATOR

RUDDER

1

E

2 RUDDER TRIM TAB ACTUATOR RUDDER TRIM TAB

DETAIL

E

CONTROL HORN

1

CABIN PRESSURE SEALS

FILL WITH SEALANT BETWEEN SEALS (BOTH SIDES)

2

TURNBUCKLES DETAIL

CANTED FUSELAGE STATION 605.98

D CABLE STOPS

2

TURNBUCKLES

1 (VIEW LOOKING DOWN) CABLE PRESSURE SEALS DETAIL

R L

L

B

B

R

D

C 1. RUDDER TRIM TAB CABLE 2. RUDDER TRIM TAB CABLE

PRESSURE SEALS FS 557.50 1

2

CABLE STOPS 2

2

DETAIL

DETAIL

C 1

A

FS 545

FS 511

1 TURNBUCKLES FS 390

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G

2

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CONDUIT TUBES

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STRAPS

R

D

1 2

2 2

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CONDUITS

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DETAIL

FS 104 FS 378

G FS 408

UC27B 030269AC.AI

Figure 201 Rudder Trim Tab Control System

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1. TIE STRAPS 2. TURNBUCKLE 3. RIGHT - HAND THREADS TERMINAL END 4. LEFT - HAND THREADS TERMINAL END 5. BLOCK (TEMPORARY) 6. CONDUIT TUBES 7. BRACKET 8. BRACKET 9. TURNBUCKLE

A

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VIEW LOOKING DOWN FS 408.25 (REF)

DETAIL 5

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FWD 7

3

9

8

4

VIEW LOOKING DOWN (CONDUIT TUBES MOVED AFT AND BLOCK TEMPORARILY INSTALLED) DETAIL

FS 378.25 (REF)

FWD UC27B 043246AD.AI

A

Figure 202 Rudder Trim Tab Forward Cable Routing

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ED

TAPERED PIN, WASHER AND NUT

CABLE DRUM SHAFT

A

WASHERS CABLE DRUM

B C

CABLE GUARD

DETAIL

FORWARD CABLE WITH RIGHT - HAND THREADS TERMINAL END

A FLAT SIDE OF DRUM

MARK ON CABLE TO MIDDLE OF CABLE DRUM SLOT

CABLE LOCK PIN INSTALLED

CABLE LOCK PIN

FLAT SIDE OF DRUM

LEFT- HAND THREADS SIDE OF CABLE

FORWARD CABLE WITH LEFT - HAND THREADS TERMINAL END (DRUM NOT FULLY WRAPPED) (CABLE GUARD NOT SHOWN)

RIGHT- HAND THREADS SIDE OF CABLE OFFSET 1.50 INCHES

DETAIL

C

(CABLE GUARD NOT SHOWN) DETAIL

B

CABLE WITH RIGHT-HAND THREADS TERMINAL END BEGINNING AT THE FLAT SIDE OF THE DRUM

FORWARD SIDE OF DRUM CABLE WITH LEFT - HAND THREADS TERMINAL END BEGINNING AT (DRUM FULLY WRAPPED) THE SPROCKET SIDE OF THE DRUM DETAIL

CABLE GUARD

FORWARD CABLE LEFT - HAND THREADS TERMINAL END

FORWARD CABLE WITH RIGHT - HAND THREADS TERMINAL END

D

TAPE (TEMPORARY)

DETAIL

E

THIS INFORMATION IS FOR THE MODEL 1900/1900C ONLY. DO NOT USE ON OTHER MODEL 1900 SERIES AIRCRAFT, BECAUSE THERE ARE CRITICAL DIFFERENCES.

Figure 203 Rudder Trim Tab Forward Cable Drum Wrapping

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

1. RIGHT SIDE PULLEY AND BRACKET 2. LEFT-HAND THREADED TERMINAL END CABLE 3. LEFT SIDE PULLEY AND BRACKET 4. RIGHT-HAND THREADED TERMINAL END CABLE

A 1 2

4

FWD 3 VIEW LOOKING UP DETAIL

UC27B 043274AB.AI

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Figure 204 (Sheet 1 of 2) Rudder Trim Tab Forward Cable Routing

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5 5

6 6 10

9 6

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B

8 5 FORWARD LOOKING AFT DETAIL

B

13

15

14

7 11 5. FORWARD (RIGHT - HAND THREADS TERMINAL END) CABLE 6. FORWARD (LEFT - HAND THREADS TERMINAL END) CABLE 7. CABLE DRUM 8. RIGHT SIDE PULLEY 9. LEFT SIDE PULLEY 10. PULLEYS AT FS 103.00 11. TAPERED PIN, WASHER AND NUT 12. FORWARD UNIVERSAL JOINT 13. CABLE DRUM SHAFT 14. CONTROL KNOB 15. WASHERS

12

6 5

DETAIL

A UC27B 043376AB.AI

Figure 204 (Sheet 2 of 2) Rudder Trim Tab Forward Cable Routing

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4 1. WASHERS 2. BRACKET 3. DRUM 4. PEDESTAL 5. DRUM GUARD

3

2

1

FWD

5 END PLAY 0.03 TO 0.063

VIEW LOOKING UP AT BOTTOM OF FWD RUDDER TRIM TAB CABLE DRUM. UC27B 045414AA.AI

Figure 205 Rudder Trim Tab Forward Drum Installation

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2. RUDDER TRIM TAB MIDDLE CABLE A. Removal CAUTION: If tension in the system is lost during cable maintenance, check the forward cable drum and rudder trim tab actuator cable drum to ensure the cable is not unwound from drums. (1) Attach a red tag to the rudder trim tab control knob with the words “Do Not Operate, Maintenance In Progress”. (2) Remove the left side passenger seats and carpets as required to gain access to the trim tab cables being removed (Ref. Chapter 25-20-00). (3) Remove the left side passenger compartment floor access panels 16G, 16H and 16I to gain access to the trim tab cables being removed (Ref. Chapter 6-50-00, FLOOR ACCESS PANELS). (4) Remove the aft floor access panels 15, 24 and 25 to gain access to the trim cables being removed (Ref. Chapter 6-50-00, FLOOR ACCESS PANELS). (5) Remove aft fuselage access panel 7 (Ref. Chapter 6-50-00, FUSELAGE ACCESS PANELS). (6) If equipped with electric elevator trim, remove the aft fuselage panel 8 (Ref. Chapter 6-50-00, FUSELAGE ACCESS PANELS) and the elevator trim servo cover over the lower pulley bracket at canted fuselage station 605.98. (7) Remove the tie straps (1) securing the conduit tubes (2) (Ref. Figure 206). (8) Slide the conduit tubes (2) forward as required to gain access to the turnbuckles (4 and 8). (9) On the right side of the upper pedestal, move the rudder trim tab control knob to approximately align the forward turnbuckles (4 and 8). (10) Attach cable block (9) to the left and right forward cable, aft of the bracket (3) (FS 378.25) to prevent the forward trim cables from moving. (11) Install cable block (6) to the rudder trim tab actuator cables (4) at the lower pulley bracket (5) at canted fuselage station 605.98 to prevent the rudder actuator trim cables from moving (Ref. Figure 207). NOTE: Each turnbuckle barrel has a groove at one end to identify the left-hand threaded end. (12) Remove the turnbuckle safety clips and attach a tag with the words “left-hand threaded terminal end” to the aft end of the outboard turnbuckle (4) (Ref. Figure 206). (13) Disconnect the left-hand threaded terminal end (5) from the outboard turnbuckle (4) and attach a feed line to the terminal end. Label the feed line with the words “left-hand threaded terminal end”. (14) Remove the turnbuckle safety clips and attach a tag with the words “right-hand threaded terminal end” to the aft end of the inboard turnbuckle (8). (15) Disconnect the right-hand threaded terminal end (7) from the inboard turnbuckle (8) and attach a feed line to the terminal end. Label the feed line with the words “right-hand threaded terminal end”. (16) Remove cable guard pins from pulley brackets. Refer to Figure 201 for a general location of the pulleys. Page 214 Nov 1/09

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL NOTE: If cable is to be reused, take care to keep the cable clean and free from damage. Coil cable loosely, no tighter than ten inches in diameter. (17) In the aft fuselage area at lower canted station 605.98 remove clips and disconnect the left-hand threaded terminal end (2) from the inboard turnbuckle (3). Mark the turnbuckle “left-hand threaded terminal end” (Ref. Figure 207). (18) In the aft fuselage area at lower canted station 605.98 remove clips and disconnect the right-hand threaded terminal end (8) from the outboard turnbuckle (7). Mark the turnbuckle “right-hand threaded terminal end”. (19) Using the feed lines, pull the left-hand threaded terminal end cable through the inboard pulley at FS 511. (20) Using the feed lines, pull the right-hand threaded terminal end cable through the outboard pulley at FS 511. (21) Using the feed lines, pull the left-hand threaded terminal end cable through the inboard pulley at FS 545. (22) Using the feed lines, pull the right-hand threaded terminal end cable through the outboard pulley at FS 545. (23) Lubricate the right-hand threaded terminal end cable with grease (1, Table 2, 27-00-00) and carefully feed the terminal end through the inboard pressure seal at FS 557.50 (Ref. Figure 201). (24) Lubricate the left-hand threaded terminal end cable with grease (1, Table 2, 27-00-00) and carefully feed the terminal end through the outboard pressure seal at FS 557.50. (25) Remove the inboard and outboard middle cables (1 and 9) from the airplane by feeding the cable through the aft fuselage access panel 7 (Ref. Chapter 6-50-00, FUSELAGE ACCESS PANELS) (Ref. Figure 207). (26) Disconnect and identify the feed lines from the inboard and outboard middle cable left and right-hand threaded terminal ends and leave feed lines in place.

B. Installation WARNING: Perform all Steps of this procedure, in the order listed. Do not skip any Steps of this procedure. Failure to do so may result in injury to personnel and damage to equipment. CAUTION: If tension in the system is lost during cable maintenance, check the forward cable drum and rudder trim tab actuator cable drum to ensure the cable is not unwound from drums. Where the cables pass through structure, the areas of possible contact between the control cables and adjacent structure must be protected with protective elements such as grommets, rub strips, block or guide fairings. Where contact of control cables does occur with the protective elements, a force no greater than eight (8) ounces shall be required to move the cable to a position of no contact. At no time should flight control cables contact metal structure with the protective elements removed. (1) Check cable for cleanliness and damage. Replace cable if necessary. Dip the cable in corrosion preventative compound (4, Table 2, 27-00-00). Remove excess corrosion preventative by wiping with a clean cloth.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL NOTE: The forward end of each cable has a longer terminal end. (2) At the aft fuselage access panel 7 (Ref. Chapter 6-50-00, FUSELAGE ACCESS PANELS) attach the feed line labeled “right-hand threaded terminal end” to the longer right-hand threads terminal end. (3) At the aft fuselage access panel 7 (Ref. Chapter 6-50-00, FUSELAGE ACCESS PANELS) attach the feed line labeled “left-hand threads terminal end” to the longer left-hand threads terminal end. NOTE: More than one person will be required to route the middle cable. Take precautions to keep the cable clean and free from damage. It is permissible to install cable guard pins as the cable is being routed. (4) Lubricate the right-hand threaded terminal end cable with grease (1, Table 2, 27-00-00) and carefully feed the terminal end through the inboard pressure seal at FS 557.50 (Ref. Figure 201). (5) Lubricate the left-hand threaded terminal end cable with grease (1, Table 2, 27-00-00) and carefully feed the terminal end through the outboard pressure seal at FS 557.50. (6) Using the feed lines, pull the cable into the aft fuselage cargo area and route left-hand threaded terminal end cable through the outboard pulley at FS 545. (7) Using the feed lines, pull the cable into the aft fuselage cargo area and route right-hand threaded terminal end cable through the inboard pulley at FS 545. (8) Route left-hand threaded terminal end cable through the outboard pulley at FS 511 to the turnbuckles in the aft fuselage area. (9) Route right-hand threaded terminal end cable through the inboard pulley at FS 511 to the turnbuckles in the aft fuselage area. (10) Lubricate turnbuckles (4 and 8) at FS 390 with grease (1, Table 2, 27-00-00) for corrosion protection prior to installation (Ref. Figure 206). (11) Remove feed line, and attach the left-hand threads terminal end (5) to the outboard turnbuckle (4). (12) Remove feed line, and attach the right-hand threads terminal end (7) to the inboard turnbuckle (8). (13) Lubricate turnbuckles (3 and 7) at lower canted station 605.98 with grease (1, Table 2, 27-00-00) for corrosion protection prior to installation (Ref. Figure 207). (14) In the aft fuselage area at lower canted station 605.98 connect the left-hand threaded terminal end (2) to the inboard turnbuckle (3). (15) In the aft fuselage area at lower canted station 605.98 connect the right-hand threaded terminal end (8) to the outboard turnbuckle (7). (16) Tension the middle cables sufficient to prevent slack. (17) Verify that the cable is routed properly as described in Steps (4) through (15). (18) Remove cable block (6) at lower canted station 605.98. (19) Remove cable block (9) at FS 378.25 (Ref. Figure 206).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (20) Remove all tags from the cables and turnbuckles. (21) Install all cable guard pins and pulleys removed during the RUDDER TRIM TAB MIDDLE CABLE REMOVAL procedure. Ensure cables are engaged in the pulley grooves and all guard pins are installed. (22) Fill the pressure seal and lubricate the cables to one inch beyond the length of cable travel through the pressure seals with grease (1, Table 2, 27-00-00). (23) Perform the RUDDER TRIM TAB SYSTEM TENSION CHECK procedure (Ref. 27-20-07). (24) Ensure turnbuckles (4 and 8) have been safety clipped and install conduit tubes (2) over the turnbuckles (4 and 8) between the conduit brackets (6) FS 408.25 and (3) FS 378.25. Secure the conduit tubes (2) with tie straps (1), forward of bracket (3) and aft of bracket (6) (Ref. Figure 206). (25) Perform the RUDDER TRIM TAB SYSTEM FUNCTIONAL CHECK procedure (Ref. 27-20-07). (26) Install the left side passenger compartment floor access panels 16G, 16H and 16I (Ref. Chapter 6-50-00, FLOOR ACCESS PANELS). (27) Install the aft floor access panels 15, 24 and 25 (Ref. Chapter 6-50-00, FLOOR ACCESS PANELS). (28) Install left side passenger compartment carpet (Ref. Chapter 25-20-01) and seats (Ref. Chapter 25-20-00). (29) Install the elevator trim servo cover over the lower pulley bracket at canted fuselage station 605.98, if equipped with electric elevator trim. (30) If equipped with electric elevator trim, install aft fuselage panel 8 and the elevator trim servo cover over the lower pulley bracket at canted fuselage station 605.98. (31) Install aft fuselage access panel 7 (Ref. Chapter 6-50-00, FUSELAGE ACCESS PANELS). (32) Remove the red tag from the rudder trim tab control knob.

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1. TIE STRAPS 2. CONDUIT TUBES 3. BRACKET 4. TURNBUCKLE 5. LEFT-HAND THREADS TERMINAL END 6. BRACKET 7. RIGHT-HAND THREADS TERMINAL END 8. TURNBUCKLE 9. BLOCK (TEMPORARY)

A

1 2

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VIEW LOOKING DOWN DETAIL

FS 408.25 (REF)

A

FWD

3

4

5

FS 378.25 (REF)

2

6

7

8

9

VIEW LOOKING DOWN (CONDUIT TUBES MOVED FORWARD AND BLOCK TEMPORARILY INSTALLED) DETAIL

A

Figure 206 Rudder Trim Tab Middle Cable

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

A 1. INBOARD MIDDLE RUDDER TRIM CABLE 2. LEFT - HAND THREADED TERMINAL END 3. INBOARD TURNBUCKLE 4. RUDDER TRIM ACTUATOR CABLES 5. LOWER PULLEY BRACKET 6. CABLE BLOCK 7. OUTBOARD TURNBUCKLE 8. RIGHT HAND THREADED TERMINAL END 9. OUTBOARD MIDDLE RUDDER TRIM CABLE

4

3 2 1 5 6 FWD

7 8

(SERVO COVER REMOVED FOR CLARITY, IF INSTALLED)

9 DETAIL

A UC27B 043869AA.AI

Figure 207 Rudder Trim Tab Middle Cable Aft Fuselage

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3. RUDDER TRIM TAB ACTUATOR AND CABLE A. Removal CAUTION: If tension in the system is lost during cable maintenance, check the forward cable drum and rudder trim tab actuator cable drums to ensure the cable is not unwound from drums. (1) Attach a red tag to the rudder trim control knob with the words “Do Not Operate, Maintenance In Progress”. (2) Remove vertical stabilizer access panel 24 (Ref. Chapter 6-50-00, STABILIZER ACCESS PANELS). (3) Remove aft fuselage access panel 7 (Ref. Chapter 6-50-00, FUSELAGE ACCESS PANELS). (4) Remove aft fuselage floorboard panel 25 forward of the aft pressure bulkhead (Ref. Chapter 6-50-00, FLOOR ACCESS PANELS). (5) Remove elevator trim servo cover over the lower pulley bracket at canted fuselage station 605.98, if installed. (6) Install cable blocks (3) to the rudder trim cables (2) at the first set of pulleys (4) forward of the aft pressure bulkhead (1) (Ref. Figure 208). (7) Remove the cable stop guard (2) from the lower pulley bracket (6) at canted fuselage station 605.98 (Ref. Figure 209). (8) Remove rudder trim tab cable retaining pins (3) from both sets of pulleys in the aft fuselage area at canted fuselage station 605.98. NOTE: Each turnbuckle barrel (5) has a groove at one end to identify the left-hand threaded end. (9) Attach a tag with the words “outboard cable left-hand threads terminal end” to the aft end of the turnbuckle in the aft fuselage area. (10) Disconnect the left-hand threads terminal end from turnbuckle and attach a feed line to the terminal end. Label the feed line with the words “outboard cable left-hand threads terminal end”. (11) Attach a tag with the words “inboard cable right-hand threads terminal end” to the aft end of the turnbuckle in the aft fuselage area. (12) Disconnect the right-hand threads terminal end from turnbuckle and attach a feed line to the terminal end. Label the feed line with the words “inboard cable right-hand threads terminal end”. NOTE: If cable is to be reused, take care to keep the cable clean and free from damage. Coil cable loosely, no tighter than ten inches in diameter. The double clevis ends on the tab actuator push-pull rod is designed to tighten the outer clevis with a binding action against the clevis bolt and the inner clevis, which removes free play of the bolt in the hole. Loosen the outer clevis first (the larger nut) before removing the clevis bolt. (13) Loosen the large jam nut on the double clevis (3) (Ref. Figure 210).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (14) Remove nut, washer and bolt (2) and disconnect the rudder tab push-pull rod double clevis (3) from the rudder trim tab horn (6). (15) Attach a feed line to the rod end double clevis (3). (16) Tape cables together at the rudder trim actuator to prevent backlash or drum movement during removal. (17) With assistance, carefully feed the rudder trim tab cables through the pulleys at canted fuselage station 605.98 and up through the vertical stabilizer bringing the cables out through the vertical stabilizer access panel. (18) Disconnect the feed lines from the cable terminal ends leaving the feed lines in place. (19) Remove bolts, washers (4) and shim (5) (if installed) from the rudder trim tab actuator (3) (Ref. Figure 211). NOTE: Note the location of bolts (4) and shim (5) (if installed) being removed for later installation. (20) Rotate the rudder trim tab actuator (3) clockwise (facing aft) and remove the elevator system cable (1) from the cable fairlead (2) on the rudder trim actuator (3). (21) Remove the rudder trim tab actuator (3) and rod from the airplane. (22) Disconnect the feed line from the rod end clevis and leave in place.

B. Drum Cable Replacement (1) Remove bolts (20), washers and nuts (19) from the rudder trim tab actuator (1) end cap (Ref. Figure 212). (2) Remove the elevator system cable fairlead bracket (21) from the rudder trim tab actuator end cap (24). (3) Remove the rudder trim tab actuator end cap (24) from the rudder trim tab actuator (1). (4) Carefully remove the cable drum end cap bearing (23) from the actuator housing (22). (5) Remove bolt, washer and nut (17) from the guide (18) on rudder trim tab actuator rod (3). (6) Rotate the rod (3) counterclockwise and unscrew the rod (3) from the cable drum (25). (7) Remove cable drum and cable (25) from the actuator housing (22). (8) Unwind the cable from the cable drum. (9) Remove the cable retaining pin from the drum and remove the cable. (10) Check cable for cleanliness and damage. Replace cable if necessary. Dip the cable in corrosion preventative compound (4, Table 2, 27-00-00). Remove excess corrosion preventative compound by wiping with a clean cloth. (11) Attach a tag labeled “outboard cable left-hand threads terminal end” to the outboard cable left-hand threads terminal end.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (12) Attach a tag labeled “inboard cable right-hand threads terminal end” to the inboard cable right-hand threads terminal end. (13) Wrap the rudder trim tab actuator left-hand threads terminal end and right-hand threads terminal end on the cable drum as follows (Ref. Figure 213): CAUTION: Do not kink the cable while locating the middle of the actuator cable. Damage to the cable will occur. (a) Align the terminal ends of the rudder trim actuator cable together and mark the middle of the rudder trim actuator cable with ink or paint. Locate the side of the rudder trim actuator cable with the left-hand threads terminal end. Position the middle of the cable in the middle of the cable drum slot. With the left-hand threads terminal end side of the cable, on the large bearing race side of the cable drum, install the cable lock pin in the middle of the cable drum slot (Ref. Detail B). (b) From the lock pin, wrap each cable 4 1/4 turns around the drum beginning with the outside grooves and work toward the middle of the drum (Ref. Detail C). (c) Position the drum into the actuator housing and tape the rudder trim tab actuator cables together just outside of the actuator housing to prevent cable backlash at the drum. When applying tape to the cable, make sure the cables are separated (not crossed) so that it is easy to identify which cable end winds off the outboard and inboard side of the drum. (14) Rotate the actuator rod (3) clockwise and screw the rod (3) into the cable drum (25) until a measurement of 1.31 ± 0.06 inches from the actuator housing (2) to the end of the rod (3) has been obtained (Ref. Figure 212). (15) Align the hole in the rudder trim tab actuator rod and install bolt, washer, and nut (17) through the guide (18) and the rudder trim tab actuator rod (3). (16) Carefully install the cable drum end cap bearing (23) in the actuator housing (22). (17) Position the end cap (24) onto the rudder trim tab actuator (1). (18) Position the elevator system cable fairlead bracket (21) onto the top of the rudder trim tab actuator end cap (24) and install bolts (20), washers and nuts (19). (19) With cable terminal ends together check the measurement from the actuator housing (2) to the center of the rod end (6) for 2.50 ± 0.03 inches. (20) If no adjustment is required, proceed to Step (21). If adjustment is required, perform the following Steps: (a) Remove safety wire from the jam nut (5) to the tab washer (4). (b) Loosen jam nut (5) on rod end (6). (c) Rotate the rod end (6) to obtain measurement of 2.50 ± 0.03 inches. (d) Tighten the jam nut (5) to the rod end (6). (e) Install safety wire from the jam nut (5) to the tab washer (6). (21) Check the measurement of the actuator rod (10) for 14.60 ± 0.06 inches in length.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

(22) If no adjustment is required, proceed to the RUDDER TRIM TAB ACTUATOR AND CABLE INSTALLATION procedure. If adjustment is required, perform the following Steps: (a) Loosen jam nut (9) on the outer double clevis (7). (b) Adjust the outer double clevis (7) until measurement is obtained. (c) Tighten jam nut (9) on the outer double clevis (7) and check measurement.

C. Installation CAUTION: If tension in the system is lost during cable maintenance, check the forward cable drum and rudder trim tab actuator cable drums to ensure the cable is not unwound from drums. Where the cables pass through structure, the areas of possible contact between the control cables and adjacent structure must be protected with protective elements such as grommets, rub strips, block or guide fairings. Where contact of control cables does occur with the protective elements, a force no greater than eight (8) ounces shall be required to move the cable to a position of no contact. At no time should flight control cables contact metal structure with the protective elements removed. (1) Attach the feed line to the rudder trim tab actuator rod double clevis end. (2) Position rudder trim tab actuator (3) and push-pull rod in place and carefully feed the rudder trim tab actuator push-pull rod through the rudder assembly (Ref. Figure 211). (3) Rotate the rudder trim tab actuator (3) clockwise (facing aft) and install the elevator system cable (1) into the cable fairlead (2) on the actuator. (4) Align the rudder trim tab actuator mounting holes and install shim (5) (if installed), washers and bolts (4). Bond shim (5) to the rudder trim tab actuator (3) flange using adhesive (170, Table 1, Chapter 91-00-00). (5) Attach feed lines to the rudder trim tab actuator cable (6) terminal ends. (6) With assistance, carefully feed the rudder trim tab cables (6) down through the vertical stabilizer and through both set of pulleys at canted fuselage station 605.98 to the turnbuckles in the aft fuselage area. (7) Lubricate all turnbuckles with grease (1, Table 2, 27-00-00) for corrosion protection prior to installation. NOTE: Each turnbuckle barrel has a groove at one end to identify the left-hand threaded end. (8) Remove feed line, and attach the left-hand threads terminal end to the turnbuckle. (9) Remove feed line, and attach the right-hand threads terminal end to the turnbuckle. (10) Tension the rudder trim cable sufficient to prevent slack. (11) Remove all tape from the cables and turnbuckles. (12) Install all cable guard pins removed from the upper and lower pulley brackets located at canted fuselage station 605.98.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (13) Install the cable stop guard (2) to the lower pulley bracket (6) located at canted fuselage station 605.98 (Ref. Figure 209). (14) Connect rudder trim tab actuator rod outer double clevis end (3) to the rudder tab horn (6) using bolt, washer and nut (2) (Ref. Figure 210). NOTE: Tighten the large nut of the double clevis first, then the small nut. (15) Remove cable blocks (3) from trim tab cables (2) in the aft fuselage area (Ref. Figure 208). (16) Rig the rudder trim tab system (Ref. 27-20-07, RUDDER TRIM TAB SYSTEM RIGGING). (17) Secure the turnbuckles with the safety clips. (18) Install servo cover over the lower pulley bracket at canted fuselage station 605.98, if installed. (19) Install vertical stabilizer access panel 24 (Ref. Chapter 6-50-00, STABILIZER ACCESS PANELS). (20) Install aft fuselage access panel 7 (Ref. Chapter 6-50-00, FUSELAGE ACCESS PANELS). (21) Install aft floorboard access panel 25 (Ref. Chapter 6-50-00, FLOOR ACCESS PANELS). (22) Remove the red tag from the rudder trim tab control knob.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

1. AFT PRESSURE BULKHEAD 2. RUDDER TRIM TAB CABLES 3. CABLE BLOCK 4. PULLEYS

A

1

2 3 4

FWD

DETAIL

A

UC27B 043139AB.AI

Figure 208 Rudder Trim Tab Cable Block

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

1. CABLE STOP 2. CABLE STOP GUARD 3. CABLE RETAINING PIN 4. RUDDER TRIM TAB CABLES 5. TURNBUCKLE 6. LOWER PULLEY BRACKET

A

FWD

2

1

5

4

6

3

(SERVO COVER REMOVED FOR CLARITY, IF INSTALLED) DETAIL

Figure 209 Cable Stop Guard and Retaining Pin

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A

UC27B 043140AA.AI

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

1. RUDDER TRIM TAB 2. BOLT, WASHER AND NUT 3. DOUBLE CLEVIS 4. RUDDER 5. VERTICAL STABILIZER 6. RUDDER TRIM TAB HORN

A

5

4

3 2 1

6

DETAIL

A

UC27B 043141AB.AI

Figure 210 Rudder Trim Tab Double Clevis

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

1. ELEVATOR SYSTEM CABLE 2. CABLE FAIRLEAD 3. RUDDER TRIM TAB ACTUATOR 4. BOLTS AND WASHERS 5. SHIM ( IF INSTALLED) 6. RUDDER TRIM TAB CABLES

A

3

4

2 1

5

6

FWD

DETAIL

A

Figure 211 Rudder Trim Tab Actuator Removal

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UC27B 043142AB.AI

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

1. RUDDER TRIM TAB ACTUATOR 2. ACTUATOR HOUSING 3. ACTUATOR ROD 4. TAB WASHER 5. ROD END JAM NUT 6. ROD END 7. OUTER CLEVIS 8. AFT CLEVIS JAM NUT 9. SMALL JAM NUT 10. PUSH-PULL ROD 11. FORWARD CLEVIS JAM NUT 12. BOLT, WASHER AND NUT 13. BUSHING 14. INSPECTION HOLE 15. LEFT CABLE 16. RIGHT CABLE 17. BOLT, WASHER AND NUT

18. GUIDE 19. WASHER, NUT 20. BOLT, WASHER 21. ELEVATOR CABLE FAIRLEAD BRACKET 22. RUDDER TRIM TAB ACTUATOR HOUSING 23. CABLE DRUM END CAP BEARING 24. RUDDER TRIM TAB ACTUATOR END CAP 25. DRUM AND CABLE

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B 1 2

6 4 5

3

20

14 13

12

11 10

DETAIL

9

A 15

C

8

7

16

18

17

22 23 24

21

20

19 DETAIL

B

25 DETAIL

C

UC27B 043144AB.AI

Figure 212 Rudder Trim Tab Actuator Cable Drum

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A B C D FWD DETAIL

LARGE BEARING RACE SIDE

A CENTER OF CABLE

CABLE LOCK PIN (PART OF CABLE ASSEMBLY)

SIDE WITH LEFT-HAND THREADS TERMINAL END.

SIDE WITH RIGHT-HAND THREADS TERMINAL END.

START WRAPPING OUTSIDE GROOVES MOVING INWARD 4 1/4 TURNS. BOTH SIDES DETAIL

B DETAIL SIDE WITH RIGHT-HAND THREADS TERMINAL END.

SIDE WITH LEFT-HAND THREADS TERMINAL END.

C

DRUM FULLY WRAPPED DETAIL

D

THIS INFORMATION IS FOR THE MODEL 1900/1900C ONLY. DO NOT USE ON OTHER MODEL 1900 SERIES AIRCRAFT, BECAUSE THERE ARE CRITICAL DIFFERENCES.

Figure 213 Rudder Trim Tab Actuator Cable Drum Wrapping

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UC27B 043245AD.AI

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

FLIGHT CONTROLS RUDDER TRIM TAB INDICATOR MAINTENANCE PRACTICES

27-20-06 200200

1. RUDDER TRIM TAB CONTROL A. Removal (1) Remove the right flight compartment pedestal side panel. (2) Position the rudder trim control dial (2) to 0 (Ref. Figure 201). (3) Loosen the setscrew (7) in the rudder trim control knob (1). (4) Remove the rudder trim control knob (1). (5) Remove the snap ring (8). (6) Pull up and remove the rudder trim control dial (2). (7) Remove the gear (3). (8) Through the right side pedestal, remove the safety wire from the upper portion of the universal joint (6) securing the upper pin (5). (9) Remove the upper pin (5) from the universal joint (6). (10) Remove the shaft (4) from the pedestal.

B. Installation (1) Install the shaft (4) into the pedestal (Ref. Figure 201). (2) Through the right side pedestal, Install the upper pin (5) into the universal joint (6) and safety wire. (3) Install the large gear (3). (4) Install the rudder trim control dial (2). (5) Install the snap ring (8). (6) Install the rudder trim control knob (1). Apply Loctite 242 to the setscrew (7) threads and tighten. (7) Perform the RUDDER TRIM TAB ADJUSTMENT procedure in this section. (8) Install the right flight compartment pedestal side panel.

C. Indicator- Adjustment The trim tab indicator should correspond to the tab position. If proper indication cannot be achieved by adjusting the tab clevis rod, adjustment of the indicator can be accomplished as follows: (1) Remove aft fuselage access panel 8 (Ref. Chapter 6-50-00, FUSELAGE ACCESS PANELS). (2) Install rig pin (2) (19, Table 7, 91-00-00) in the rudder aft torque tube (1) (Ref. Figure 202).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (3) Loosen the setscrew (7) in the side of the rudder trim control knob (1) (Ref. Figure 201). (4) Remove the rudder trim control knob (1). (5) Remove snap ring (8). (6) Pull the rudder trim control dial (2) up until it turns freely. (7) Turn the dial (2) to the desired position and lower to engage the gear (3). (8) Install the snap ring (8). (9) Install the rudder trim control knob (1). Apply Loctite 242 to the setscrew (7) threads and tighten. (10) Remove the rig pin (2) from the rudder aft torque tube (1) (Ref. Figure 202). (11) Install aft fuselage access panel 8 (Ref. Chapter 6-50-00, FUSELAGE ACCESS PANELS).

D. Inspection (1) Position the rudder trim control dial (2) to 0 (Ref. Figure 201). (2) Loosen the setscrew (7) in the rudder trim control knob (1). (3) Remove the rudder trim control knob (1). (4) Remove the snap ring (8). (5) Pull up and remove the rudder trim control dial (2). (6) Inspect the gear (3) and the shaft (4) for distortion and missing teeth. (7) Install the rudder trim control dial (2). (8) Install the snap ring (8). (9) Install the rudder trim control knob (1). Apply Loctite 242 to the setscrew (7) threads and tighten.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

1. RUDDER TRIM CONTROL KNOB 2. RUDDER TRIM CONTROL DIAL 3. GEAR 4. SHAFT 5. PIN UPPER 6. UNIVERSAL JOINT 7. SETSCREW 8. SNAP RING

A

1 8 2

7 3

4

5 6 DETAIL

A

UC27B 041808AA.AI

Figure 201 Rudder Trim Tab Control

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

A

1. RUDDER AFT TORQUE SHAFT 2. RIG PIN

1 2

VIEW

NOTE: EARLIER VERSIONS OF THE TORQUE SHAFT SECTOR MAY NOT HAVE LIGHTENING HOLES INSTALLED.

A

UC27B 041711AC.AI

Figure 202 Rudder Aft Torque Tube Installation

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27-20-06

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

FLIGHT CONTROLS RUDDER TRIM TAB CONTROL SYSTEM MAINTENANCE PRACTICES

27-20-07 200200

1. PROCEDURES A. Rigging WARNING: Perform all Steps of this procedure, in the order listed. Do not skip any Steps of this procedure. Failure to do so may result in injury to personnel and damage to equipment. The rudder control system must be properly rigged before the rudder trim tab system can be rigged. Failure to do so may result in injury to personnel and damage to equipment. CAUTION: Where the cables pass through structure, the areas of possible contact between the control cables and adjacent structure must be protected with protective elements such as grommets, rub strips, blocks or guide fairings. Where contact of the control cables does occur with the protective elements, a force no greater than eight (8) ounces shall be required to move the cables to a position of no contact. At no time should flight control cables contact metal structure with the protective elements removed. (1) Remove left side passenger compartment seats as necessary to access floorboard panels for rudder trim tab cable turnbuckle access (Ref. Chapter 25-20-00, PASSENGER SEAT REMOVAL). (2) Remove left side passenger compartment carpet as necessary to access floorboard panels for rudder trim tab cable turnbuckle access (Ref. Chapter 25-20-01, CARPET REMOVAL AND INSTALLATION). (3) Remove left side passenger compartment floorboard number 16H panel, for rudder trim tab cable turnbuckle access (Ref. Chapter 6-50-00, FLOOR ACCESS PANELS). (4) Remove the aft fuselage access panels 7 and 8 (Ref. Chapter 6-50-00, FUSELAGE ACCESS PANELS). (5) Remove vertical stabilizer panel 24, if required to shim rudder trim tab actuator (Ref. Chapter 6-50-00, STABILIZER ACCESS PANELS). (6) Airplanes equipped with Mechanical Steering (Ref. Chapter 32-50-00, MECHANICAL STEERING - MAINTENANCE PRACTICES) must have the steering system as if the airplane was in flight. The airplane must be placed on jacks and the steering disconnect actuator must be in the extended position. This is accomplished by performing the following: (a) Perform THREE-POINT JACKING (PREFERRED PROCEDURE) procedure (Ref. Chapter 7-10-00). (b) Apply electrical power to the airplane. The mechanical steering disconnect actuator should extend. (c) Verify that the mechanical steering disconnect actuator located on the left side of the nose landing gear wheel well is in the extended position. (d) Remove electrical power from the airplane.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (7) Perform RUDDER TRAVEL BOARD INSTALLATION procedure (Ref. 27-00-02). (8) Install rig pin (2) (7, Table 1, 27-00-00) in the hole in the rudder aft torque shaft (1) assembly. Using minimum force, try to manually move the rudder to verify proper rig pin installation. (Ref. Figure 204). (9) Perform RUDDER TRIM TAB TRAVEL BOARD INSTALLATION procedure (Ref. 27-00-02). (10) Through the left side aft fuselage floorboard panels remove the tie straps (1) securing the conduit tubes (5) over the rudder trim cables (Ref. Figure 205). (11) Slide the conduit tubes (5) aft as required to gain access to the turnbuckles (2 and 8). (12) Check the rudder trim cable tension by performing the following Steps: NOTE: Do not measure the tension when the temperature is varying rapidly or with the airplane located in direct sunlight. (a) Allow the temperature of the airframe to stabilize before measuring and adjusting the cable tension. (b) Measure the temperature in the compartment next to the rudder trim tab cables near the turnbuckles in the aft passenger compartment. (c) Refer to Rudder Trim Tab Cable Tension Graph Figure 201 and read the pounds of tension required for the measured temperature. It is permissible to measure tension at any point of the rudder trim system. (d) Position a cable tensiometer (4, Table 1, 27-00-00) at least three inches from the turnbuckles and measure the cable tension of both cables. Cable diameter is noted in Figure 201. NOTE: Cable tension tolerance is +3/ -2 lbs of the tension found in Figure 201. (13) If no adjustment is needed, proceed to Step (14). If adjustment is needed, perform the following Steps: (a) Remove the safety clips from the rudder trim tab cable turnbuckles. WARNING: If cable tension at any time is below 5 pounds, check all rudder trim system cable drums and pulleys for proper cable engagement. (b) Adjust the cable tension by adjusting turnbuckles at fuselage station 390 or the aft fuselage area equally to the tension value needed at the current temperature found in Figure 201. NOTE: Each turnbuckle barrel has a groove at one end to identify the left-hand threaded end. (c) Using the rudder trim control knob on the pedestal cycle the trim tab system three times to equalize the cable tension throughout the system and check the cable tension. If tension is out of limits, repeat Step (13) (b). (d) Install the safety clips to the rudder trim tab cable turnbuckles. (14) Slide the conduit tubes (5) forward as required to cover the rudder trim tab cable turnbuckles (2 and 8) (Ref. Figure 205).

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27-20-07

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (15) Secure the conduit tubes (5) with tie straps (1), forward of bracket (6) and aft of bracket (7). (16) Remove the rig pin from the rudder aft torque shaft (1). (17) Using minimum force, move the rudder to the left or right position to access the rudder trim tab actuator housing and rod ends. (18) Using the rudder trim control knob on the pedestal move the trim tab cables to obtain 2.50 inches from the rudder trim actuator housing to the center of the rod end. (Ref. Figure 202). (19) Move rudder back to neutral (0° deflection). (20) Install rig pin (2) (7, Table 1, 27-00-00) in the rudder aft torque shaft (1) assembly. Using minimum force try to manually move the rudder to verify proper rig pin installation. (Ref. Figure 204). NOTE: With the dimension required in Step (18) properly set, the rudder rig pin installed and the rudder at neutral, the rudder trim tab must also be at neutral (0° deflection). (21) Check the rudder trim tab for neutral (0° deflection). (22) If no adjustment is required, proceed to Step (23). If adjustment is required, perform the following Steps: (a) Loosen large jam nut (9) on clevis (10) and remove nut, washer and bolt. Discard nut (Ref. Figure 202). (b) Disconnect the rod clevis (10) from the rudder trim tab horn (4). NOTE: The double clevis ends on the trim tab actuator push-pull rod is designed to tighten the outer clevis with a binding action against the clevis bolt and the inner clevis, which removes free play of the bolt in the hole. Loosen the outer clevis first (the larger nut) before removing the clevis bolt. (c) Adjust the push-pull rod (7) by loosening the jam nut (8) from the clevis (10) and rotating the clevis (10) to bring the rudder trim tab to 0° position. (d) After adjustment, the push-pull rod stud must be visible through the inspection hole in the clevis (10). (e) Tighten the jam nut (8), taking care not to alter push-pull rod (7) adjustment. (f) Temporarily connect the rod clevis (10) to the rudder trim tab horn (4) using bolt, washer and nut. (23) Rotate the rudder trim control knob on the pedestal to the full nose left stop. Verify the cable stop (2) contacts the cable stop guard (3) located in the aft fuselage area (Ref. Figure 203). (24) Using the rudder trim tab travel board, check the rudder trim tab for a deflection of 15° to 16.5° right from neutral. (25) Rotate the rudder trim control knob on the pedestal to the full nose right stop. Verify the cable stop (2) contacts the cable stop guard (3) located in the aft fuselage area. (26) Using the rudder trim tab travel board, check the rudder trim tab for a deflection of 15° to 16.5° left from neutral.

27-20-07

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (27) If no adjustment is required, proceed to Step (28). If adjustment is required, perform the following Steps: (a) Remove safety wire (1) from the cable stop(s) (2) located in the aft fuselage area. (b) Loosen the cable stop(s) (2) and adjust as needed to achieve proper deflection. (c) Tighten cable stop(s) (2) and torque 40 to 45 inch-pounds and safety wire (1). NOTE: Wrap five turns on each end of safety wire on end opposite to cable stop contact area. (d) Repeat Steps (23) through (27). (28) Using the rudder trim control knob on the pedestal move the rudder trim tab to 0° (neutral) position, per the travel board. (29) Using rudder travel board with the rudder trim travel board, perform the following: (a) Remove the rig pin from the rudder aft torque shaft (1). (b) Disconnect the trim tab from the push-pull rod and align the trim tab with the rudder. (c) Move the rudder surface to the full left position. Measurement should be 25° +1°/ -0° deflection. (d) Connect the trim tab and check measurement on the trim travel board. The trim tab must be 10° ± 1° left at full left rudder. (e) Disconnect the trim tab from the push-pull rod and align the trim tab with the rudder. (f) Move the rudder surface to the full right position. Measurement should be 25° +1°/ -0° deflection. (g) Connect the trim tab and check measurement on the trim travel board. The trim tab must be 10° ± 1° right at full right rudder. (30) If no adjustment is required, proceed to Step (31). If adjustment is required, perform the following Steps (a) A maximum of one laminated shim (5) (P/N 101-524078-3) may be placed between the adapter and the base of the actuator (Ref. Figure 206). (b) Adjustment of the actuator is accomplished by removing laminations from the shim as required to obtain the correct anti-servo trim tab position. NOTE: To increase the anti-servo, install shim under the left side of the actuator flange. To decrease anti-servo, install the shim under the right side of the actuator flange. Add or remove shim laminations under only one side of the actuator. (c) Repeat Steps (28) through (30). NOTE: The bolt must be installed with the head of the bolt up. (31) Install the rod clevis (10) to the rudder trim tab horn (4) using bolt, washer and a new nut.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (32) Tighten large aft jam nut (9) to aft clevis (10). (33) Install a rig pin (2) (7, Table 1, 27-00-00) in the rudder aft torque shaft (1) (Ref. Figure 204). (34) Rotate the rudder trim tab control knob on the pedestal so the trim tab is positioned at neutral (0° deflection). The 0 mark on the trim tab indicator must align with the triangle mark on the pedestal edgelighted panel. If required, perform the RUDDER TRIM TAB INDICATOR ADJUSTMENT procedure (Ref. 27-20-06). (35) Turn the rudder trim tab control knob on the pedestal counterclockwise, verify that the trim tab moves right. (36) Turn the rudder trim tab control knob on the pedestal clockwise, verify that the trim tab moves left. (37) Remove the rig pin from the rudder aft torque shaft (1). (38) Perform the RUDDER OPERATIONAL CHECK procedure (Ref. 27-20-02). (39) Remove the rudder trim tab travel board. (40) Perform the RUDDER TRAVEL BOARD REMOVAL procedure (Ref. 27-00-02). (41) Perform LOWERING THE AIRPLANE AFTER THREE-POINT JACKING procedure (Ref. Chapter 7-10-00). (42) Install vertical stabilizer panel 24, if removed (Ref. Chapter 6-50-00, STABILIZER ACCESS PANELS). (43) Install the aft fuselage access panels 7 and 8 (Ref. Chapter 6-50-00, FUSELAGE ACCESS PANELS). (44) Install left side passenger compartment floorboard panel number 16H (Ref. Chapter 6-50-00, FLOOR ACCESS PANELS). (45) Install left side passenger compartment carpet removed for floorboard access (Ref. Chapter 25-20-01, CARPET REMOVAL AND INSTALLATION). (46) Install left side passenger compartment seats removed for floorboard access (Ref. Chapter 25-20-00, PASSENGER SEAT INSTALLATION).

B. Operational Check (1) Rotate the rudder trim tab control knob counter clockwise and make sure that the rudder trim tab moves right smoothly with no unusual noise or binding. (2) Rotate the rudder trim tab control knob clockwise and make sure that the rudder trim tab moves left smoothly with no unusual noise or binding. (3) If requirements are not met, perform the RUDDER TAB CONTROL RIGGING procedure in this section.

C. Cable Tension Check (1) Remove aft fuselage access panel 7 (Ref. Chapter 6-50-00, FUSELAGE ACCESS PANELS).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (2) Rotate the rudder trim tab control knob to the full left and full right position three cycles to equalize system tension. (3) Check the cable tension by performing the following Steps: NOTE: Do not measure the tension when the temperature is varying rapidly or with the airplane located in direct sunlight. (a) Allow the temperature of the airframe to stabilize before measuring and adjusting the cable tension. (b) Measure the temperature in the compartment next to the rudder trim tab control cables. (c) Refer to Rudder Trim Tab Cable Tension Graph Figure 201 and read pounds of tension for the measured temperature. (d) Position a cable tensiometer (4, Table 1, 27-00-00) on the rudder trim tab cables and measure the cable tension. Cable diameter is noted in Figure 201. NOTE: Trim tab cable tension tolerance is +3/ -2 pounds. (e) If cable tension does not require adjustment, proceed to Step (4). If cable tension requires adjustment, perform the RUDDER TRIM TAB CONTROL RIGGING procedure in this section. (4) Install fuselage access panel 7 (Ref. Chapter 6-50-00, FUSELAGE ACCESS PANELS). (5) Perform the RUDDER TRIM TAB OPERATIONAL CHECK procedure in this section.

D. Functional Check (1) Remove aft fuselage access panel 8 (Ref. Chapter 6-50-00, FUSELAGE ACCESS PANELS). (2) Install rig pin (2) (7, Table 1, 27-00-00) in the rudder aft torque shaft (1) (Ref. Figure 204). (3) Perform the RUDDER TRAVEL BOARD INSTALLATION procedure (Ref. 27-00-02). Verify that the rudder is at 0°. If the rudder is not at 0°, perform the RUDDER CONTROL SYSTEM RIGGING procedure (Ref. 27-20-02). (4) Perform the RUDDER TRIM TAB TRAVEL BOARD INSTALLATION procedure (Ref. 27-00-02). (5) Rotate the trim tab control knob counter clockwise to the full left position and make sure that the trim tab moves to the full right position 15° to 16.5° smoothly with no unusual noise or binding. (6) Rotate the trim tab control knob clockwise to the full right position and make sure that the trim tab moves to the full left position 15° to 16.5° smoothly with no unusual noise or binding. (7) If rudder trim tab requires adjustment, perform the RUDDER TRIM TAB CONTROL RIGGING procedure in this section. (8) Remove rig pin (1) from the rudder aft torque shaft (1) (Ref. Figure 204).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

(9) Install the aft fuselage panel 8 (Ref. Chapter 6-50-00, FUSELAGE ACCESS PANELS). (10) Remove the rudder trim tab travel board. (11) Perform the RUDDER TRAVEL BOARD REMOVAL procedure (Ref. 27-00-02).

TOLERANCE: +3, -2 POUNDS OF TENSION 1/16" DIAMETER RUDDER TAB CABLE TENSION GRAPH

POUNDS OF TENSION

40

30

20

10

0 UC27B 042490AA.AI

Figure 201 Rudder Trim Tab Cable Tension Graph

27-20-07

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1. RUDDER TRIM TAB 2. RUDDER 3. VERTICAL STABILIZER 4. RUDDER TRIM TAB CONTROL HORN 5. RUDDER TRIM TAB ACTUATOR 6. BOLT, WASHER AND NUT 7. PUSH-PULL ROD 8. ADJUSTMENT STUD 9. AFT CLEVIS JAMNUT 10. AFT CLEVIS 2 11. FORWARD CLEVIS 12. FORWARD CLEVIS JAMNUT

A

3

1

B 4

DETAIL

A 5 * 2. 50 S E H C IN

6

7

12 10

9

8

DETAIL

B

Figure 202 Rudder Trim Tab Actuator Adjustment

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11

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UC27B 042197AA.AI

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

1. SAFETY WIRE 2. CABLE STOP 3. CABLE STOP GUARD 4. TRIM TAB CABLES 5. TURNBUCKLE

A

FWD

3

2

1

5

4

(SERVO COVER REMOVED FOR CLARITY, IF INSTALLED) DETAIL

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UC27B 042199AA.AI

Figure 203 Rudder Trim Cable Stop

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

A

1. RUDDER AFT TORQUE SHAFT 2. RIG PIN

1 2

VIEW

NOTE: EARLIER VERSIONS OF THE TORQUE SHAFT SECTOR MAY NOT HAVE LIGHTENING HOLES INSTALLED.

A

UC27B 041711AC.AI

Figure 204 Rudder Aft Torque Shaft Rig Pin Installation

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1. TIE STRAPS 2. TURNBUCKLE 3. LEFT - HAND THREADS TERMINAL END 4. RIGHT - HAND THREADS TERMINAL END 5. CONDUIT TUBES 6. BRACKET 7. BRACKET 8. TURNBUCKLE

1

A

5

1

VIEW LOOKING DOWN DETAIL 5

A

FWD

FS 408.25 (REF) 4

8

7

2

FS 378.25 (REF) 3

VIEW LOOKING DOWN (CONDUIT TUBES MOVED AFT AND BLOCK TEMPORARILY INSTALLED) DETAIL

6

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Figure 205 Rudder Trim Cable Turnbuckle Location

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1. ELEVATOR SYSTEM CABLE 2. CABLE FAIRLEAD 3. RUDDER TRIM TAB ACTUATOR 4. BOLTS AND WASHERS 5. SHIM ( IF INSTALLED) 6. RUDDER TRIM TAB CABLES

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Figure 206 Rudder Trim Tab Actuator Removal

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UC27B 043142AB.AI

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

FLIGHT CONTROLS FLIGHT CONTROLS ASSIST SYSTEMS DESCRIPTION AND OPERATION

27-21-00 00

1. GENERAL A. Stability Augmentation System (Yaw Damper) The Stability Augmentation Computer System is installed on airplanes not equipped with an autopilot, to aid the pilot with directional control (yaw damper). The yaw damper portion of this computer system senses changes in compass headings, which it converts into directions sent to the electric rudder servo. The electric rudder servo in turn operates the rudder control cables, moving the rudder in the appropriate direction to stabilize the yaw axis of the airplane. To activate the yaw damper portion of this system, the pedestal mounted yaw damper control switch must be moved to the YAW DAMP position. Activation of this system also requires that the airplane's MASTER SWITCH-BATT switch be in the ON position and either No. 1 or No. 2 inverter be selected. The system may be disengaged by momentarily pressing the control wheel mounted YD DISC switch. A test circuit, which tests the functional integrity of the yaw damper system, is an integral function of the computer and is energized by pressing the YD ENGAGE annunciator light on the instrument panel. A positive test of the yaw damper system is indicated when the YD ENGAGE annunciator is extinguished and the yaw damper system is disengaged. In order to engage the yaw damper system after testing, the yaw damper switch must then be returned to the YAW DAMP position. If the airplane is equipped with a Hawker Beechcraft Corporation installed autopilot, the autopilot will incorporate the yaw damper capabilities. Refer to the appropriate autopilot flight manual supplement for details which may vary from the above description. Refer to Figure 1 for Yaw Damper System Wiring Diagram.

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Figure 1 Yaw Damper System Wiring Diagram

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

FLIGHT CONTROLS FLIGHT CONTROL ASSIST SYSTEMS MAINTENANCE PRACTICES

200200

1. PROCEDURES Information concerning location, removal and installation instructions for the electrical components of this system are found in Chapter 39.

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FLIGHT CONTROLS ELEVATOR MAINTENANCE PRACTICES

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1. PROCEDURES A. Removal NOTE: Removal of the elevator static wicks will prevent accidental damage (Ref. Chapter 23-60-00, STATIC DISCHARGE - MAINTENANCE PRACTICES). (1) Remove access panels 15 and 18 (Ref. Chapter 6-50-00, STABILIZER ACCESS PANELS). (2) Remove access panel 34 on the left and right side of the vertical stabilizer (Ref. Chapter 6-50-00, STABILIZER ACCESS PANELS). (3) Move the elevator surface (1) and install a rig pin (3) (7, Table 1, Chapter 27-00-00) in the elevator aft bellcrank through vertical stabilizer (4). Look at the opposite side of the vertical stabilizer to verify that the rig pin has protruded. Using minimum force try to manually move the elevators up and down to verify that the rig pin is correctly installed and that the elevators do not move (Ref. Figure 204). (4) Remove elevator mount bolt access plugs (2) from the top and bottom of the elevator surface (1) (Ref. Figure 201). NOTE: The double clevis end (3) on trim tab push-pull rod (4) is designed to tighten the outer clevis with a binding action against clevis bolt (6) and the inner clevis, which prevents turning of the bolt in the hole. Loosen the outer clevis first, (the larger nut) (5) before removing the clevis bolt (6) (Ref. Figure 203). (5) Loosen clevis jam nut (5) and remove the bolt, nut and washer (6). Disconnect push-pull rod (4) from the elevator trim tab control horn (2). Do not change the adjustment of the push-pull rod (4) end. (6) Disconnect the three bonding jumpers (9) between the elevator (1) and the horizontal stabilizer hinge brackets (8) (Ref. Figure 201). NOTE: Observe the number and position of washers (7) and the position of the safety wire bracket (2) in order to facilitate installation (Ref. Figure 202). (7) Remove safety wire (3), bolt (4), safety wire bracket (2) and washers (7), which attach the torque tube (9) to the support bracket (8). (8) Remove the nuts and washers (5) from the mount bolts (3) (Ref. Figure 201). (9) Remove the bolt, nut, and washer (10) and disconnect elevator push-pull rod (5) from the elevator control horn (6) (Ref. Figure 202). NOTE: Note the length and position of each bolt (3) for reference upon installation. Bushings (11) may stay with the bolts during removal; if they do, note their length and position (Ref. Figure 201). (10) With assistance, remove the mount bolts (3) from brackets (4) and carefully remove the elevator (1) from the horizontal stabilizer.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (11) Perform the ELEVATOR INBOARD, CENTER AND OUTBOARD HINGE INSPECTION procedure in this section. Repair or replace parts as required.

B. Installation WARNING: Any time the push-pull rods are installed or adjusted, the inspection hole near the ends of the push-pull rods must be checked to ascertain that the threads of the end fittings are visible. NOTE: Any repair, modification, painting, or replacement of the elevator or tab will require balancing (Ref. Chapter 55-20-00, BALANCING PROCEDURES). If the same elevator is being installed that was removed, inspected in accordance with the ELEVATOR INBOARD, CENTER AND OUTBOARD HINGE INSPECTION procedure in this section and no maintenance has been performed upon this elevator, then proceed to Step (2). (1) Perform the ELEVATOR INBOARD, CENTER AND OUTBOARD HINGE INSPECTION procedure in this section. Repair or replace parts as required. (2) If hinge bolts (6), nuts, washers, and cotter pin (10) were loosened or replaced, they must be tightened to a torque of 40 to 50 inch-pounds. Start the procedure at the inboard hinge and work outboard to the outboard hinge. After torquing, the bolt shall not rotate under 15 to 20 inch-pounds of torque applied to the bolt. It is permissible to tighten to 60 inch-pounds if cotter pin holes do not align. Do not loosen the nut to align cotter pin holes. If torque requirements cannot be met, add or remove washers as required and repeat this Step (Ref. Figure 201). NOTE: The bolts (3) must be installed with the heads up. Ensure the bolts (3) and bushings (11) are installed in the position noted in the ELEVATOR REMOVAL procedure. (3) With assistance, guide the elevator (1) into position on the horizontal stabilizer. Ensure all hinges are aligned. Install the attaching bolts (3), bushings (11), washers and nuts (5). Do not tighten at this time. (4) Install bolt (4), washers (7), and safety wire bracket (2). A gap of 0.060 to 0.092 inch must be maintained between the support bracket (8) and the safety wire bracket (2). Use AN960-516 or AN960-516L washers as required to control the noted gap. It is permissible to install one AN960-516 or AN960-516L washer between the support bracket (8) and the torque tube (9); if this is done, a NAS1305-23H bolt must be used in place of the NAS1305-22H bolt. When the above conditions are satisfied, torque the bolt to 56 to 78 inch-pounds (Ref. Figure 202). (5) Torque the inboard hinge bracket mount bolt first and work outboard to the outboard hinge bracket. Torque the mount bolts (3) 50 to 80 inch-pounds. After torquing, the bolt shall not rotate when 25 to 35 inch-pounds of torque is applied to the bolt (3). If requirements can not be met replace the mount bolt, nut, washer and bushing (3, 5 and 11) and repeat this Step (Ref. Figure 201). (6) Connect the three bonding jumpers (9) between the elevator (1) and the horizontal stabilizer hinge bracket (8). (7) Perform the ELECTRICAL BONDING CHECK procedure (Ref. Chapter 20-00-01). (8) Align the elevator push-pull rod (5) with the elevator control horn (6) and install the bolt, nut, and washer (10) (Ref. Figure 202).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (9) Align the elevator trim tab double clevis end (3) with the trim tab control horn (2) and install the bolt, nut, and washer (6) and tighten the clevis jam nut (5) (Ref. Figure 203). (10) With full elevator down and elevator trim tab at trailing edge full down, check that there is clearance between the elevator trim tabs and the top of the rudder at full left and full right rudder. (11) With assistance, move the elevator from full down to full up as many times as necessary to assure clearance between the elevators and the horizontal stabilizer. Inspect for 0.137 inch minimum clearance between the leading edge (nose) of the elevators and any part of the horizontal stabilizer trailing edge at any position of the elevators. (12) Move the elevators from full down to full up as many times as necessary to assure clearance between the elevators and the aft fairing. Inspect for 0.12 inch minimum clearance between the inboard end of the elevators and the aft fairing (tail cone) at any position of the elevators. (13) Move the elevator from full down to full up as many times as necessary to assure clearance between the elevators and the horizontal stabilizer. Inspect the outboard end of the elevators where the balance weights are installed. There must be clearance between the elevator and the horizontal stabilizer at any position of the elevators. Examine the leading edge and the inboard side of the balance weights for protruding screw heads. (14) Perform the ELEVATOR FREEPLAY CHECK procedure in this section. (15) Perform the ELEVATOR TRIM TAB FUNCTIONAL CHECK procedure (Ref. 27-30-05). (16) Perform the ELEVATOR FUNCTIONAL CHECK procedure (Ref. 27-30-02). If the elevators do not pass the ELEVATOR FUNCTIONAL CHECK, perform the ELEVATOR CONTROL SYSTEM RIGGING procedures (Ref. 27-30-02). CAUTION: Carefully lower the elevator surface. Do not allow the elevator to free fall to the down position. This could cause damage to the elevator system. (17) Remove the elevator aft bellcrank rig pin (3) (Ref. Figure 204). (18) Install access panel 34 on the left and right side of the vertical stabilizer (Ref. Chapter 6-50-00, STABILIZER ACCESS PANELS). (19) Install access panels 15 and 18 (Ref. Chapter 6-50-00, STABILIZER ACCESS PANELS).

C. Freeplay Check NOTE: Movement or jarring of the airplane will invalidate the elevator freeplay check. The airplane should be placed in a hangar and no personnel in or on the airplane during the freeplay check. (1) Remove access panel 34 on the left and right side of the vertical stabilizer (Ref. Chapter 6-50-00, STABILIZER ACCESS PANELS). (2) Move the elevator surface (1) and install a rig pin (3) (7, Table 1, 27-00-00) in the aft elevator bellcrank through the vertical stabilizer (4). Look at the opposite side of the vertical stabilizer to verify that the rig pin has protruded. Using minimum force try to manually move the elevators up and down to verify that the rig pin is correctly installed and that the elevators do not move (Ref. Figure 204).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (3) Visually inspect the elevator for any damage, for security of the hinge attach points and for tightness of the actuating system. (4) Perform the ELEVATOR TRAVEL BOARD INSTALLATION procedure (Ref. 27-00-02). (5) Apply tape (for paint protection) on the top surface one inch forward of the elevator trailing edge just left or right of the elevator travel board (6). This will be the point of pressure against the elevator (3) by the push-pull scale (Ref. Figure 205). (6) Apply tape in the corresponding position on the bottom surface of the elevator (3). (7) Attach a scale (8) or a dial indicator (1) to the elevator travel board (6) to measure the up and down movement at the elevator (3) trailing edge. (8) If using the dial indicator (1) perform the following Steps: (a) Position the dial indicator (1) so the stem (2) is 0.50 inch from the trailing edge of the elevator and is depressed 0.10 inch when in contact with the elevator (3) surface initially. Turn the rotating face of the dial indicator (1) to zero. Do not reset the dial indicator during this procedure. (b) With a push-pull scale (6, Table 1, 27-00-00) against the top of the elevator (3), apply four pounds of downward load. Record the dial reading. (c) Apply four pounds upward load on the bottom surface of the elevator (3). Record the dial reading. (d) Proceed to Step (10). (9) If using the scale perform the following Steps: (a) Attach the scale (8) to the elevator travel board (6) with tape (7). (b) With a push-pull scale (6, Table 1, 27-00-00) against the top of the elevator (3), apply four pounds of downward load. Record the scale reading. (c) Apply four pounds upward load on the bottom surface of the elevator (3). Record the scale reading. (d) Proceed to Step (10). NOTE: The maximum freeplay travel limit is the total difference between the upward and downward load readings. (10) The maximum freeplay travel limit is 0.12 inch. Excess movement must be corrected. (11) If freeplay limits are exceeded, inspect all components for corrosion, cracks, wear, condition of fasteners and loose or missing rivets. Inspect the elevator horn attachment and inboard torque tube attachment for corrosion, cracks, wear, condition of fasteners and loose or missing rivets. Perform the ELEVATOR INBOARD, CENTER AND OUTBOARD HINGE INSPECTION procedure in this section. Repair or replace as required. (12) Perform Steps (3) through (11) on the opposite elevator. (13) Remove the elevator travel board.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL CAUTION: Carefully lower the elevator surface. Do not allow the elevator to free fall to the down position. This could cause damage to the elevator system. (14) Remove the elevator aft bellcrank rig pin (3) (Ref. Figure 204). (15) Install access panel 34 on the left and right side of the vertical stabilizer (Ref. Chapter 6-50-00, STABILIZER ACCESS PANELS).

D. Inspection This inspection is to be performed with the elevator installed. (1) Inspect inboard, center and outboard hinge support (4) rivets in the elevator upper and lower skin (13) and in the elevator spar (14) (Ref. Figure 201). At the outboard hinge, the two outboard spar rivets do not need to be inspected if the other rivets in the outboard hinge support meet inspection requirements. If the two outboard spar rivets need to be inspected, remove the plug in the rib just forward of the hinge support and use a borescope. Replace any loose or missing rivets (Ref. Model 1900 Airliner Series Structural Repair Manual, Chapter 55-90-07, Figure 1). (2) With assistance, move the elevator from full down to full up as many times as necessary to assure clearance between the elevators and the horizontal stabilizer. Inspect for 0.137 inch minimum clearance between the leading edge (nose) of the elevators and any part of the horizontal stabilizer trailing edge at any position of the elevators. (3) Move the elevators from full down to full up as many times as necessary to assure clearance between the elevators and the aft fairing. Inspect for 0.12 inch minimum clearance between the inboard end of the elevators and the aft fairing (tail cone) at any position of the elevators. (4) Move the elevator from full down to full up as many times as necessary to assure clearance between the elevators and the horizontal stabilizer. Inspect the outboard end of the elevators where the balance weights are installed. There must be clearance between the elevator and the horizontal stabilizer at any position of the elevators. Examine the leading edge and the inboard side of the balance weights for protruding screw heads. (5) Perform the ELEVATOR OPERATIONAL CHECK procedure (Ref. 27-30-02).

E. Inboard, Center and Outboard Hinge Inspection Perform this procedure at the time the elevator assembly is removed and installed or if the ELEVATOR FREEPLAY CHECK limits were exceeded. Inspect the rivets attaching the hinge support (4) to the elevator (Ref. Figure 201). Replace any loose or missing rivets (Ref. Model 1900 Airliner Series Structural Repair Manual, Chapter 55-90-07, Figure 1). The following information in Table 201 is for the inspection of the inboard, center and outboard hinge attaching hardware for the left and right elevator assemblies.

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Table 201 Acceptable Dimensions for Left and Right Elevator Assembly Hinge Attaching Hardware

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ITEM

DIAMETER

Mount Bolt (Vertical) (Ref. Figure 1, Item 3)

0.2492 to 0.2483 inch

Bushing (Ref. Figure 1, Item 11)

Inside 0.2495 to 0.2505 inch Outside 0.3735 to 0.3720 inch

Clevis (Ref. Figure 1, Item 7)

Vertical Hole 0.250 to 0.254 inch Horizontal Hole 0.250 to 0.254 inch Thickness 0.720 to 0.700 inch

Bearing (Ref. Figure 1, Item 12)

Inside 0.2500 to 0.2495 inch

Hinge Mount Bolt (Horizontal) (Ref. Figure 1, Item 6)

0.2492 to 0.2483 inch

Hinge Support (Ref. Figure 1, Item 4)

Hole for Bushing (11) 0.375 to 0.379 inch Bolt Hole 0.250 to 0.254 inch

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Figure 201 Elevator Hinge Installation

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0.060 - 0.092 1. ELEVATOR 2. SAFETY WIRE BRACKET 3. SAFETY WIRE 4. BOLT 5. PUSH-PULL ROD 6. CONTROL HORN 7. WASHERS 8. SUPPORT BRACKET 9. TORQUE TUBE 10. BOLT, NUT, AND WASHER

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Figure 202 Elevator Torque Tube Installation

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1. ELEVATOR TRIM TAB 2. ELEVATOR TRIM TAB CONTROL HORN 3. DOUBLE CLEVIS END 4. PUSH-PULL ROD 5. CLEVIS JAM NUT 6. BOLT, NUT, AND WASHER 7. ELEVATOR

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Figure 203 Elevator Trim Tab Connection

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1. ELEVATOR 2. RUDDER 3. ELEVATOR AFT BELLCRANK RIG PIN 4. VERTICAL STABILIZER 5. HORIZONTAL STABILIZER

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VIEW LOOKING UP LEFT HAND SIDE DETAIL

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Figure 204 (Sheet 1 of 2) Elevator Aft Bellcrank Rig Pin Installation

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

6. ELEVATOR AFT BELLCRANK 7. ELEVATOR AFT RIG PIN HOLE

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VERTICAL STABILIZER (REF)

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Figure 204 (Sheet 2 of 2) Elevator Aft Bellcrank Rig Pin Installation

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1. DIAL INDICATOR 2. INDICATOR STEM 3. ELEVATOR 4. SMALL C-CLAMP 5. BRACKET 6. ELEVATOR TRAVEL BOARD 7. TAPE 8. SCALE 1 2

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Figure 205 Elevator Freeplay Scale and Dial Indicator Installation

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

FLIGHT CONTROLS ELEVATOR CABLES MAINTENANCE PRACTICES

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1. ELEVATOR FORWARD CONTROL CABLES A. Removal (1) Attach a red tag to the control wheel with the words “Do Not Operate the Elevator System, Maintenance In Progress”. (2) Remove the pilot seat (Ref. Chapter 25-10-00, SEAT REMOVAL). (3) Remove the pilot carpet. (4) Remove flight compartment floorboards 2, 4 and 21 (Ref. Chapter 6-50-00, FLOOR ACCESS PANELS). (5) Remove the left side passenger seats (Ref. Chapter 25-20-00, SEAT REMOVAL). (6) Remove the left side passenger carpet (Ref. Chapter 25-20-01). (7) Remove the left side cabin floorboards 16A THRU 16I, 15, 24 and 25 (Ref. Chapter 6-50-00, FLOOR ACCESS PANELS). (8) Remove aft fuselage access panels 7 and 8 (Ref. Chapter 6-50-00, FUSELAGE ACCESS PANELS). (9) Insert a piece of safety wire into each elevator aft control cable (6 and 7) and secure the opposite end to the airplane structure to function as cable blocks. (10) Remove safety clips (5) from the elevator control cable turnbuckles (1 and 2). (11) Attach a tag with the words “left-hand threads terminal end” to the forward end of the turnbuckle (1). (12) Attach a tag with the words “right-hand threads terminal end” to the forward end of the turnbuckle (2). (13) Disconnect the forward elevator down control cable (4) from the turnbuckle (1) and attach a feed line to the left-hand threads terminal end (10). Label the feed line with the words “left-hand threads terminal end”. (14) Disconnect the forward elevator up control cable (3) from the turnbuckle (2) and attach a feed line to the right-hand threads terminal end (9). Label the feed line with the words “right-hand threads terminal end”. (15) Remove the pressure seals located at the aft pressure bulkhead at FS 557.50. Refer to Figure 201 for general location of the pressure seals. (16) Remove the cable retaining pins from the elevator system pulleys. Refer to Figure 201 for general location of the pulleys.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL CAUTION: With the control column pulled to the aft position, allowing the control column to free fall to the forward position can cause damage to the elevator system. (17) With assistance, remove safety wire from screw and disconnect the elevator up cable from the forward elevator bellcrank. (18) With assistance, remove safety wire from screw and disconnect the elevator down cable from the forward elevator bellcrank. NOTE: It may be necessary to remove some pulley(s) if cable passage is restricted. (19) With assistance, feed the forward elevator control cables through the pulleys, bringing the feed lines to the forward elevator bellcrank. NOTE: If cable is to be reused, take care to keep the cable clean and free from damage. Coil cable loosely no tighter than 29 inches in diameter. (20) Disconnect the feed lines and leaves them in place. (21) Remove the cables from the airplane.

B. Installation WARNING: Perform all Steps of this procedure, in the order listed. Do not skip any Steps of this procedure. Failure to do so may result in injury to personnel and damage to equipment. CAUTION: Where the cables pass through structure, the areas of possible contact between the control cables and adjacent structure must be protected with protective elements such as grommets, rub strips, blocks or guide fairings. Where contact of control cables does occur with the protective elements, a force no greater than eight (8) ounces shall be required to move the cable to a position of no contact. At no time should flight control cables contact metal structure with the protective elements removed. (1) Check cable for damage and replace cable if necessary. If a used cable is installed, cable should be cleaned with solvent (2, Table 2, 27-00-00) and then dipped in corrosion preventive compound (4, Table 2, 27-00-00). Excess should be removed by wiping with a clean rag. (2) Attach the feed line labeled “left-hand threads terminal end” to the forward elevator down control cable (4) left-hand threads terminal end (10) (Ref. Figure 202). (3) Attach the feed line labeled “right-hand threads terminal end” to the forward elevator up control cable (3) right-hand threads terminal end (9). NOTE: More than one person will be required to route the elevator fuselage cables. Take precautions to keep the cable clean and free from damage. (4) Route both forward elevator cables from the forward elevator bellcrank through the fuselage as follows (Ref. Figure 201): NOTE: It is permissible to install cable guard pins and any removed pulleys as the cable is being routed. (a) Route forward elevator down cable through the inboard pulley at FS 303.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (b) Route forward elevator down cable through the upper pulley at FS 465. (c) Route forward elevator down cable through the upper right side pulley at FS 515. (d) Route forward elevator down cable through the pressure seal retaining plate and through the right side pressure seal hole. (e) Route forward elevator down cable through the right side pulley at FS 563. (f) Route forward elevator up cable through the pulley at FS 108. (g) Route forward elevator up cable through the outboard pulley at FS 303. (h) Route forward elevator up cable through the lower pulley at FS 465. (i) Route forward elevator up cable through the lower left side pulley at FS 515. (j) Route forward elevator up cable through the pressure seal retaining plate and through the left side pressure seal hole. (k) Route forward elevator up cable through the left side pulley at FS 563. (l) Ensure all pulleys are installed and that all cable guard pins are installed in the pulley brackets. CAUTION: With the control column pulled to the aft position, allowing the control column to free fall to the forward position can cause damage to the elevator system. Do not over torque the cable locking plate attachment screws or damage to the bellcranks will occur. Maximum torque will not exceed 15 inch-pounds. (5) With assistance, connect the forward elevator down control cable to the forward elevator bellcrank upper groove by installing the cable locking plate and attaching screw. Safety wire the screw to the bellcrank. (6) With assistance, connect the forward elevator up control cable to the forward elevator bellcrank lower groove by installing the cable locking plate and attaching screw. Safety wire the screw to the bellcrank. (7) Lubricate the turnbuckles with grease (1, Table 2, 27-00-00) for corrosion protection prior to installation. (8) Remove the feed line and connect the forward elevator down control cable (4) to the left-hand threads side of the elevator down turnbuckle (1) in the aft fuselage area (Ref. Figure 202). (9) Remove the feed line and connect the forward elevator up control cable (3) to the right-hand threads side of the elevator up turnbuckle (2) in the aft fuselage area. (10) Tension the cables to prevent slack. (11) Ensure that the fuselage cables are routed properly by verifying that the cables have been routed exactly as described in Step (4). Ensure the cables are engaged in the pulley grooves and all pulley guard pins are installed. (12) Remove the safety wire from the aft elevator control cables.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (13) Remove all tape from the cables and turnbuckles. (14) Fill the pressure bulkhead cable seals with grease (1, Table 2, 27-00-00) and install the pressure seals (Ref. Figure 201). (15) Lubricate the elevator forward control cables to one inch beyond the length of travel through the pressure seal with grease (1, Table 2, 27-00-00). (16) Perform ELEVATOR CONTROL SYSTEM OPERATIONAL CHECK procedure (Ref. 27-30-02). (17) Perform ELEVATOR CONTROL SYSTEM RIGGING procedures (Ref. 27-30-02). (18) Ensure that safety clips (5) are installed on both turnbuckles (1 and 2) (Ref. Figure 202). (19) Remove the red tag from the control wheel. (20) Connect and rig the autopilot servo cables (if equipped) (Ref. Chapter 22-10-00). (21) Install aft fuselage access panels 7 and 8 (Ref. Chapter 6-50-00, FUSELAGE ACCESS PANELS). (22) Install the left side cabin floorboards 16A THRU 16I, 15, 24 and 25 (Ref. Chapter 6-50-00, FLOOR ACCESS PANELS). (23) Install the left side passenger carpet (Ref. Chapter 25-20-01). (24) Install the left side passenger seats (Ref. Chapter 25-20-00, SEAT REMOVAL). (25) Install flight compartment floorboards 2, 4 and 21 (Ref. Chapter 6-50-00, FLOOR ACCESS PANELS). (26) Install the pilot carpet. (27) Install the pilot seat (Ref. Chapter 25-10-00, SEAT REMOVAL).

2. ELEVATOR AFT CONTROL CABLES A. Removal NOTE: Take care not to drop the cables or the lead lines into the vertical stabilizer since recovery is difficult. Secure the lead lines with tape or ties when the cables are disconnected from the lead lines. (1) Attach a red tag to the control wheel with the words “Do Not Operate the Elevator System, Maintenance In Progress”. (2) Remove horizontal access panels 16, 17, 20 and 25 (Ref. Chapter 6-50-00, STABILIZER ACCESS PANELS). (3) Remove aft fuselage access panels 7 and 8 (Ref. Chapter 6-50-00, FUSELAGE ACCESS PANELS). (4) Disconnect the autopilot servo cables (if equipped) (Ref. Chapter 22-10-00). (5) Install cable blocks (3) to the elevator forward control cables (5 and 6) in the aft fuselage area aft of the elevator system pulleys at FS 563.00 (Ref. Figure 203).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (6) Remove safety clips (5) from the elevator control cable turnbuckles (1 and 2) in the aft fuselage area (Ref. Figure 202). (7) Attach a tag with the words “right-hand threads terminal end” to the aft end of the turnbuckle (1). (8) Attach a tag with the words “left-hand threads terminal end” to the aft end of the turnbuckle (2). (9) Disconnect the aft elevator down control cable (6) from the turnbuckle (1) and attach a feed line to the terminal end. Label the feed line with the words “right-hand threads terminal end”. (10) Disconnect the aft elevator up control cable (7) from the turnbuckle (2) and attach a feed line to the terminal end. Label the feed line with the words “left-hand threads terminal end”. (11) Remove safety wire from screws and disconnect the aft elevator control cables from aft elevator bellcrank. (12) With assistance, feed the aft elevator control cables up through the vertical stabilizer, bringing the feed lines to the aft elevator bellcrank. (13) Disconnect the feed lines and leave them in place. NOTE: If cable is to be reused, take care to keep the cable clean and free from damage. Coil cable loosely no tighter than 29 inches in diameter. (14) Remove the cables from the airplane.

B. Installation WARNING: Perform all Steps of this procedure, in the order listed. Do not skip any Steps of this procedure. Failure to do so may result in injury to personnel and damage to equipment. CAUTION: Where the cables pass through structure, the areas of possible contact between the control cables and adjacent structure must be protected with protective elements such as grommets, rub strips, blocks or guide fairings. Where contact of control cables does occur with the protective elements, a force no greater than eight (8) ounces shall be required to move the cable to a position of no contact. At no time should flight control cables contact metal structure with the protective elements removed. (1) Check cable for damage and replace cable if necessary. If a used cable is installed, cable should be cleaned with solvent (2, Table 2, 27-00-00) and then dipped in corrosion preventive compound (4, Table 2, 27-00-00). Excess should be removed by wiping with a clean rag. (2) Attach the feed line labeled “right-hand threads terminal end” to the aft elevator down control cable terminal end. (3) Attach the feed line labeled “left-hand threads terminal end” to the aft elevator up control cable terminal end. NOTE: More than one person will be required to route the elevator empennage cables. Take precautions to keep the cable clean and free from damage. (4) With assistance, using the feed lines, route the aft elevator cables through the vertical stabilizer.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL CAUTION: Do not over torque the cable locking plate attachment screws or damage to the bellcranks will occur. Maximum torque will not exceed 15 inch-pounds. (5) Connect the aft elevator control cables to the aft elevator bellcrank as follows: (a) Connect the aft elevator up cable to the aft side of the aft elevator bellcrank by installing the cable locking plate and attaching screw. Safety wire the screw to the bellcrank. (b) Connect the aft elevator down cable to the forward side of the aft elevator bellcrank by installing the cable locking plate and attaching screw. Safety wire the screw to the bellcrank. (6) Lubricate the turnbuckles with grease (1, Table 2, 27-00-00) for corrosion protection prior to installation. (7) Remove the feed line and connect the aft elevator down control cable (6) to the right-hand threads side of the elevator down turnbuckle (1) in the aft fuselage area (Ref. Figure 202). (8) Remove the feed line and connect the aft elevator up control cable (7) to the to the left-hand threads side of the elevator up turnbuckle (2) in the aft fuselage area. (9) Tension the cables to prevent slack. (10) Make sure that the aft elevator down cable (6) and aft elevator up cable (7) are routed properly. (11) Remove the cable blocks (3) from the elevator forward control cables (5 and 6) in the aft fuselage area from the elevator pulleys at FS 563.00 (Ref. Figure 203). (12) Remove all tape from the cables and turnbuckles. (13) Perform ELEVATOR CONTROL OPERATIONAL CHECK procedures (Ref. 27-30-02). (14) Perform ELEVATOR CONTROL SYSTEM RIGGING procedures (Ref. 27-30-02). (15) Ensure that safety clips (5) are installed on both turnbuckles (1 and 2) (Ref. Figure 202). (16) Connect and rig the autopilot servo cables (if equipped) (Ref. Chapter 22-10-00). (17) Remove the red tag from the control wheel. (18) Install aft fuselage access panels 7 and 8 (Ref. Chapter 6-50-00, FUSELAGE ACCESS PANELS). (19) Install horizontal access panels 16, 17, 20 and 25 (Ref. Chapter 6-50-00, STABILIZER ACCESS PANELS).

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AFT BELLCRANK

DOWN TENSION SPRING

A AFT ELEVATOR DOWN CABLE

C

AFT ELEVATOR UP CABLE

RIG PIN HOLE PRESSURE SEALS DETAIL

AFT BELLCRANK DETAIL

C

TURNBUCKLES DETAIL

B

A

C

FS 570 ELEVATOR AUTOPILOT SERVO FS 563

RIG PIN HOLE STOP BOLT

STOP BOLT

B

ELEVATOR DOWN CABLE

FS 515 ELEVATOR UP CABLE

FORWARD BELLCRANK DETAIL

FORWARD BELLCRANK

FORWARD ELEVATOR DOWN CABLE

D

E

FS 465

ELEVATOR UP CABLE FS 303

ELEVATOR DOWN CABLE

AFT

FORWARD ELEVATOR UP CABLE

D FS 108

OUTBOARD DETAIL

E

UE27B 051641AA.AI

Figure 201 Elevator Control System

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This Page Intentionally Left Blank

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1. ELEVATOR DOWN TURNBUCKLE 2. ELEVATOR UP TURNBUCKLE 3. FORWARD ELEVATOR UP CABLE 4. FORWARD ELEVATOR DOWN CABLE 5. SAFETY CLIPS 6. AFT ELEVATOR DOWN CABLE 7. AFT ELEVATOR UP CABLE 8. ELEVATOR UP LH THREADS TERMINAL END 9. ELEVATOR UP RH THREADS TERMINAL END 10. ELEVATOR DOWN LH THREADS TERMINAL END 11. ELEVATOR DOWN RH THREADS TERMINAL END

A 1

5

6

2

8

7

RUDDER CABLE (REF)

11 5

9 10

4 3

FWD

VIEW LOOKING INBOARD FROM LEFT HAND SIDE

DETAIL

A

UC27B 050012AA.AI

Figure 202 Elevator Control Cable Turnbuckles

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1. AFT PRESSURE BULKHEAD 2. ELEVATOR SYSTEM PULLEYS (FS 563.00) 3. CABLE BLOCKS 4. RUDDER CABLE PULLEY 5. ELEVATOR UP CABLE 6. ELEVATOR DOWN CABLE

A

6 3 2 1

5

4

DETAIL

A

Figure 203 Elevator Control Cable Block Installation

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UC27B 051643AA.AI

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

FLIGHT CONTROLS ELEVATOR CONTROL SYSTEM MAINTENANCE PRACTICES

27-30-02 200200

1. PROCEDURES WARNING: After May 2014, Mandatory Service Bulletin 27-4119 must be accomplished prior to performing the following procedures.

A. Preparation (1) Remove left lower tailcone panel 7 (Ref. Chapter 06-50-00, FUSELAGE ACCESS PANELS). (2) Remove the aft horizontal stabilizer access plate 18 (Ref. Chapter 06-50-00, STABILIZER ACCESS PANELS). (3) Remove aft fairing 15 (Ref. Chapter 06-50-00, STABILIZER ACCESS PANELS). (4) Remove the lower aft rig pin hole access panels 34 on the left and right side of the upper vertical stabilizer (Ref. Chapter 06-50-00, STABILIZER ACCESS PANELS). (5) Remove the left flight compartment seat (Ref. Chapter 25-10-00). (6) Remove the left flight compartment carpet. (7) Remove the left flight compartment inboard seat track. (8) Remove the left flight compartment floorboard panels 2, 21, and 22 (Ref. Chapter 06-50-00, FLOOR ACCESS PANELS). (9) Remove the bottom forward fuselage access panel 3 aft of the nose landing gear wheel well, if necessary for forward push-pull tube adjustment (Ref. Chapter 06-50-00, FUSELAGE ACCESS PANELS). (10) Remove the right forward pedestal side panel to access the bobweight.

B. Rigging WARNING: Perform all Steps of this procedure, in the order listed. Do not skip any Steps of this procedure. Failure to do so may result in injury to personnel and damage to equipment. The gust lock pin must not be used for rigging the Elevator control system except when performing the Gust Lock Pin Check portion of this procedure. CAUTION: Where the cables pass through structure, the areas of possible contact between the control cables and adjacent structure must be protected with protective elements such as grommets, rub strips, blocks or guide fairings. Where contact of control cables does occur with the protective elements, a force no greater than eight (8) ounces shall be required to move the cable to a position of no contact. At no time should flight control cables contact metal structure with the protective elements removed. (1) Disconnect the autopilot elevator servo cables from the elevator primary control cables, if installed (Ref. Figure 211).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (a) Disconnect the elevator servo lower cable (3) by removing the safety clips (6) and completely loosen turnbuckle (7). (b) Disconnect the elevator servo upper cable (2) from the bridle clamp (5). (c) Position the upper and lower elevator servo cables to prevent interference with the movement of the elevator control cables. NOTE: Steps (2) through (12) are for the left and right elevator rigging adjustments. (2) Perform ELEVATOR TRAVEL BOARD INSTALLATION procedure (Ref. 27-00-02). NOTE: One travel board may be used and moved from one side to the other. (3) Move the elevator surface (1) and install a rig pin (3) (7, Table 1, 27-00-00) in the aft elevator bellcrank through the vertical stabilizer (4). Using minimum force try to manually move the elevators up and down to verify that the rig pin is correctly installed. Ensure that the elevators do not move. Look at the opposite side of the vertical stabilizer to verify that the rig pin has protruded (Ref. Figure 208). (4) If using travel board P/N D807 101-610000-1/-2 at HSS 35.00, perform the following: (a) Loosen clevis jam nut (5) and disconnect the pushrods (4) from the elevator trim tab control horns (2). Do not change the adjustment of the pushrod ends (3) (Ref. Figure 209). (b) Align the trailing edge of the elevator trim tab (1) to the trailing edge of the elevator surface. WARNING: When using HSS 50.00 travel board, make measurements from the trailing edge of the elevator assemblies only. Improper rigging will occur if measurements are made from elevator trim tab trailing edge. When using HSS 35.00 travel board, ensure elevator trim tab trailing edge aligns with the elevator trailing edge before reading travel board. (5) Use the travel board(s) to measure the position of the elevators (Ref. 27-00-02, READING A TRAVEL BOARD). Both elevators must be at 0° deflection (neutral). (6) If no adjustments are required, proceed to Step (7). If adjustment is required, perform the following Steps: (a) Remove nut, washer and bolt (8) attaching the push-pull rod(s) (10) to the elevator control horn(s) (4) (Ref. Figure 204). (b) Loosen jam nut(s) (6) on the elevator pushrod(s) (10). (c) Rotate rod end(s) (9) until elevator(s) are at 0° deflection (neutral). (d) Verify that the threads of the rod end(s) (9) are visible through the inspection hole(s) (12) at the end of the pushrod(s) (10) after adjustment is completed. (e) Tighten jam nut(s) (6) on the elevator pushrod(s) (10). (f) Install nut, washer and bolt (8) attaching the push-pull rod(s) (10) to the elevator horn(s) (4). (7) Remove rig pin (3) from the aft elevator bellcrank (6) (Ref. Figure 208).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (8) Lower the elevators to the down position. With the primary down stop bolt(s) (5) contacting the elevator control horn(s), use the travel board(s) to measure the position of the elevators. The elevators must have down travel of 14° +1°/ -0° and the lower stop bolts (5) must be contacting their respective elevator control horns in the down stop bolt contact area (11) (Ref. Figure 204). (9) If no adjustment is required, proceed to Step (10). If adjustments are required, perform the following Steps: (a) Remove safety wire (7) from the stop bolt(s) (5) (Ref. Figure 204). (b) Loosen jam nut(s) (2) on the stop bolt(s) (5). (c) Adjust stop bolt(s) (5) so both elevators are at 14° +1°/ -0° and so the elevator control horn down stop bolt contact area (11) on both elevator control horns are contacting their respective down stop bolt(s) (5). (d) When adjustment is complete, tighten jam nut(s) (2) on stop bolt(s) (5) and safety wire (7) (178, Table 1, 91-00-00) the stop bolt(s). (10) With assistance, pull the cockpit control wheel (yoke) aft and hold to maintain the elevators in the up position. (11)

Use the travel board(s) to measure the position of the elevators. The elevators must have up travel of 20° +1°/ -0° and the upper stop bolts (5) must be contacting their respective elevator control horns in the up stop bolt contact area (3) (Ref. Figure 204).

(12) If no adjustments are required, proceed to Step (13). If adjustments are required, perform the following Steps: (a) Remove safety wire (7) from stop bolt(s) (5) (Ref. Figure 204). (b) Loosen jam nut(s) (2) on stop bolt(s) (5). (c) Adjust stop bolt(s) (5) so the elevators are at 20° +1°/ -0° and so the elevator control horn up stop bolt contact area (3) on both elevator control horns are contacting their respective up stop bolts (5). (d) When adjustment is complete, tighten jam nut(s) (2) on stop bolt(s) (5) and safety wire (7) (178, Table 1, 91-00-00) the stop bolt(s). (13) Move the elevator surface (1) and install a rig pin (3) (7, Table 1, 27-00-00) in the aft elevator bellcrank through the vertical stabilizer (4). Using minimum force try to manually move the elevators up and down to verify that the rig pin is correctly installed. Ensure that the elevators do not move. Look at the opposite side of the vertical stabilizer to verify that the rig pin has protruded. (Ref. Figure 208). (14) Check the cable tension by performing the following Steps: NOTE: The forward bellcrank rig pin must NOT be installed when measuring the cable tensions to calculate the combined (average) tension (Tc). (a) Allow the temperature of the airframe to stabilize before measuring and adjusting the cable tension.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL NOTE: Do not measure the tension when the temperature is varying rapidly or with the airplane located in direct sunlight. (b) Measure the temperature in the compartment next to the elevator cables (3 and 4) near the turnbuckles (1 and 2) (Ref. Figure 206). (c) Refer to Figure 202, Elevator Cable Tension Graph, and read the pounds of tension for the measured temperature. (d) Position a cable tensiometer (4, Table 1, 27-00-00) at least three inches from the turnbuckles (1 and 2) and measure the cable tension of both cables (3 and 4) (Ref. Figure 206). Cable diameter is noted in Figure 202. (e) Record the values as T up (aft cable) and T down (fwd cable). Calculate the combined (average) tension (Tc) by adding the T up and T down tensions together and divide by two. Example: Tc = (T up + T down) ÷ 2. NOTE: The combined tension (Tc) must be within a ± 8 pounds tolerance of the tension found in Figure 202. Cable tension adjustments, if needed, are performed later in this procedure. Do not adjust tension at this time. (15) Check that the forward elevator bellcrank is synchronized with the aft elevator bellcrank by performing the following Steps: (a) Install a rig pin in the forward elevator bellcrank rig pin hole (1). The rig pin should be inserted through the forward elevator bellcrank (4) with minimum forward and aft force of the control column (Ref. Figure 207). If rig pin can not be installed with minimum forward and aft force of the control column, proceed to Step (18). (b) Visually inspect to ensure that the rig pin is properly installed through the forward elevator bellcrank rig pin hole (3) to ensure that both the forward and aft elevator bellcranks are synchronized (Ref. Figure 201). (c) Raise and lower the forward rig pin while trying to move the control column forward and aft. The rig pin should move with minimum movement of the control column. (d) Remove the forward elevator bellcrank rig pin. (16) If the tension meets requirements per Step (14) (e) and the forward rig pin can be installed per Step (15) (a), proceed to Step (19) and continue checking the Elevator System. (17) If the tension does not meet requirements per Step (14) (e) and/or the forward elevator bellcrank rig pin cannot be installed per Step (15) (a), proceed to Step (18) for adjustment. WARNING: Changing the cable tension may affect other parts of the elevator system. Do not adjust the cable tension per Step (18) only. If the cable tension must be adjusted, the elevator system must be rigged per the entire ELEVATOR CONTROL SYSTEM RIGGING procedure. (18) If adjustment is needed, perform the following Steps: (a) Remove safety clips (5) from the elevator turnbuckles (1 and 2) (Ref. Figure 206).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL WARNING: When relaxing cable tension, do not relax cable tension below 10 pounds. If cable tension is below 10 pounds, check all elevator system pulleys for proper cable engagement. (b) Install the forward elevator bellcrank rig pin. If necessary, with the aft elevator bellcrank pinned, slowly relax tension of the turnbuckles until the forward elevator bellcrank rig pin can be installed. (c) With the forward and aft elevator bellcranks rig pins installed, position a cable tensiometer (4, Table 1, 27-00-00) at least three inches from the turnbuckles and measure the cable tension of both cables. Cable diameter is noted in Figure 202. Both cables should be at the nominal tension value ± 2 pounds found in Figure 202. (d) Adjust the cable tension by adjusting each turnbuckle (1 and 2) to the tension value ± 2 pounds needed at the current temperature found in Figure 202. The turnbuckles (1 and 2) must be adjusted in opposite directions to achieve the cable tensions (Ref. Figure 206). NOTE: (Information Only) The down cable tension is increased and the up cable tension is decreased at this time, because the bobweight will change the cable tension when the forward elevator bellcrank rig pin is removed. (e) Increase the forward (down) elevator cable tension by 1/2 additional turn. Decrease the aft (up) elevator cable tension by 1/2 turn. Both turnbuckles will be adjusted in the same direction. (f) Measure the elevator cable tensions to verify that the forward (down) elevator cable tension has increased and the aft (up) elevator cable tension has decreased. If this has not been achieved, return to Step (18) (d). If this has been achieved, proceed to Step (18) (g). (g) Remove the forward elevator bellcrank rig pin. (h) Position a cable tensiometer (4, Table 1, 27-00-00) at least three inches from the turnbuckles and measure the cable tension of both cables. Cable diameter is noted in Figure 202. (i) Record the values as T up (aft cable) and T down (fwd cable). Calculate the combined (average) tension (Tc) by adding the T up and T down tensions together and divide by two. Example: Tc = (T up + T down) ÷ 2. NOTE: The combined tension (Tc) must be within a ± 8 pounds tolerance of the tension found in Figure 202. (j) If the combined cable tension (Tc) is above the maximum, repeat Step (18) and set the cable tensions slightly lower while meeting the requirements of each Step. (k) If the combined cable tension (Tc) is below the minimum, repeat Step (18) and set the cable tensions slightly higher while meeting the requirements of each Step. (l) Install safety clips (5) on the turnbuckles (1 and 2) (Ref. Figure 206).

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(m) Example: The cable tension was found to be low. With the aft elevator bellcrank rig pin installed, the cable tension did not need to be relaxed when installing the forward elevator bellcrank rig pin. The temperature is measured at 59° Fahrenheit. Figure 202 shows 66 pounds as the desired tension. Each turnbuckle is turned one to two turns tighter. The tension is still low, so each turnbuckle is turned another turn tighter. With both rig pins installed, the cable tension of the forward cable is 68 pounds and the aft cable is 65 pounds, so both cables are within ± 2 pounds of the 66 pounds per Figure 202. The forward elevator cable turnbuckle is tightened 1/2 additional turn. The aft elevator cable turnbuckle is loosened 1/2 turn. The tensions of both cables are measured to verify that the forward cable tension increased and the aft cable tension decreased. The forward elevator bellcrank rig pin is removed and the cable tension is measured. The up (aft) cable measures 57 pounds and the down (forward) cable tension measures 75 pounds. The combined tension is calculated: Tc = (T up + T down) ÷ 2 Example: T up = 57 lbs, T down = 75 lbs Tc = (T up + T down) ÷ 2 Tc = 57 + 75 = 132 = 66 lbs = acceptable 2 2 (19) Verify that the safety clips (5) are installed on both turnbuckles (1 and 2) (Ref. Figure 206). (20) Remove the aft elevator bellcrank rig pin (3) (Ref. Figure 208). (21) With assistance, manually push up on the elevator surface until the elevator control horns contact the up stop bolts in the tail. Use a go/no-go scale (8, Table 1, 27-00-00) to verify that the forward elevator bellcrank up stop bolt (5) has a 0.37 ± 0.06 inch clearance from the stop bracket (4) located in the structure (Ref. Figure 201). CAUTION: Carefully lower the elevator surface. Do not allow the elevator to free fall to the down position. This could cause damage to the elevator system. (22) Apply clay to the forward elevator down stop bolt (6) and slowly lower the elevator surface until the elevator control horns rest on the down stop bolts in the tail. Verify that the forward elevator bellcrank down stop bolt (6) has a 0.31 +0.12/ -0 inch clearance from the stop bolt (1) located in the structure. (23) If adjustment is needed, loosen the jam nut(s) (7) on the stop bolt(s) (5 and/or 6) and adjust the bolt(s). Check the clearance. When adjustment is completed, tighten the stop bolt jam nut(s) (7). WARNING: After May 2014, Ensure Mandatory Service Bulletin 27-4119 has been accomplished and Kit 114-5060 has been installed. CAUTION: If the bobweight stop bolt(s) (6) align with a depression in the bobweight (8), measure the clearance from the deepest impression (in-line with the stop bolt) to the face of the stop bolt (6). If the impression (depression) in the bobweight is more than 0.080 inch deep, contact Beechcraft Technical Support.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (24) (Without Kit 114-5060 installed) Measure the bobweight stop bolt (6) clearance with the elevator control horns resting on the down stops in the tail. The bobweight stop bolt (6) must have 0.56 +0/ -0.12 inch clearance with the bobweight (8) (Ref. Figure 203, Sheet 1). (25) (With Kit 114-5060 installed) Measure the bobweight stop bolt(s) (6) clearance with the elevator control horns resting on the down stops in the tail. The bobweight stop bolt(s) (6) must have 0.56 +0/ -0.12 inch clearance with the bobweight (8) (Ref. Figure 203, Sheet 2). (a) Measure the clearance between each stop bolt (6) and the bob weight (8). The clearances must be within 0.032 inch of each other. (26) If adjustment is needed, loosen the jam nut(s) (5) on the stop bolt(s) (6) and adjust the bolt(s) (6). Check the clearance. When adjustment is completed, tighten the jam nut(s) (5). (27) Remove nuts, washers and bolts (8) attaching the push-pull rods (10) to the elevator horns (4) and discard the nuts (Ref. Figure 204). (28) With the aft elevator bellcrank push-pull rod disconnected from the elevator horn, the forward bellcrank down stop bolt (6) must make contact with the stop bolt (1) in the structure (Ref. Figure 201). NOTE: The forward bellcrank down stop bolt (6) (Figure 201) must make contact before the bobweight makes contact with the stop bolt(s) (6) Without Kit 114-5060 (Ref. Figure 203, Sheet 1), With Kit 114-5060 (Ref. Figure 203, Sheet 2). (29) Verify that the bobweight stop bolt(s) (6) have clearance. If the bobweight stop bolt(s) have clearance proceed to Step (31). If the bobweight stop bolt(s) (6) make contact first, connect the elevator push-pull rods to the elevator horns and then adjust the bobweight stop bolt(s) (6) to the maximum clearance per Step (24) Without Kit 114-5060 (Ref. Figure 203, Sheet 1) or Step (25) With Kit 114-5060 (Ref. Figure 203, Sheet 2) and adjust the forward elevator bellcrank down stop bolt (6) to its minimum clearance per Step (22) (Ref. Figure 201). (30) If more adjustment is needed, perform the following Steps to adjust the push-pull rod assembly between the control column and the forward elevator bellcrank to achieve the required clearance. If no adjustment is needed, proceed to Step (31). (a) Measure the center to center length of the push-pull rod assembly (8). The length must be 15.17 +0.19/ -0.06 inches (Ref. Figure 201). (b) If adjustment is needed, remove nut, washer, and bolt (6) from the control column (4) and discard the nut. Loosen jam nut (2) at the base of the rod end (3). Rotate the rod end (3) until measurement is met (Ref. Figure 205). (c) After adjustment, tighten jam nut (2). The rod end (3) should have zero to seven threads maximum exposed beyond the jam nut when the adjustment is complete. Verify threads are visible through the tube inspection hole (5). (d) Install push-pull rod assembly (1) to the control column (4) with a bolt, washer and new self locking nut (6). (31) Install bolts, washers and new self locking nuts (8) attaching the push-pull rods (10) to the elevator horns (4) (Ref. Figure 204).

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(32) Gust Lock Pin Check - Install the gust lock (control lock) pin into the control column (Ref. Chapter 27-70-00). Verify the elevators are 7° to 15° down (the elevators do not need to be on the down stops with the gust lock pin installed). However, if the elevators are full down and resting on the elevator control horn stop bolts (5), the force required to push on the pilot’s control wheel while inserting the gust lock pin must be a maximum of ten pounds. If the force is too high, adjust (shorten) the push-pull tube per Step (30), then return to Step (24) or Step (25) to check the bobweight clearance. Remove the gust lock pin. (33) Connect the autopilot elevator servo cables to the elevator control cables, if installed (Ref. Figure 211). (a) Connect the elevator servo upper cable (2) to the bridle clamp (5). (b) Torque the elevator servo upper cable bridle clamp bolts to 55 ± 5 inch-pounds. Check for a minimum gap of 0.005 inch between the clamp halves. (c) Connect the elevator servo lower cable (3) using turnbuckle (7). (d) Perform the ELEVATOR SERVO CABLE TENSIONING procedure (Ref. Chapter 22-10-00). WARNING: Verify elevator movement in the following Step by moving only the cockpit control column. (34) Final Travel Check - With assistance, move a cockpit control wheel (yoke) aft and verify the elevator surfaces move up 20° +1°/ -0° and that the elevator stop bolts contact the elevator control horns. Move the control wheel forward and verify the elevator surfaces move down 14° +1°/ -0° and that the elevator stop bolts contact the elevator control horns. If these requirements are not met, repeat this rigging procedure in its entirety. (35) If travel board P/N D807 101-610000-1/-2 at HSS 35.00 was used, connect the elevator trim tab pushrods (4) to the elevator trim tab control horns (2) and tighten clevis jam nut (5) (Ref. Figure 209). (36) Remove the travel board(s) and tape from the horizontal stabilizer(s). (37) Perform the ELEVATOR CONTROL SYSTEM FRICTION TEST procedure in this section.

C. Follow on Maintenance (1) Install left lower tailcone panel 7 (Ref. Chapter 06-50-00, FUSELAGE ACCESS PANELS). (2) Install aft fairing 15 (Ref. Chapter 06-50-00, STABILIZER ACCESS PANELS). (3) Install the aft horizontal stabilizer access panel 18 (Ref. Chapter 06-50-00, STABILIZER ACCESS PANELS). (4) Install the lower aft rig pin hole access panels 34 on the left and right side of the upper vertical stabilizer (Ref. Chapter 06-50-00, STABILIZER ACCESS PANELS). (5) Install the bottom forward fuselage access panel 3, if removed (Ref. Chapter 06-50-00, FUSELAGE ACCESS PANELS). (6) Install the right forward pedestal side panel.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (7) Install the left flight compartment floorboard panels 2, 21, and 22 (Ref. Chapter 06-50-00, FLOOR ACCESS PANELS). (8) Install the left flight compartment inboard seat track. (9) Install the left flight compartment carpet. (10) Install the left flight compartment seat (Ref. Chapter 25-10-00, SEAT INSTALLATION).

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15.17 +0.19/ -0.06 INCHES

2

8 1

0.37 ± 0.06 INCH

0.31 +0.12/ -0 INCH

4

3

6

7

5

7

FWD

1. STOP BOLT IN STRUCTURE 2. ELEVATOR FORWARD BELLCRANK 3. RIG PIN HOLE 4. STOP BRACKET 5. BELLCRANK UP STOP BOLT 6. BELLCRANK DOWN STOP BOLT 7. JAM NUT 8. PUSH-PULL ROD ASSEMBLY

NOTE: THIS INFORMATION IS FOR THE MODEL 1900/1900C ONLY. DO NOT USE FOR OTHER MODEL 1900 SERIES AIRCRAFT. UC27B 034847AC.AI

Figure 201 Elevator Forward Bellcrank Stops

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POUNDS OF TENSION

3/16" DIAMETER ELEVATOR CABLE TENSION GRAPH

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Figure 202 Elevator Cable Tension Graph

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1. FORWARD BELLCRANK 2. RIG PIN HOLE 3. ELEVATOR DOWN CABLE 4. ELEVATOR UP CABLE 5. JAM NUT 6. STOP BOLT 7. LINK ASSEMBLY 8. BOBWEIGHT

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FS 108

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0.56 +0/ -0.12 INCH 5 PRESSURE BULKHEAD

CONTROL COLUMN & BOBWEIGHT FWD

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DETAIL

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A UC27B 130146AA.AI

Figure 203 (Sheet 1 of 2) Control Column Bobweight Adjustments, Without Kit 114-5060

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1. FORWARD BELLCRANK 2. RIG PIN HOLE 3. ELEVATOR DOWN CABLE 4. ELEVATOR UP CABLE 5. JAM NUT 6. STOP BOLT 7. LINK ASSEMBLY 8. BOBWEIGHT

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4 7

FS 108

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0.56 +0/ -0.12 INCH 5 PRESSURE BULKHEAD

CONTROL COLUMN & BOBWEIGHT FWD

DETAIL

6

DETAIL

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A UC27B 034843AD.AI

Figure 203 (Sheet 2 of 2) Control Column Bobweight Adjustments, With Kit 114-5060

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1. LEFT ELEVATOR CONTROL HORN SUPPORT 2. JAM NUT 3. UP STOP BOLT CONTACT AREA 4. LEFT ELEVATOR CONTROL HORN 5. ELEVATOR CONTROL HORN STOP BOLTS 6. JAM NUT 7. SAFETY WIRE 8. BOLT 9. ROD END 10. ELEVATOR AFT BELLCRANK PUSH-PULL ROD 11. DOWN STOP BOLT CONTACT AREA 12. INSPECTION HOLE 2

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FWD VIEW LOOKING UP AND INBOARD AT LEFT HAND SIDE (LEFT SIDE SHOWN, RIGHT SIDE OPPOSITE) DETAIL

A UC27B 034932AB.AI

Figure 204 Elevator Horn Stop Bolts and Push-Pull Rods

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1. ELEVATOR FORWARD BELLCRANK PUSH-PULL ROD 2. JAM NUT 3. ROD END 4. CONTROL COLUMN 5. INSPECTION HOLE 6. BOLT, WASHER, NUT

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VIEW LOOKING UP DETAIL

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Figure 205 Elevator Forward Bellcrank Push-Pull Rod Adjustment

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1. ELEVATOR DOWN TURNBUCKLE 2. ELEVATOR UP TURNBUCKLE 3. ELEVATOR UP CABLE 4. ELEVATOR DOWN CABLE 5. SAFETY CLIPS

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2

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VIEW LOOKING INBOARD FROM LEFT HAND SIDE DETAIL

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Figure 206 Elevator Cable Turnbuckle Adjustment

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

1. FORWARD ELEVATOR BELLCRANK RIG PIN HOLE 2. TRIM WHEEL 3. PEDESTAL 4. FORWARD ELEVATOR BELLCRANK (UNDER STRUCTURE)

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VIEW LOOKING DOWN LEFT SIDE OF PEDESTAL DETAIL

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UC27B 034841AA.AI

Figure 207 Elevator Forward Bellcrank Rig Pin Installation

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1. ELEVATOR 2. RUDDER 3. ELEVATOR AFT BELLCRANK RIG PIN 4. VERTICAL STABILIZER 5. HORIZONTAL STABILIZER

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4

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VIEW LOOKING UP LEFT HAND SIDE DETAIL

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Figure 208 (Sheet 1 of 2) Elevator Aft Bellcrank Rig Pin Installation

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

6. ELEVATOR AFT BELLCRANK 7. ELEVATOR AFT RIG PIN HOLE

6

FWD 7

VERTICAL STABILIZER (REF)

DETAIL

B UC27B 040558AB.AI

Figure 208 (Sheet 2 of 2) Elevator Aft Bellcrank Rig Pin Installation

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1. ELEVATOR TRIM TAB 2. ELEVATOR TRIM TAB CONTROL HORN 3. PUSHROD END 4. PUSHRODS 5. CLEVIS JAM NUT

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4 ELEVATOR ASSEMBLY (REF)

HORIZONTAL STABILIZER (REF)

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Figure 209 Elevator Trim Tab Connection

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

D. Operational Check CAUTION: With the control column pulled to the aft position, allowing the control column to free fall to the forward position can cause damage to the elevator system. (1) Pull the pilot’s control wheel aft and make sure that the elevator travels to the full up position with no unusual noise or binding. (2) Move the pilot’s control wheel forward and make sure that the elevator travels to the full down position with no unusual noise or binding. (3) Repeat Steps (1) and (2) with the copilot’s control wheel. (4) If requirements are not met, perform the ELEVATOR CONTROL SYSTEM RIGGING procedure in this section.

E. Cable Tension Check (1) Remove the elevator aft bellcrank rig pin hole access panels 34 on the left and right side of the upper vertical stabilizer (Ref. Chapter 06-50-00, STABILIZER ACCESS PANELS). (2) Remove the left aft fuselage panel 7 (Ref. Chapter 06-50-00, FUSELAGE ACCESS PANELS). (3) Move either control wheel to the full forward and full aft position three to four times to equalize system tension. (4) Disconnect the autopilot elevator servo cables from the primary control cables, if installed (Ref. Figure 211). (a) Disconnect the elevator servo lower cable (3) by removing the safety clips (6) and completely loosen turnbuckle (7). (b) Disconnect the elevator servo upper cable (2) from the bridle clamp (5). (c) Position the upper and lower elevator servo cable to prevent interference with the movement of the elevator control cables. (5) Move the elevator surface (1) and install a rig pin (3) (7, Table 1, 27-00-00) in the aft elevator bellcrank through the vertical stabilizer (4). Using minimum force try to manually move the elevators up and down to verify that the rig pin is correctly installed. Ensure that the elevators do not move. Look at the opposite side of the vertical stabilizer to verify that the rig pin has protruded (Ref. Figure 208). (6) Check the cable tension by performing the following Steps: NOTE: The forward bellcrank rig pin must NOT be installed when measuring the cable tensions to calculate the combined (average) tension (Tc). Do not measure the tension when the temperature is varying rapidly or with the airplane located in direct sunlight. (a) Allow the temperature of the airframe to stabilize before measuring and adjusting the cable tension.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (b) Through the aft fuselage access panel 7, measure the temperature in the compartment next to the elevator cables (3 and 4) near the turnbuckles (1 and 2) (Ref. Figure 206). (c) Refer to the Elevator Cable Tension Graph Figure 202 and read the pounds of tension for the measured temperature. (d) Position a cable tensiometer (4, Table 1, 27-00-00) at least three inches from the turnbuckles (1 and 2) and measure the cable tension of both cables (3 and 4) (Ref. Figure 206). Cable diameter is noted in Figure 202. (e) Record the values as T up (aft cable) (3) and T down (fwd cable) (4). Calculate the combined (average) tension (Tc) by adding the T up and T down tensions together and divide by two. Example: Tc = (T up + T down) ÷ 2. NOTE: The combined tension (Tc) must be within a ± 8 pounds of the tension found in Figure 202. (7) If tension does not meet requirements, perform the ELEVATOR CONTROL SYSTEM RIGGING procedure in this section. (8) Connect the autopilot elevator servo cables to the elevator control cables, if installed (Ref. Figure 211). (a) Connect the elevator servo upper cable (2) to the bridle clamp (5). (b) Torque the elevator servo upper cable bridle clamp bolts to 55 ± 5 inch-pounds. Check for a minimum gap of 0.005 inch between the clamp halves. (c) Connect the elevator servo lower cable (3) using turnbuckle (7). (d) Perform the ELEVATOR SERVO CABLE TENSIONING procedure (Ref. Chapter 22-10-00). (9) Remove the aft elevator bellcrank rig pin (3) (Ref. Figure 208). (10) Install the elevator aft bellcrank rig pin hole access panels 34 on the left and right side of the upper vertical stabilizer (Ref. Chapter 06-50-00, STABILIZER ACCESS PANELS). (11) Install the left aft fuselage panel 7 (Ref. Chapter 06-50-00, FUSELAGE ACCESS PANELS). (12) Perform the ELEVATOR OPERATIONAL CHECK procedure in this section.

F. Functional Check (1) Install an elevator travel board (Ref. 27-00-02, ELEVATOR TRAVEL BOARD INSTALLATION AT HSS 50.00). WARNING: Verify elevator movement in the following Steps by moving only the cockpit control column. NOTE: One travel board may be used and moved from one side to the other. CAUTION: With the control column pulled to the aft position, allowing the control column to free fall to the forward position can cause damage to the elevator system.

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(2) If using travel board P/N D807 101-610000-1/-2 at HSS 35.00, perform the following: (a) Loosen clevis jam nut (5) and disconnect pushrod (4) from the elevator trim tab control horn (2). Do not change the adjustment of the pushrod end (3) (Ref. Figure 209). (b) Align the trailing edge of the elevator trim tab (1) to the trailing edge of the elevator surface. (3) Pull a cockpit control wheel (yoke) aft and verify the elevator surfaces move up 20° +1°/ -0° (Ref. 27-00-02, READING A TRAVEL BOARD) and that the elevator stop bolts contact the elevator control horns. Make sure the elevator system moves smoothly without any unusual noise or binding. (4) Move the control wheel (yoke) forward and verify the elevator surfaces move down 14° +1°/ -0° and that the elevator stop bolts contact the elevator control horns. Make sure the elevator system moves smoothly without any unusual noise or binding. (5) Repeat Steps (1) thru (4) for the opposite side. (6) If the elevator surfaces do not achieve desired travel, perform the ELEVATOR CONTROL SYSTEM RIGGING procedure in this section. (7) If travel board P/N D807 101-610000-1/-2 at HSS 35.00 was used, connect the elevator trim tab pushrod (4) to the elevator trim tab control horn (2) and tighten clevis jam nut (5) (Ref. Figure 209). (8) Remove the travel board(s).

G. Friction Check NOTE: Take all force readings with the elevator control system completely installed: downsprings attached, cables rigged with tension applied, bobweight mounted on control yoke and elevator servo cables attached. (1) Remove aft fuselage panel 7 (Ref. Chapter 06-50-00, FUSELAGE ACCESS PANELS). (2) To obtain the maximum allowable system friction Fs (max), measure the tension of both elevator cables in the aft fuselage. Take the higher of the two readings and multiply it by a factor of 0.106. For example: Fs (max) = 0.106 X Highest Cable Tension Fs (max) = 0.106 X 98 lbs Fs (max) = 10.4 lbs (3) Attach tie straps around the pilot’s or copilot’s control wheel and attach a push-pull gage so that the gage is approximately center of the control wheel. Alternative method is to use a push-pull gage attached and centered between the inboard grips of the pilot and copilots control wheels. NOTE: Do not use tape on tube due to close fit between the control wheel tube and the subpanel support bushing.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (4) Determine method to identify when the control wheel passes through neutral. Suggested method is to move the elevators to neutral, install rig pin and use a non-permanent marker to mark the control wheel tube (3) (Ref. Figure 212). Remove the rig pin. (5) Smoothly pull the control wheel aft from full forward (elevator down) position until the elevator passes through the neutral position and record the reading from the push-pull gage. From the full aft (elevator up) position, slowly let the control wheel move forward until the elevator passes through neutral and record the reading from the push-pull gage. Repeat four more times for a total of 5 cycles, recording each cycle. (6) Average the 5 readings from pulling the control wheel aft from full down (elevator down) and record this value as F up. (7) Average the 5 readings from moving the control wheel forward through neutral and record this value as F down. (8) With the maximum allowable system friction established, use the force measurements to calculate the actual system friction. System friction value is obtained by taking the difference of the F up and F down force values divided by 2, or: Fs = (F up - F down) ÷ 2 Example: F up = 46 lbs, F down = 28 lbs Fs = (F up - F down) ÷ 2 Fs = 46 - 28 = 18 = 9 = acceptable 2 2 (9) Combined downspring and bobweight force range (Fc) is 37.5 ± 1.5 pounds. (10) Combined downspring and bobweight force is obtained by adding the F up and F down values together and dividing by 2, or: Fc = (F up + F down) ÷ 2 Example: F up = 46 lbs, F down = 28 lbs Fc = (F up + F down) ÷ 2 Fc = 46 + 28 = 74 = 37 = acceptable 2 2 (11) If adjustment is not required, proceed to Step (12). If adjustment is required, perform the following Steps: (a) Remove the top horizontal access panel 17 (Ref. Chapter 06-50-00, STABILIZER ACCESS PANELS). NOTE: The downsprings can be installed in different adjustment holes. Downsprings do not need to be adjusted equally.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (b) If Fc is above the range of 37.5 ± 1.5 lbs, decrease the downspring (1) tension (adjust the spring shorter) by moving the spring end to another adjustment hole (2) in the adjustment link (4) (Ref. Figure 210). (c) If Fc is below the proper range, increase the downspring (1) tension (adjust the spring longer) by moving the spring end to another adjustment hole (2) in the adjustment link (4). (d) Repeat Steps (3) through (11). (e) Install the top horizontal access panel 17 (Ref. Chapter 06-50-00, STABILIZER ACCESS PANELS). (12) If system friction-force cannot be brought under the limits: (a) Check cable installation, pulley bearings, push-pull rod ends, bellcrank bearings, the control yoke and all linkages associated with the elevator control system. (b) Perform the ELEVATOR INSPECTION procedures (Ref. 27-30-00). (13) Check the outboard leading edge of the elevators for elevator counterweight screw and elevator balance weight bolt interference. (14) If a mark for neutral was made on the control wheel tube (3), clean the mark off the tube (Ref. Figure 212). (15) Install aft fuselage panel 7 (Ref. Chapter 06-50-00, FUSELAGE ACCESS PANELS).

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1. DOWNSPRING 2. ADJUSTMENT HOLES 3. AFT ELEVATOR BELLCRANK 4. ADJUSTMENT LINK

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FWD

VERTICAL STABILIZER (REF)

DETAIL

A

Figure 210 Elevator Aft Bellcrank Downspring Adjustment

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

1

A

1. ELEVATOR AFT BELLCRANK 2. ELEVATOR SERVO UPPER CABLE 3. ELEVATOR SERVO LOWER CABLE 4. AUTOPILOT ELEVATOR SERVO ASSEMBLY 5. BRIDLE CLAMP 6. SAFETY CLIPS 7. TURNBUCKLE 8. LEFT-HAND THREAD GROOVE

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DETAIL

DETAIL

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UC27B 061564AA.AI

Figure 211 Elevator Servo Installation

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1. CONTROL WHEEL 2. SUPPORT 3. CONTROL WHEEL TUBE

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3 UC27E 062173AA.AI

Figure 212 Control Wheel Installation

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2. BOBWEIGHT AND STOP A. Inspection CAUTION: Carefully lower the elevator surface. Do not allow the elevator to free fall to the down position. This could cause damage to the elevator system. (1) With the control column in its furthest forward, (control column in the relaxed, elevator down) position, inspect the alignment of the bobweight (5) with the stop bolt(s) (4) Without Kit 114-5060 (Ref. Figure 214, Sheet 1), With Kit 114-5060 (Ref. Figure 214, Sheet 2). (a) Check alignment of the bobweight (5) with the stop bolt(s) (4). With side pressure applied by hand ensure no part of the stop bolt(s) (4) protrudes beyond the face of the bobweight (5) on either edge. (b) Inspect for evidence of scraping along either side of the bobweight (5) by the stop bolt(s) (4). (c) Inspect the condition of the stop bolt(s) (4) and stop bracket (6). Inspect for evidence of damage or deformation by contact with the bobweight bellcrank assembly (3 and 5). (2) Correct any discrepancies revealed in Step (1) above. If the stop bolt(s) (4) or stop bracket (6) are bent, replace the parts. Contact BC Technical Support for assistance. (3) Inspect the link assembly (4) for proper orientation (Ref. Figure 213). NOTE: Correct orientation is when the link assembly attach point (3) bolt is positioned aft of the control column. Attach point (3) bolt center, must remain above a line drawn between attach point (1) and (2) bolt centers as the bobweight moves to its maximum aft position. (4) If link assembly is properly oriented proceed to Step (5). If not perform the BOBWEIGHT LINK ASSEMBLY inspection in this section. (5) With the control column resting in the forward (elevator down) position so the elevator control horns are on the primary stops, inspect the bobweight clearance with the stop bolt(s) (6). The bobweight stop bolt(s) (6) must have 0.56 +0.00/ -0.12 inch clearance with the face of the bobweight (8). If not, adjust the stop bolt(s) (6) Without Kit 114-5060 (Ref. Figure 203, Sheet 1), With Kit 114-5060 (Ref. Figure 203, Sheet 2). (6) Perform the ELEVATOR CONTROL SYSTEM Operational Check procedure in this section. (a) If the stop bolt bracket or bobweight is replaced, perform the ELEVATOR CONTROL SYSTEM Functional Check procedure in this section.

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3. BOBWEIGHT LINK ASSEMBLY A. Inspection (1) Remove access panel (3), located aft of the nose wheel well. Refer to the Chapter 06-50-00, FUSELAGE ACCESS PANELS illustration in the Airplane Access Panels - Description and Operation section. WARNING: Whenever any part of this system is dismantled, adjusted, repaired or replaced, a detailed investigation must be made on completion to make sure that distortion, tools, rags, or any other loose articles, or foreign matter that could impede the free movement and safe operation of the system are not present, and the systems and installation in the work area are clean. (2) Inspect link assembly (4) at attachment to the control column for modification in accordance with Mandatory Service Bulletin 27-3739. There must be a large diameter washer (7) installed on both sides of the link assembly bearing. If not, comply with Service Bulletin 27-3739 modification instructions (Ref. Figure 213). (3) Remove both bobweight stop bolt(s) (4) from the stop bracket (6) Without Kit 114-5060 (Ref. Figure 214, Sheet 1), With Kit 114-5060 (Ref. Figure 214, Sheet 2). CAUTION: Carefully lower the elevator surface. Do not allow the elevator to free fall to the down position. This could cause damage to the elevator system. (4) With an assistant in the cockpit manipulating the control column, remove cotter pin (if installed), nut, washer, and bolt (6) from the control column (4) and discard the nut from the push-pull rod assembly between the control column and the forward elevator bellcrank (Ref. Figure 205). (5) Slowly allow the control column to move to its furthest forward (elevator down) position. (6) Inspect the link assembly (4) for proper orientation (Ref. Figure 213). NOTE: Correct orientation is when the link assembly attach point (3) bolt is positioned aft of the control column. Attach point (3) bolt center, must remain above a line drawn between attach point (1) and (2) bolt centers as the bobweight moves to its maximum aft position. (7) If link assembly (4) is properly oriented proceed to Step (8). If not inspect the following and contact BC Technical Support for assistance. (a) Remove the link assembly (4). (b) Inspect the Control column at the link assembly attach point (1) and the bobweight bellcrank at the link assembly attach holes (3). These attach holes are 0.250 to 0.254 inch diameter when new. (c) Inspect the link assembly. If the link assembly bearing is loose, see the Model 1900 Structural Repair Manual (Ref. Chapter 27). (d) Inspect the bobweight support structure at the bobweight bellcrank attach point (2). The holes are 0.250 to 0.254 inch diameter when new. (e) Correct any discrepancies revealed in Steps (a), (b), (c) and (d) above.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (f) Install the link assembly (4). (g) Repeat Step (6). (8) Inspect the bobweight support structure (2) at the aft side of bobweight bellcrank (3) for contact (Ref. Figure 214). (a) Evidence of damage or deformation by contact with the bobweight bellcrank assembly will be found on the aft side of the bob weight support structure in the lower area as noted by “Check This Area” arrow, Without Kit 114-5060 (Ref. Figure 214, Sheet 1), With Kit 114-5060 (Ref. Figure 214, Sheet 2). NOTE: The aft side of the bob weight support structure has an approximately 1.8 inch wide flat surface. Contact with the bobweight bellcrank assembly may damage or deform this aft surface. The outboard side of the bob weight support structure may have upholstery (13) material. This upholstery (13) may be removed as needed so the sheet metal is exposed to aid in this inspection (Ref. Figure 213). (b) If the structure is distorted more than 1.00 inch aft from the center of the bobweight bellcrank pivot (1) bolt hole contact BC Technical Support for repair Without Kit 114-5060 (Ref. Figure 214, Sheet 1), With Kit 114-5060 (Ref. Figure 214, Sheet 2). (9) With an assistant in the cockpit manipulating the control column, Install the push-pull rod assembly (1) to the control column (4) with bolt, washer, new nut and cotter pin (6) (if installed) (Ref. Figure 205). (10) Install the bobweight stop bolt(s) (6) and check clearance with the elevator control horns resting on the down stops in the tail. Adjust the bobweight stop bolt(s) (6) to have 0.56 +0.00/ -0.12 inch clearance with the face of the bobweight (8) Without Kit 114-5060 (Ref. Figure 203, Sheet 1), With Kit 114-5060 (Ref. Figure 203, Sheet 2). (11) Install access panel 3, located aft of the nose wheel well. Refer to the Chapter 06-50-00, FUSELAGE ACCESS PANELS illustration in the Airplane Access Panels - Description and Operation section. (12) Perform the ELEVATOR CONTROL SYSTEM Functional Check in this section.

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ATTACH POINT ATTACH POINT ATTACH POINT LINK ASSEMBLY SUPPORT STRUCTURE BOBWEIGHT BELLCRANK ASSEMBLY WASHERS WASHERS NUT WASHER BOLT CONTROL COLUMN UPHOLSTERY

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Figure 213 Link Assembly, Proper Orientation

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

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1.00 MAXIMUM

6 CHECK THIS AREA

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BOBWEIGHT BELLCRANK PIVOT BOBWEIGHT SUPORT STRUCTURE BOBWEIGHT BELLCRANK STOP BOLT BOBWEIGHT STOP BRACKET

DETAIL

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Figure 214 (Sheet 1 of 2) Bobweight Support Structure, Without Kit 114-5060

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6 CHECK THIS AREA

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BOBWEIGHT BELLCRANK PIVOT BOBWEIGHT SUPORT STRUCTURE BOBWEIGHT BELLCRANK STOP BOLT BOBWEIGHT STOP BRACKET

DETAIL

A UC27B 120167AB.AI

Figure 214 (Sheet 2 of 2) Bobweight Support Structure, With Kit 114-5060

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FLIGHT CONTROLS ELEVATOR TRIM TABS MAINTENANCE PRACTICES

27-30-03 200200

1. PROCEDURES A. Removal (1) Perform the ELEVATOR REMOVAL procedure (Ref. 27-30-00). (2) Disconnect the bonding jumper (2) from the inboard end of the tab (6). Remove the screw, nut, and washers (1). Discard the nut. (Ref. Figure 201). (3) Remove the hinge pin retention bolt (5) which secures the inboard end of the hinge pin (4) to the tab (6). NOTE: Lubricating the hinge and hinge pin with lubricant (106, Table 1, Chapter 91-00-00) will facilitate hinge pin removal. (4) While supporting the tab (6) remove the hinge pin (4) from the hinge. Remove the tab (6) from the elevator (3).

B. Installation NOTE: Repair, modification, painting or replacement of the elevator or the elevator tab requires balancing (Ref. Chapter 55-20-00). (1) Lubricate the elevator tab hinge and the hinge pin (4) with lubricant (106, Table 1, Chapter 91-00-00) (Ref. Figure 201). (2) Position the elevator tab (6) on the elevator and install the hinge pin (4). (3) Secure the inboard end of the hinge pin (4) to the tab with the retention bolt (5). (4) Connect the bonding jumper (2) to the inboard end of the tab (6) with the screw, washers and new nut (1). (5) Perform the ELEVATOR INSTALLATION procedure (Ref. 27-30-00). (6) Perform the ELECTRICAL BONDING CHECK procedure (Ref. Chapter 20-00-01). (7) With full elevator down and elevator trim tab at trailing edge full down, check that there is clearance between the elevator trim tabs and the top of the rudder at full left and full right rudder. (8) Perform the ELEVATOR TRIM TAB FUNCTIONAL CHECK procedure (Ref. 27-30-05). If the trim tab was replaced perform the ELEVATOR TRIM TAB RIGGING procedure (Ref. 27-30-05). (9) Perform the ELEVATOR TRIM TAB FREEPLAY CHECK procedure in this section.

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A 1. SCREW, NUT, AND WASHERS 2. BONDING JUMPER 3. ELEVATOR 4. HINGE PIN 5. HINGE PIN RETENTION BOLT 6. TAB

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UC27B 042321AA.AI

Figure 201 Elevator Trim Tab Installation

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

C. Freeplay Checks NOTE: Movement or jarring of the airplane will invalidate the elevator trim tab freeplay readings. The airplane should be placed in a hangar and no personnel in or on the airplane during the freeplay check. (1) Obtain a copy of Table 201. (2) Visually inspect the elevator trim tabs for any damage, security of hinge attach points and for tightness of the actuating systems. Inconsistencies should be remedied prior to checking the freeplay of the tabs. (3) Remove left and right outboard elevator access plates 22 (Ref. Chapter 6-50-00, STABILIZER ACCESS PANELS). (4) Lubricate both elevator trim tab actuators with grease (1, Table 2, 27-00-00). (5) Remove access panel 34 on the left and right side of the vertical stabilizer (Ref. Chapter 6-50-00, STABILIZER ACCESS PANELS). (6) Move the elevator surface (1) and install a rig pin (3) (7, Table 1, 27-00-00) in the aft elevator bellcrank through the vertical stabilizer (4). Using minimum force try to manually move the elevators up and down to verify that the rig pin is correctly installed. Ensure that the elevators do not move. Look at the opposite side of the vertical stabilizer to verify that the rig pin has protruded (Ref. Figure 202). (7) Align the elevator trim tab trailing edge with elevator trailing edge. (8) Apply tape (for paint protection) on the top surface 3.50 inches aft of the tab hinge line along the centerline of the tab actuator. Apply tape in the corresponding position on the bottom surface of the tab. This will be the point of pressure against the tab by the push-pull scale (6, Table 1, 27-00-00). WARNING: Ensure the trim tab freeplay check fixture is securely attached to the elevator before releasing supporting pressure to prevent damage to the equipment or injury to personnel. (9) Install the trim tab freeplay check fixture (1) (1, Table 1, 27-00-00) onto the elevator (4) with the dial indicator stem (5) positioned on the top surface of the elevator tab (3) 3.00 inches aft of the tab hinge line on the outboard edge of the elevator tab (3) (Ref. Figure 203). (10) Position the dial indicator (2) so the stem (5) is depressed 0.10 inch when in contact with the tab (3) surface initially. Turn the rotating face of the dial indicator (2) to zero. Do not reset the dial indicator during this procedure. (11) With a push-pull scale (6, Table 1, 27-00-00) perform Steps (a) thru (d) and record the dial readings on a copy of Table 201. (a) Apply three pounds of downward load on the top surface of the tab. Record the dial reading as A. (b) Release half the load until a 1.5 pound downward load is obtained. Record the dial reading as B.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (c) Apply three pounds upward load on the bottom surface of the tab. Record the dial reading as C. (d) Release half the load until a 1.5 pound upward load is obtained. Record the dial reading as D. (12) Perform the calculations to the data on Table 201 as follows: (a) Record A, B, C and D as positive numbers. (b) Multiply B by 2 and record as 2B. (c) Subtract A from 2B and record as X. (d) Multiply D by 2 and record as 2D. (e) Subtract C from 2D and record as Y. NOTE: The results of X and Y can be negative numbers. (f) Add X and Y and record as E. (13) If deflection of the tab is within allowable limits, the tab and its linkage are in good condition. (14) If the freeplay is excessive, disconnect the trim tab actuator rod and visually inspect the bolts and bushing for indications of excessive wear. Replace excessively worn parts. (15) If all associated linkage is in good condition (no excessive wear) the actuator needs to be checked for excessive play and/or replaced. (16) Remove the trim tab freeplay check fixture (1) from the elevator. Remove the tape from the elevator (Ref. Figure 203). (17) Repeat Steps (2), (7) through (16) on the opposite elevator trim tab. (18) Install left and right outboard elevator access plates 22 (Ref. Chapter 6-50-00, STABILIZER ACCESS PANELS). CAUTION: Carefully lower the elevator surface. Do not allow the elevator to free fall to the down position. This could cause damage to the elevator system. (19) Remove rig pin (3) from the vertical stabilizer (4) (Ref. Figure 202). (20) Install access panel 34 on the left and right side of the vertical stabilizer (Ref. Chapter 6-50-00, STABILIZER ACCESS PANELS).

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Table 201 Elevator Trim Tab Freeplay Limits Serial Number: _______________

Date: _______________

Left Elevator Trim Tab (__________) X 2

= (__________)

- (__________)

= (__________)

2B

A

X

= (__________)

- (__________)

= (__________)

D

2D

C

Y

(__________)

+ (__________)

= (__________)

X

Y

E

B (__________) X 2

(E = 0.006 inch maximum) Right Elevator Trim Tab (__________) X 2

= (__________)

- (__________)

= (__________)

2B

A

X

= (__________)

- (__________)

= (__________)

D

2D

C

Y

(__________)

+ (__________)

= (__________)

X

Y

E

B (__________) X 2

(E = 0.006-inch maximum)

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1. ELEVATOR 2. RUDDER 3. AFT ELEVATOR BELLCRANK RIG PIN 4. VERTICAL STABILIZER 5. HORIZONTAL STABILIZER 6. TRIM TAB

A 5

6 1

2 4

3

VIEW LOOKING UP LEFT HAND SIDE DETAIL

A

Figure 202 Elevator Aft Bellcrank Rig Pin Installation

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UC27B 041403AA.AI

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

A 1. TRIM TAB FREEPLAY CHECK FIXTURE 2. DIAL INDICATOR 3. ELEVATOR TAB 4. ELEVATOR 5. DIAL INDICATOR STEM

4 2 1 3

5

DETAIL

A

UC27B 043403AB.AI

Figure 203 Trim Tab Freeplay Fixture Installation

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2. ELEVATOR ELECTRIC TRIM A. Operational Check (1) Apply electrical power to the airplane. (2) Set the elevator trim switch, located on the pedestal, to the ON position. (3) Actuate each switch of the dual trim switch, on the pilot’s and copilot’s control wheel, independently to the NOSE DOWN and the NOSE UP position and verify that the trim system does not activate. CAUTION: While performing this procedure, do not keep the trim buttons (switches) depressed after the trim tab has reached its full limit of travel. (4) Actuate both trim switches on the pilot’s control wheel to the NOSE UP position and note the trim wheel movement in the proper direction as well as full travel. Verify visually that the trim tab itself travels to the proper position (trim tab full down). (5) Actuate both trim switches on the pilot’s control wheel to the NOSE DOWN position and note the trim wheel movement in the proper direction as well as full travel. Verify visually that the trim tab itself travels to the proper position (trim tab full up). (6) Repeat Steps (4) and (5) on the copilot’s control wheel. NOTE: Review Steps (7) and (8) before proceeding with this procedure. Time critical actions are involved. (7) Actuate both trim switches on the copilot’s control wheel to the NOSE UP position; the trim wheel begins moving. After 3 to 5 seconds, perform Step (8). (8) Actuate both trim switches on the pilot’s control wheel to the NOSE DOWN position. As soon as the trim wheel reverses its direction of travel, release ALL trim buttons. This verifies pilot override. (9) Actuate both trim buttons on the pilot’s control wheel to the NOSE UP position and, while trim tab is in travel, press the red disconnect switch on the control wheel to the second detent position and release. Note the PITCH TRIM OFF annunciator is illuminated and the elevator trim system is deactivated. (10) Activate the elevator trim system by setting the elevator trim switch to OFF then ON and repeat Step (9) on the copilot’s control wheel. (11) Set the elevator trim switch to the OFF position and manually rotate the elevator trim wheel to the stops in both directions to check for freedom of movement. Repeat this Step with the elevator trim switch set to the ON position and verify freedom of movement. (12) Remove electrical power from the airplane.

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FLIGHT CONTROLS ELEVATOR TRIM TAB CABLES MAINTENANCE PRACTICES

27-30-04 200200

1. FORWARD ELEVATOR TRIM TAB CABLE A. Removal CAUTION: If tension in the system is lost during cable maintenance, check the forward cable drum and elevator trim tab actuator cable drums to ensure the cable is not unwound from drums (Ref. Figure 201). (1) Remove the lower aft rig pin access panels 34 on the left and right side of the upper vertical stabilizer (Ref. Chapter 6-50-00, STABILIZER ACCESS PANELS). (2) Move the elevator surface (1) to the neutral position (0°) and install a rig pin (3) (7, Table 1, 27-00-00) in the elevator aft bellcrank through vertical stabilizer (4). Look at the opposite side of the vertical stabilizer to verify that the rig pin has protruded. Using minimum force try to manually move the elevators up and down to verify that the rig pin is correctly installed and that the elevators do not move (Ref. Figure 202). (3) Remove belly access panel 3 just aft of the nose landing gear wheel well (Ref. Chapter 6-50-00, FUSELAGE ACCESS PANELS). (4) Remove both flight compartment carpet and seats (Ref. Chapter 25-10-00). (5) Remove the left and right pedestal side access panels. (6) Remove floor access panels 1 left, 2 left, 4, 21 and 23 to gain access to the trim cables being removed (Ref. Chapter 6-50-00, FLOOR ACCESS PANELS). (7) Remove the environmental outlet duct assembly located under the pilot seat. (8) Remove the left side passenger seats and carpets as required to gain access to the trim cables being removed (Ref. Chapter 25-20-00). (9) Remove the left side passenger compartment floor access panels 16 as required to gain access to the trim cables being removed (Ref. Chapter 6-50-00, FLOOR ACCESS PANELS). (10) Remove the tie straps (1) securing the conduit tubes (6) (Ref. Figure 203). (11) Slide the conduit tubes (6) forward as required to gain access to the turnbuckles (2 and 9). (12) On the left side of the upper pedestal, move the elevator trim control wheel to approximately align the forward cable terminal ends (3 and 4). (13) Attach cable block (5) to the left and right aft cable, forward of the bracket (8) (FS 408.25) to prevent all of the trim cables aft of the forward cable from moving. (14) If installed, identify, tag and disconnect the Supplemental Type Certificate (STC) Flight Data Recorder (FDR) trim sensor bridle clamp (7) located on the forward cables between the main and rear wing spars (Ref. Figure 204). Refer to the STC holders instructions. NOTE: Each turnbuckle barrel has a groove at one end to identify the left-hand threaded end.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (15) Attach a tag with the words “forward cable left-hand threads terminal end” to the forward end of the turnbuckle (9) (Ref. Figure 203). (16) Disconnect the left-hand threads terminal end (4) from turnbuckle (9) and attach a feed line to the terminal end. Label the feed line with the words “left-hand threads terminal end”. (17) Attach a tag with the words “forward cable right-hand threads terminal end” to the forward end of the turnbuckle (2). (18) Disconnect the right-hand threads terminal end (3) from turnbuckle (2) and attach a feed line to the terminal end. Label the feed line with the words “right-hand threads terminal end”. (19) Remove cable guard pins from pulley brackets. Refer to Figure 201 for general location of the pulleys. NOTE: If reusing cable, take care to keep the cable clean and free from damage. Coil cable loosely no tighter than a ten inch diameter. (20) While pulling the feed lines through the fuselage, withdraw both left and right-hand threads terminal ends through the fuselage and out of the belly access panel 3 (Ref. Chapter 6-50-00, FUSELAGE ACCESS PANELS). (21) On the left side of the pedestal, loosen bolts (11) on the idler sprocket bracket (10) and relieve tension from the chain (8) (Ref. Figure 206). NOTE: When bolt (13) is removed from the left side of the pedestal, washer (14), located between the pedestal frame and the sprocket (12), may fall. (22) Remove safety wire from bolts (13) and remove bolts (13) at each end of the trim shaft (16). Remove washer (14) between pedestal and sprocket (12). Lift the trim shaft assembly as needed to remove the chain (8) from the sprocket (12), and remove the trim shaft assembly from the pedestal with the forward cable attached. (23) Remove the sprocket (12), cable guard (15) and drum (3) from the trim shaft (16). (24) Unwrap cable from the drum and remove the cable lock pin (Ref. Figure 205). (25) Remove the forward cable from the airplane while routing the feed lines through the right side of the pedestal, and into the cockpit. (26) Disconnect the feed lines from forward cable left and right-hand threads terminal ends.

B. Installation WARNING: Perform all Steps of this procedure, in the order listed. Do not skip any Steps of this procedure. Failure to do so may result in injury to personnel and damage to equipment. CAUTION: If tension in the system is lost during cable maintenance, check the forward cable drum and elevator trim tab actuator cable drums to ensure the cable is not unwound from drums (Ref. Figure 201). (1) Check cable for cleanliness and damage. Dip the cable in corrosion preventive compound (11, Table 1, Chapter 91-00-00). Remove excess corrosion preventative by wiping with a clean cloth.

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27-30-04

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (2) Attach a tag labeled “left-hand threads terminal end” to the forward cable left-hand threads terminal end. (3) Attach a tag labeled “right-hand threads terminal end” to the forward cable right-hand threads terminal end. (4) Wrap the forward left-hand threads terminal end and right-hand threads terminal end on the cable drum as follows (Ref. Figure 205): CAUTION: Do not kink the cable while locating the middle of the forward cable. Damage to the cable will occur. (a) Align the terminal ends of the forward cable together and mark the middle of the forward cable with ink or paint. Locate the side of the forward cable with the left-hand threads terminal end. Position the middle of the cable in the middle of the cable drum slot (Ref. Detail B). With the left-hand threads terminal end side of the cable, on the flat side of the cable drum, install the cable lock pin in the middle of the cable drum slot (Ref. Detail C). (b) From the lock pin, wrap each cable 2 1/4 turns around the drum beginning with the outside grooves and work toward the middle of the drum (Ref. Detail D). (c) Position the cable guard over the drum and tape the forward cables together just outside of the cable guard to prevent cable backlash at the drum (Ref. Detail E). When applying tape to the cable, make sure the cables are separated so that it is easy to identify which cable end winds off the forward and aft side of the drum. (5) Attach the right-hand threads terminal end to the feed line labeled “right-hand threads terminal end”. (6) Attach the left-hand threads terminal end to the feed line labeled “left-hand threads terminal end”. NOTE: More than one person will be required to route the forward cable. Take precautions to keep the cable clean and free from damage. (7) Pull the feed lines from the belly access panel 3 (Ref. Chapter 6-50-00, FUSELAGE ACCESS PANELS), draw the forward cables through the right side of the pedestal and then proceeding inside of pedestal to the left side and then out of the belly access panel, until the drum is close to the pedestal. (8) On the left side of the upper pedestal, move the elevator trim control wheel (7) to approximately 0 position (Ref. Figure 206). (9) Install the cable drum (3) with guard (15) and sprocket (12) on the trim shaft (16). Carefully position the trim shaft assembly and washer (14) in the pedestal, and install the lower end of the chain (8) over the sprocket (12). (10) Install and safety wire the bolts (13), one on each side of the pedestal.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

(11) Identify the forward cable with the left-hand threads terminal end (2) and make sure it winds off the forward side of the drum as installed. Identify the forward cable with the right-hand threads terminal end (1) and make sure it winds off the aft side of the drum as installed. Route the cable from the drum as follows (Ref. Figures 204 and 206): NOTE: It is permissible to install cable guard pins as the cable is being routed. (a) Route cable (2) over the outboard pulley (4) (1st set of pulleys, under the pedestal drum) (Ref. Figure 206). (b) Route cable (1) over the inboard pulley (4) (1st set of pulleys, under the pedestal drum). (c) Route cable (2) through aft pulley (5) (2nd set of pulleys, under the pedestal, just above the belly access panel). (d) Route cable (1) through forward pulley (5) (2nd set of pulleys, under the pedestal, just above the belly access panel). (e) Using the feed lines, pull the cable into the fuselage and continue with Step (f). (f) Route cable (2) over the bottom pulley (6) (3rd set of pulleys, at FS 103.00 under the pilot). This cable continues as the bottom cable under the pilot and into the forward cabin through the fairleads. (g) Route cable (1) over the top pulley (6) (3rd set of pulleys, at FS 103.00 under the pilot). This cable continues as the top cable under the pilot and into the forward cabin through the fairleads. (12) Using the feed lines, pull the cable ends through the fuselage and conduit tubes (6) to the turnbuckle connections (Ref. Figure 203). (13) Lubricate all turnbuckles with grease (23, Table 1, Chapter 91-00-00) for corrosion protection prior to installation. NOTE: Each turnbuckle barrel has a groove at one end to identify the left-hand threaded end. (14) Remove feed line, and attach the left-hand threads terminal end (4) to the turnbuckle (9). (15) Remove feed line, and attach the right-hand threads terminal end (3) to the turnbuckle (2). (16) Tension the forward cable sufficient to prevent slack. (17) Make sure that the forward cable is routed properly and is engaged in the pulleys. Verify Steps (11) (a) thru (11) (d) and Steps (11) (f) thru (11) (g). (18) Remove slack from the tab control chain (8) and tighten the bolts (11) on the idler sprocket bracket (10) (Ref. Figure 206). (19) Remove cable block (5) (Ref. Figure 203). (20) Remove all tape from the cable, turnbuckles, and pulleys. (21) Install all cable guard pins removed during the FORWARD - ELEVATOR TRIM TAB CABLE REMOVAL procedure.

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27-30-04

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (22) Ensure bolts (13) have been safety wired (Ref. Figure 206). (23) Move the top of the trim control wheel and note that the top cable, aft of the 3rd set of pulleys (6), under the pilot’s seat, moves the same direction as the top of the trim control wheel. No binding is allowed. (24) Perform the ELEVATOR TRIM TAB RIGGING procedure (Ref. 27-30-05). (25) Move the top of the trim control wheel to position the elevator trim tabs at neutral with the elevators at the neutral position. If necessary, adjust the indicator in the flight compartment to 0 while the tabs and elevators are at neutral. The 0 mark on the trim control wheel indicator must align with the triangle mark on the pedestal edgelighted panel. Perform the ELEVATOR TRIM TAB INDICATOR ADJUSTMENT procedure (Ref. 27-30-08). (26) If installed, connect the Supplemental Type Certificate (STC) Flight Data Recorder (FDR) trim sensor bridle clamp (7) (Ref. Figure 204). Refer to the STC holders instructions. (27) Ensure turnbuckles (2 and 9) have been safetied. Install conduit tubes (6) over the turnbuckles (2 and 9) between brackets (8) FS 408.25 and (7) FS 378.25. Secure the conduit tubes (6) with tie straps (1), forward of bracket (7) and aft of bracket (8) (Ref. Figure 203). (28) Move the top of the trim control wheel forward (airplane nose down) and verify the trim tab trailing edge moves to the full up position. Looking inboard from the pilot side at the trim wheel, turn the wheel counterclockwise, proceed to the tail and verify the trim tab moved up. (29) Move the top of the trim control wheel aft (airplane nose up) and verify the trim tab trailing edge moves to the full down position. Looking inboard from the pilot side at the trim wheel, turn the wheel clockwise, proceed to the tail and verify the trim tab moved down. (30) Install the left side passenger compartment floor access panels 16 (Ref. Chapter 6-50-00, FLOOR ACCESS PANELS). (31) Install the left side passenger seats and carpets (Ref. Chapter 25-20-00). (32) Install the environmental outlet duct assembly located under the pilot seat. (33) Install floor access panels 1 left, 2 left, 4, 21 and 23 (Ref. Chapter 6-50-00, FLOOR ACCESS PANELS). (34) Install left and right pedestal side access panels. (35) Install flight compartment carpet and seats (Ref. Chapter 25-10-00). (36) Install belly access panel 3 just aft of the nose landing gear wheel well (Ref. Chapter 6-50-00, FUSELAGE ACCESS PANELS). (37) Remove the rig pin from the elevator aft bellcrank (Ref. Figure 202). (38) Install the lower aft rig pin access panels 34 on the left and right side of the upper vertical stabilizer (Ref. Chapter 6-50-00, STABILIZER ACCESS PANELS). CAUTION: With the control column pulled to the aft position, allowing the control column to free fall to the forward position can cause damage to the elevator system.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (39) Pull the pilot’s control wheel aft and make sure that the elevator travels to the full up position with no unusual noise or binding. (40) Move the pilot’s control wheel forward and make sure that the elevator travels to the full down position with no unusual noise or binding. (41) Perform the ELEVATOR ELECTRIC TRIM OPERATIONAL CHECK (Ref. 27-30-03). If airplane is not equipped with electric trim, perform the ELEVATOR TRIM TAB OPERATIONAL CHECK (Ref. 27-30-05).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

1

19

20 13

20

25 2

20

18

15

16

13

13

J D

4

24

5

DETAIL

17

18

LEFT SIDE SHOWN, RIGHT SIDE TYPICAL

B

DETAIL

J

14

15

13 15

14

6

15

G14 E

C

20

16

25

4

14

14

7 3

E G15

13

A

20

B

15

G

13

E

14 CABLE STOPS

F

PRESSURE SEAL DETAIL

5

16

17

F

27 FS 408.25

H

LOOKING AFT DETAIL

D

26 TURNBUCKLE 15 13 DETAIL 14

14

DETAIL

11 9

10

13

12

G

G

C 21

22

21

11 12 DETAIL

A

12

J

23

6

8

DETAIL

FS 378.25

12 11 VIEW LOOKING DOWN DETAIL

H

FWD

1. TAB CONTROL WHEEL 2. IDLER SPROCKET 3. IDLER SPROCKET BRACKET 4. CABLE DRUM 5. FORWARD CABLE (LEFT-HAND THREADS TERMINAL END) 6. FORWARD CABLE (RIGHT-HAND THREADS TERMINAL END) 7. TRIM SHAFT 8. 1ST PULLEY (UNDER PEDESTAL) 9. 2ND PULLEY (UNDER PEDESTAL) 10. 3RD PULLEY (FS 103.00) 11. FORWARD CABLE 12. FAIR LEAD 13. LEFT-HAND THREADS TERMINAL END 14. RIGHT-HAND THREADS TERMINAL END 15. TURNBUCKLE 16. RIGHT VERTICAL CABLE 17. LEFT VERTICAL CABLE 18. HORIZONTAL CABLE 19. ELEVATOR TRIM TAB ACTUATOR 20. ACTUATOR CABLE 21. STRAPS 22. CONDUIT TUBES 23. CLIPS 24. WASHER 25. BOLTS 26. LEFT AFT CABLE UC27B 27. RIGHT AFT CABLE 034894AB.AI

Figure 201 Elevator Trim Tab Control System

27-30-04

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This Page Intentionally Left Blank

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

A

1

1. RIG PIN VIEW LOOKING UP LEFT HAND SIDE DETAIL

A UC27B 034878AA.AI

Figure 202 (Sheet 1 of 2) Elevator Aft Bellcrank Rig Pin Installation

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

6. ELEVATOR AFT BELLCRANK 7. ELEVATOR AFT RIG PIN HOLE

6

FWD 7

VERTICAL STABILIZER (REF)

DETAIL

B UC27B 040558AB.AI

Figure 202 (Sheet 2 of 2) Elevator Aft Bellcrank Rig Pin Installation

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

1. TIE STRAPS 2. TURNBUCKLE 3. RIGHT - HAND THREADS TERMINAL END 4. LEFT - HAND THREADS TERMINAL END 5. BLOCK (TEMPORARY) 6. CONDUIT TUBES 7. BRACKET 8. BRACKET 9. TURNBUCKLE

A

1 6 1

VIEW LOOKING DOWN DETAIL

FS 408.25 (REF) 5

2

A 3

FWD

7

FS 378.25 (REF) 6

9 4 VIEW LOOKING DOWN (CONDUIT TUBES MOVED FORWARD AND BLOCK TEMPORARILY INSTALLED) 8

DETAIL

FWD UC27B 034861AA.AI

A

Figure 203 Forward Trim Tab Cable

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1. FORWARD CABLE (RIGHT-HAND THREADS TERMINAL END) 2. FORWARD CABLE (LEFT-HAND THREADS TERMINAL END) 3. FAIRLEADS 4. TURNBUCKLES 5. LEFT - HAND THREADS TERMINAL END 6. RIGHT - HAND THREADS TERMINAL END 7. BRIDLE CLAMP

A 3

3

4 3

5

1 3 3

6

2 FLIGHT DATA RECORDER (REF) (IF INSTALLED)

2

7 1

DETAIL

A UC27B 034835AA.AI

Figure 204 Forward Trim Tab Cable

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

D A

SPROCKET

B C

E DETAIL

FORWARD CABLE WITH LEFT - HAND THREADS TERMINAL END

A MIDDLE OF CABLE AND MIDDLE OF CABLE DRUM SLOT

SIDE WITH RIGHT-HAND THREADS TERMINAL END

CABLE LOCK PIN INSTALLED

CABLE LOCK PIN

FLAT SIDE OF DRUM FORWARD CABLE WITH RIGHT - HAND THREADS TERMINAL END

(CABLE GUARD NOT SHOWN) DETAIL

SIDE WITH LEFT - HAND THREADS TERMINAL END

B

(DRUM NOT FULLY WRAPPED) (CABLE GUARD NOT SHOWN) DETAIL

C

CABLE WITH RIGHT-HAND THREADS TERMINAL END BEGINNING AT THE SPROCKET SIDE OF THE DRUM

CABLE WITH LEFT-HAND THREADS TERMINAL END BEGINNING AT THE FLAT SIDE OF THE DRUM

CABLE GUARD SPROCKET SIDE OF DRUM FORWARD CABLE RIGHT - HAND THREADS TERMINAL END

FORWARD CABLE WITH LEFT - HAND THREADS TERMINAL END

TAPE (TEMPORARY)

DETAIL

(DRUM FULLY WRAPPED)

E

DETAIL

D

UC27B 034837AA.AI

Figure 205 Forward Trim Tab Cable

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

1

2

3

4

A

B 1 6

5 2

2

1 FORWARD LOOKING AFT DETAIL

7

B

8

17

1. FORWARD (RIGHT - HAND THREADS TERMINAL END) CABLE 2. FORWARD (LEFT - HAND THREADS TERMINAL END) CABLE 3. DRUM 4. 1ST SET OF PULLEYS (UNDER PEDESTAL) 5. 2ND SET OF PULLEYS (UNDER PEDESTAL) 6. 3RD SET OF PULLEYS (FS 103.00) 7. TAB CONTROL WHEEL 8. CHAIN 9. IDLER SPROCKET 10. IDLER SPROCKET BRACKET 11. BOLTS 12. LOWER SPROCKET 13. BOLTS 14. WASHER 15. CABLE GUARD 16. TRIM SHAFT 17. UPPER SPROCKET

15

11 3 12

2

1

14 13

Figure 206 Forward Trim Tab Cable

27-30-04

16

10

DETAIL

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13

9

A

UC27B 034834AB.AI

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

2. AFT ELEVATOR TRIM TAB CABLE A. Removal CAUTION: If tension in the system is lost during cable maintenance, check the forward cable drum and elevator trim tab actuator cable drums to ensure the cable is not unwound from drums (Ref. Figure 201). NOTE: Removal procedures for the left and right aft elevator trim cables are the same, only the procedure for the left aft elevator trim cable is provided, unless otherwise indicated. (1) Remove the left side passenger seats and carpets as required to gain access to the trim cables being removed (Ref. Chapter 25-20-00). (2) Remove the left side passenger compartment floor access panels 16 as required to gain access to the trim cables being removed (Ref. Chapter 6-50-00, FLOOR ACCESS PANELS). (3) Remove floor access panel 15 (Ref. Chapter 6-50-00, FLOOR ACCESS PANELS). (4) Remove cargo compartment access panels aft of cargo door. (5) Remove stabilizer access panels 20 and 25 (Ref. Chapter 6-50-00, STABILIZER ACCESS PANELS). (6) Remove fuselage access panels 7 and 8 (Ref. Chapter 6-50-00, FUSELAGE ACCESS PANELS). (7) Remove the lower aft rig pin access panels 34 on the left and right side of the upper vertical stabilizer (Ref. Chapter 6-50-00, STABILIZER ACCESS PANELS). (8) Move the elevator surface (1) to the neutral position (0°) and install a rig pin (3) (7, Table 1, 27-00-00) in the elevator aft bellcrank through vertical stabilizer (4). Look at the opposite side of the vertical stabilizer to verify that the rig pin has protruded. Using minimum force try to manually move the elevators up and down to verify that the rig pin is correctly installed and that the elevators do not move (Ref. Figure 202). (9) Remove the tie straps (1) securing the conduit tubes (2) (Ref. Figure 207). (10) Slide the conduit tubes (2) forward as required to gain access to the turnbuckles (5 and 7). (11) On the left side of the upper pedestal, move the elevator trim control wheel so that turnbuckle (4) is easily accessible (Ref. Figure 208, Detail C). (12) Attach cable block (4) to both ends of the forward cable (10), aft of bracket (3) (FS 378.25) to prevent the forward cable (10) from moving (Ref. Figure 207). (13) Attach cable block (6) on the left vertical cable (5) at the left upper vertical cable pulley to maintain tension on the trim cables that are not being removed (Ref. Figure 208, Detail D). NOTE: Each turnbuckle barrel has a groove at one end to identify the left-hand threaded end. (14) Attach a tag with the words “left-hand threads terminal end” to the aft end of the turnbuckle (7) (Ref. Figure 207). (15) Disconnect the left-hand threads terminal end (8) from turnbuckle (7) and attach a feed line to the terminal end.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (16) Disconnect the right-hand threads terminal end (3) from turnbuckle (4) (Ref. Figure 208, Detail C). (17) Remove cable guard pins from the aft cable pulley brackets. NOTE: If reusing cable, take care to keep the cable clean and free from damage. Coil cable loosely no tighter than a ten inch diameter. (18) Lubricate the left aft cable (1) with grease (23, Table 1, Chapter 91-00-00) as necessary to pull the cable terminal end through the pressure seal (4) (Ref. Figure 209). (19) With the right-hand threads terminal end (3), pull the cable and feed line through the frames, aft pressure bulk head and pulleys, to remove the left aft cable (1) from the airplane. Use care while drawing the left-hand threads terminal end through the pressure seal to prevent damage to the seal. Wipe excessive grease from airplane, aft cable and cable terminal end with a clean cloth (Ref. Figure 208). (20) Disconnect the feed lines from left aft cable (11) left-hand threads terminal end (8) (Ref. Figure 207).

B. Installation CAUTION: If tension in the system is lost during cable maintenance, check the forward cable drum and elevator trim tab actuator cable drums to ensure the cable is not unwound from drums (Ref. Figure 201). NOTE: Installation procedures for the left and right aft elevator trim cables are the same, only the procedure for the left aft elevator trim cable is provided, unless otherwise indicated. (1) Clean the cable assembly with a clean cloth saturated with solvent (2, Table 1, Chapter 91-00-00). (2) Check cable for cleanliness and damage. Dip the cable in corrosion preventive compound (11, Table 1, Chapter 91-00-00). Remove excess corrosion preventative by wiping with a clean cloth. (3) Attach the left aft cable (1) right-hand threads terminal end (3) to the feed line, in the passenger compartment. The feed line was installed during the AFT ELEVATOR TRIM TAB CABLE REMOVAL procedure (Ref. Figure 208). (4) Lubricate the left aft cable (1) with grease (23, Table 1, Chapter 91-00-00) as required to pull the cable terminal end through the pressure seal (4). With assistance, route the left aft cable (1) through pulleys (3), aft pressure bulk head and frames, by pulling the feed line aft. Use care while drawing the right-hand threads terminal end through the pressure seal (4) to prevent damage to the pressure seal (4). Wipe excessive grease from right-hand threads terminal end, left aft cable and airplane with a clean cloth after drawing the left aft cable (1) through the pressure seal (4) (Ref. Figure 209). (5) Lubricate all turnbuckles with grease (23, Table 1, Chapter 91-00-00) for corrosion protection prior to installation. NOTE: Each turnbuckle barrel has a groove at one end to identify the left-hand threaded end. (6) Remove feed line and attach the left aft cable (1) right-hand threads terminal end (3) to turnbuckle (4) (Ref. Figure 208). (7) Attach the left aft cable (11) left-hand threads terminal end (8) to turnbuckle (7) (Ref. Figure 207). Page 216 Nov 1/09

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (8) Tension the left aft cable (11) sufficient to prevent slack. Ensure the left aft cable is routed properly and engaged in the pulleys. (9) Install the cable guard pins removed during the AFT ELEVATOR TRIM TAB CABLE REMOVAL procedure. (10) Remove the cable blocks (Ref. Figure 207, Item 4 and Figure 208, Item 6). (11) Fill the pressure seal (4) and lubricate the left aft cable (1) to one inch beyond the length of travel through the pressure seal with grease (23, Table 1, Chapter 91-00-00) (Ref. Figure 209). (12) Remove all tags from turnbuckles. (13) Move the top of the trim control wheel forward (airplane nose down) and verify the trim tab trailing edge moves to the full up position. Move the top of the trim control wheel aft (airplane nose up) and verify the trim tab trailing edge moves to the full down position. (14) Perform the ELEVATOR TRIM TAB RIGGING procedure (Ref. Chapter 27-30-05). (15) Ensure turnbuckle (7) has been safetied and slide conduit tubes (2) over the turnbuckles (7 and 5) between brackets (9) FS 408.25 and (3) FS 378.25. Secure the conduit tubes (2) with tie straps (1), forward of bracket (3) and aft of bracket (9) (Ref. Figure 207). (16) Looking inboard from the pilot side at the trim wheel, turn the wheel counterclockwise and proceed to the tail and verify the trim tab moved up. (17) Looking inboard from the pilot side at the trim wheel, turn the wheel clockwise and proceed to the tail and verify the trim tab moved down. (18) Install the left side passenger compartment floor access panels 16 (Ref. Chapter 6-50-00, FLOOR ACCESS PANELS). (19) Install floor access panel 15 (Ref. Chapter 6-50-00, FLOOR ACCESS PANELS). (20) Install cargo compartment access panels aft of cargo door. (21) Install the left side passenger seats and carpets (Ref. Chapter 25-20-00). (22) Install fuselage access panels 7 and 8 (Ref. Chapter 6-50-00, FUSELAGE ACCESS PANELS). (23) Install stabilizer access panels 20 and 25 (Ref. Chapter 6-50-00, STABILIZER ACCESS PANELS). (24) Remove the rig pin from the elevator aft bellcrank (Ref. Figure 202). (25) Install the lower aft rig pin access panels 34 on the left and right side of the upper vertical stabilizer (Ref. Chapter 6-50-00, STABILIZER ACCESS PANELS). CAUTION: With the control column pulled to the aft position, allowing the control column to free fall to the forward position can cause damage to the elevator system. (26) Pull the pilot’s control wheel aft and make sure that the elevator travels to the full up position with no unusual noise or binding.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (27) Move the pilot’s control wheel forward and make sure that the elevator travels to the full down position with no unusual noise or binding. (28) Perform the ELEVATOR ELECTRIC TRIM OPERATIONAL CHECK (Ref. Chapter 27-30-03). If airplane is not equipped with electric trim, perform the ELEVATOR TRIM TAB OPERATIONAL CHECK (Ref. Chapter 27-30-05).

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1. TIE STRAPS 2. CONDUIT TUBES 3. BRACKET 4. CABLE BLOCK 5. TURNBUCKLE 6. RIGHT - HAND THREADS TERMINAL END

7. TURNBUCKLE 8. LEFT - HAND THREADS TERMINAL END 9. BRACKET 10. FORWARD CABLE 11. LEFT AFT CABLE 12. RIGHT AFT CABLE

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UC27B 040770AA.AI

Figure 207 Aft Trim Tab Cable FS 393.25

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AFT PRESSURE BULKHEAD (REF)

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1. LEFT AFT CABLE 2. PULLEY 3. RIGHT-HAND THREADS TEMINAL END 4. TURNBUCKLE 5. LEFT VERTICAL CABLE 6. CABLE BLOCK 7. PULLEY 8. PRESSURE SEAL

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Figure 208 Aft Trim Tab Cable FS 541.0 to CS 605.98

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UC27B 040771AA.AI

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

A 1. LEFT AFT CABLE 2. RIGHT AFT CABLE 3. PULLEYS 4. PRESSURE SEALS

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UC27B 040736AA

Figure 209 Aft Trim Tab Cable FS 523.5

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3. VERTICAL ELEVATOR TRIM TAB CABLE A. Removal CAUTION: If tension in the system is lost during cable maintenance, check the forward cable drum and elevator trim tab actuator cable drums to ensure the cable is not unwound from drums (Ref. Figure 201). NOTE: The left vertical cable is connected to the right actuator cable and the right vertical cable is connected to the left actuator cable. Removal procedures for the left and right vertical trim tab cables are the same, only the left vertical trim tab cable procedure is provided, unless otherwise indicated. Both terminal ends of the left vertical cable are left-hand threads terminal ends and both terminal ends of the right vertical cable are right-hand threads terminal ends. (1) Remove the stabilizer access panels 18, 21, and 22 (Ref. Chapter 6-50-00, STABILIZER ACCESS PANELS). (2) Remove fuselage access panels 7 and 8 (Ref. Chapter 6-50-00, FUSELAGE ACCESS PANELS). (3) Remove the lower aft rig pin access panels 34 on the left and right side of the upper vertical stabilizer (Ref. Chapter 6-50-00, STABILIZER ACCESS PANELS). (4) Move the elevator surface (1) to the neutral position (0°) and install a rig pin (3) (7, Table 1, 27-00-00) in the elevator aft bellcrank through vertical stabilizer (4). Look at the opposite side of the vertical stabilizer to verify that the rig pin has protruded. Using minimum force try to manually move the elevators up and down to verify that the rig pin is correctly installed and that the elevators do not move (Ref. Figure 202). NOTE: For airplanes without an electric elevator trim tab system, proceed to Step (8). (5) Move the electric elevator trim tab system to position the electric trim turnbuckles (15) so they are easily accessible (Ref. Figure 210). NOTE: With one electric elevator trim tab system cable disconnected, do not move the elevator trim tab system cables, to prevent the electric elevator trim tab system cable (14) from unwinding from the electric elevator trim tab system drum (13). (6) While maintaining tension on the electric elevator trim tab system drum (13), disconnect, one at a time, the left and right turnbuckles (15), and secure each cable (14) to the structure to prevent cable (14) from unwinding from the drum (13). (7) Remove the bridle cable clamps (8) from the vertical cables. (8) Move turnbuckle (17) so that cable block (5) can be installed and turnbuckle (10) is easily accessible. (9) Attach cable block (12) on the left aft cable (11) to prevent slack in the trim tab cables forward of the vertical cable. (10) Attach cable block (5) on both cable ends of the right actuator cable (4) against trim tab actuator housing (7), to prevent the actuator drum (6) from unwinding.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (11) Disconnect the vertical cable lower terminal end (9) from the turnbuckle (10) and attach a feed line to terminal end (9). (12) Disconnect vertical cable upper terminal end (2) from the turnbuckle (3). (13) Disconnect cable guard pins from the vertical cable pulley brackets. NOTE: If reusing cable, take care to keep the cable clean and free from damage. Coil cable loosely no tighter than a ten inch diameter. (14) Remove vertical cable from airplane, drawing the feed line through stabilizer access panel (16). (15) Disconnect the feed line from the vertical cable.

B. Installation CAUTION: If tension in the system is lost during cable maintenance, check the forward cable drum and elevator trim tab actuator cable drums to ensure the cable is not unwound from drums (Ref. Figure 201). NOTE: The left vertical cable is connected to the right actuator cable and the right vertical cable is connected to the left actuator cable. Installation procedures for the left and right vertical trim tab cables are the same, only the procedure for the left vertical trim tab cable is provided, unless otherwise indicated. Both terminal ends of the left vertical cable are left-hand threads terminal ends and both terminal ends of the right vertical cable are right-hand threads terminal ends. (1) Check cable for cleanliness and damage. Dip the cable in corrosion preventive compound (11, Table 1, Chapter 91-00-00). Remove excess corrosion preventative by wiping with a clean cloth. (2) Attach the left vertical cable (1) to the feed line at the stabilizer access panel (16) (Ref. Figure 210). (3) With assistance, route the vertical cable down through vertical stabilizer to the turnbuckle (10). (4) Lubricate all turnbuckles with grease (23, Table 1, Chapter 91-00-00) for corrosion protection prior to installation. (5) Connect the left vertical cable upper terminal end (2) to the actuator cable turnbuckle (3). (6) Install all cable guard pins removed during the VERTICAL ELEVATOR TRIM TAB CABLE REMOVAL procedure. (7) Remove the feed line and connect the vertical cable lower terminal end (9) to the turnbuckle (10). (8) Tension the left vertical cable (1) sufficient to prevent slack. Ensure the left vertical cable is routed properly and engaged in the pulleys. (9) Remove cable blocks (5 and 12). (10) Move the top of the trim control wheel forward (airplane nose down) and verify the trim tab trailing edge moves to the full up position. Move the top of the trim control wheel aft (airplane nose up) and verify the trim tab trailing edge moves to the full down position.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (11) Perform the ELEVATOR TRIM TAB RIGGING procedure (Ref. Chapter 27-30-05). (12) Ensure turnbuckles (3 and 10) have been safetied. (13) Looking inboard from the pilot side at the trim wheel, turn the wheel counterclockwise, proceed to the tail and verify the trim tab moved up. (14) Looking inboard from the pilot side at the trim wheel, turn the wheel clockwise, proceed to the tail and verify the trim tab moved down. NOTE: For airplanes without an electric elevator trim tab system, proceed to Step (17). (15) Perform the ELEVATOR ELECTRIC TRIM TAB CABLE RIGGING (ALL AIRPLANES WITHOUT COLLINS APS-65H AUTOPILOT SYSTEM) procedure (Ref. Chapter 27-30-07). (16) Ensure turnbuckles (15) have been safetied. (17) Install fuselage access panels 7 and 8 (Ref. Chapter 6-50-00, FUSELAGE ACCESS PANELS). (18) Install the stabilizer access panels 18, 21, and 22 (Ref. Chapter 6-50-00, STABILIZER ACCESS PANELS). (19) Remove the rig pin from the elevator aft bellcrank (Ref. Figure 202). (20) Install the lower aft rig pin access panels 34 on the left and right side of the upper vertical stabilizer (Ref. Chapter 6-50-00, STABILIZER ACCESS PANELS). CAUTION: With the control column pulled to the aft position, allowing the control column to free fall to the forward position can cause damage to the elevator system. (21) Pull the pilot’s control wheel aft and make sure that the elevator travels to the full up position with no unusual noise or binding. (22) Move the pilot’s control wheel forward and make sure that the elevator travels to the full down position with no unusual noise or binding. (23) Perform the ELEVATOR ELECTRIC TRIM OPERATIONAL CHECK (Ref. Chapter 27-30-03). If airplane is not equipped with electric trim, perform the ELEVATOR TRIM TAB OPERATIONAL CHECK (Ref. Chapter 27-30-05).

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10 1. LEFT VERTICAL CABLE 2. VERTICAL CABLE UPPER TERMINAL END 3. TURNBUCKLE 4. RIGHT ACTUATOR CABLE 5. CABLE BLOCK 6. ACTUATOR DRUM 7. TRIM ACTUATOR HOUSING 8. BRIDLE CLAMP 9. VERTICAL CABLE LOWER TERMINAL END 10. TURNBUCKLE 11. LEFT AFT CABLE 12. BLOCK 13. DRUM 14. ELECTRICAL TRIM TAB SYSTEM CABLES 15. TURNBUCKLES 16. STABILIZER ACCESS PANEL 17. TURNBUCKLE

11

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ELEVATOR CABLES REMOVED FOR CLARITY 12

DETAIL

B UC27B 034838AA.AI

Figure 210 Vertical Trim Tab Cable

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4. HORIZONTAL ELEVATOR TRIM TAB CABLE A. Removal CAUTION: If tension in the system is lost during cable maintenance, check the forward cable drum and elevator trim tab actuator cable drums to ensure the cable is not unwound from drums (Ref. Figure 201). (1) Remove the stabilizer access panels 18, 21, and 22 (Ref. Chapter 6-50-00, STABILIZER ACCESS PANELS). (2) Remove the lower aft rig pin access panels 34 on the left and right side of the upper vertical stabilizer (Ref. Chapter 6-50-00, STABILIZER ACCESS PANELS). (3) Move the elevator surface (1) to the neutral position (0°) and install a rig pin (3) (7, Table 1, 27-00-00) in the elevator aft bellcrank through vertical stabilizer (4). Look at the opposite side of the vertical stabilizer to verify that the rig pin has protruded. Using minimum force try to manually move the elevators up and down to verify that the rig pin is correctly installed and that the elevators do not move (Ref. Figure 202). (4) Move the elevator trim control wheel to position the horizontal cable turnbuckles (3 and 9) so that they are easily accessible (Ref. Figure 211). (5) Attach cable blocks (5) on both cable ends of the left and right actuator cables (4 and 8) against actuator housing (7), to prevent the actuator drum (6) from unwinding. (6) Disconnect the horizontal cable left-hand threads terminal end (10) from turnbuckle (9) and right-hand threads terminal end (2) from the turnbuckle (3). (7) Remove cable guard pin from horizontal cable pulley (11) bracket. NOTE: If cable is to be reused, take care to keep the cable clean and free from damage. Coil cable loosely no tighter than a ten inch diameter. (8) Remove horizontal cable from airplane.

B. Installation CAUTION: If tension in the system is lost during cable maintenance, check the forward cable drum and elevator trim tab actuator cable drums to ensure the cable is not unwound from drums (Ref. Figure 201). (1) Check cable for cleanliness and damage. Dip the cable in corrosion preventive compound (11, Table 1, Chapter 91-00-00). Remove excess corrosion preventative by wiping with a clean cloth. (2) Lubricate turnbuckles (3 and 9) with grease (23, Table 1, Chapter 91-00-00) (Ref. Figure 211). (3) Position the horizontal cable (1) on the horizontal cable pulley (11) and route the horizontal cable (1) through fairleads and horizontal stabilizer. (4) Connect the horizontal cable left-hand threads terminal end (10) to turnbuckle (9) and right-hand threads terminal end (2) to turnbuckle (3).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (5) Tension the horizontal cable (1) sufficient to prevent slack. Ensure the horizontal cable is routed properly and engaged in the pulleys. (6) Install cable guard pin in horizontal pulley (11) bracket. (7) Remove cable blocks (5). (8) Move the top of the trim control wheel forward (airplane nose down) and verify the trim tab trailing edge moves to the full up position. Move the top of the trim control wheel aft (airplane nose up) and verify the trim tab trailing edge moves to the full down position. (9) Perform the ELEVATOR TRIM TAB RIGGING procedure (Ref. Chapter 27-30-05). (10) Ensure turnbuckles (3 and 9) have been safetied. (11) Looking inboard from the pilot side at the trim wheel, turn the wheel counterclockwise, proceed to the tail and verify the trim tab moved up. (12) Looking inboard from the pilot side at the trim wheel, turn the wheel clockwise, proceed to the tail and verify the trim tab moved down. (13) Install the stabilizer access panels 18, 21, and 22 (Ref. Chapter 6-50-00, STABILIZER ACCESS PANELS). (14) Remove the rig pin from the elevator aft bellcrank (Ref. Figure 202). (15) Install the lower aft rig pin access panels 34 on the left and right side of the upper vertical stabilizer (Ref. Chapter 6-50-00, STABILIZER ACCESS PANELS). CAUTION: With the control column pulled to the aft position, allowing the control column to free fall to the forward position can cause damage to the elevator system. (16) Pull the pilot’s control wheel aft and make sure that the elevator travels to the full up position with no unusual noise or binding. (17) Move the pilot’s control wheel forward and make sure that the elevator travels to the full down position with no unusual noise or binding. (18) Perform the ELEVATOR ELECTRIC TRIM OPERATIONAL CHECK (Ref. Chapter 27-30-03). If airplane is not equipped with electric trim, perform the ELEVATOR TRIM TAB OPERATIONAL CHECK (Ref. Chapter 27-30-05).

5. ELEVATOR TRIM TAB CONTROL WHEEL CHAIN A. Removal (1) Remove the left flight compartment seat (Ref. 25-10-00, SEAT REMOVAL). (2) Remove the pedestal left side access panels, located below the elevator trim tab control wheel. (3) Rotate the elevator trim tab control wheel (11) until the chain master link is accessible (Ref. Figure 212). (4) Loosen bolts (8) and nuts (15) and move the idler sprocket (10) to provide slack in the chain (1).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL NOTE: A drop cloth should be used to keep items dropped from the pedestal from falling below the floor. (5) Remove the master link from the chain (1). NOTE: Chain guide screws (5, 7) may be loosened and moved aside to facilitate chain (1) removal. Leave the old chain in place until ready to install a new chain. The chain being removed will be used to guide the new chain on to and around the sprockets.

B. Installation (1) Use the master link and temporarily splice the new chain to the old chain that was left in place. (2) Using the old chain, slowly pull the new chain in to position around sprockets (6 and 12) (Ref. Figure 212). (3) Remove the master link and discard the old chain. (4) Install the master link in the new chain (1). (5) Position the trim bracket (4) and the idler sprocket bracket (9). Align the idler sprocket (10) with the chain (1) and press the idler sprocket (10) against the chain (1) to remove slack and tighten bolts (8) and nuts (15). (6) Adjust the chain guide screws (5, 7) to give a clearance of 0.03 to 0.09 inch between the screws (5, 7) and the chain (1). (7) Rotate the elevator trim tab control wheel (11) through full range of motion to verify smooth operation with no unusual noises or binding. (8) Install the pedestal left side access panel. (9) Install the left flight compartment seat (Ref. Chapter 25-10-00, SEAT INSTALLATION). (10) Perform the ELEVATOR TRIM TAB INDICATOR ADJUSTMENT procedure (Ref. 27-30-08).

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B C D DETAIL

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DETAIL

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UC27B 034840AA.AI

Figure 211 Horizontal Trim Tab Cable

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1. CHAIN 2. WASHER 3. PIVOT SHAFT 4. TRIM BRACKET 5. SCREW 6. LOWER SPROCKET 7. SCREW 8. BOLTS 9. IDLER SPROCKET BRACKET 10. IDLER SPROCKET 11. ELEVATOR TRIM TAB CONTROL WHEEL 12. UPPER SPROCKET 13. WASHER 14. SCREWS 15. NUTS

A B

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ADJUST SCREW IN SLOT TO GIVE 0.03 TO 0.09 INCH CLEARANCE BETWEEN CHAIN AND SCREW

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UC27B 084497AA.AI

Figure 212 Elevator Trim Tab Control Wheel Chain Installation

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FLIGHT CONTROLS ELEVATOR TRIM TAB CONTROL SYSTEM MAINTENANCE PRACTICES

27-30-05 200200

1. ELEVATOR TRIM TAB A. Rigging WARNING: Perform all Steps of this procedure, in the order listed. Do not skip any Steps of this procedure. Failure to do so may result in injury to personnel and damage to equipment. (1) Remove the left and right access panels 18, 21 and 22 (Ref. Chapter 6-50-00, STABILIZER ACCESS PANELS). (2) Remove access panel 34 on the left and right side of the upper vertical stabilizer (Ref. Chapter 6-50-00, STABILIZER ACCESS PANELS). WARNING: The elevator control system must be properly rigged before the elevator trim tab system can be rigged. Failure to do so may result in injury to personnel and damage to equipment. CAUTION: Where the cables pass through structure, the areas of possible contact between the control cables and adjacent structure must be protected with protective elements such as grommets, rub strips, block or guide fairings. Where contact of control cables does occur with the protective elements, a force no greater than eight (8) ounces shall be required to move the cable to a position of no contact. At no time should flight control cables contact metal structure with the protective elements removed. (3) If installed, perform the ELEVATOR ELECTRIC TRIM TAB CABLE DISCONNECTION procedure (Ref. 27-30-07). (4) If installed, identify, tag and disconnect the Supplemental Type Certificate (STC) Flight Data Recorder (FDR) trim sensor bridle clamp located on the forward cables between the main and rear spars. Refer to the STC holders instructions. (5) Move the elevator surface (1) and install a rig pin (3) (7, Table 1, 27-00-00) in the elevator aft bellcrank through the vertical stabilizer (4). Using minimum force try to manually move the elevators up and down to verify that the rig pin is correctly installed. Ensure that the elevators do not move. Look at the opposite side of the vertical stabilizer to verify that the rig pin has protruded (Ref. Figure 203). NOTE: One travel board may be used and moved from one side to the other. (6) Perform ELEVATOR TRAVEL BOARD INSTALLATION procedure (Ref. 27-00-02). Both elevators must be at neutral (0° deflection). Apply tape to the horizontal stabilizer upper surface to mark the location of the travel board feet to ensure proper location when moving the travel board from one side to the other. (a) If using the travel board at HSS 50.00, install a digital protractor (5, Table 1, 27-00-00) attached to the elevator trim tab surface. (b) If using travel board at HSS 35.00, install the elevator trim travel board to the elevator travel board.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (7) Check the cable tension by performing the following Steps: NOTE: Do not measure the tension when the temperature is varying rapidly or with the airplane located in direct sunlight. (a) Allow the temperature of the airframe to stabilize before measuring and adjusting the cable tension. (b) Measure the temperature in the compartment next to the elevator trim cables in the horizontal stabilizer area. (c) Refer to Elevator Trim Cable Tension Graph Figure 201, and read the pounds of tension for the measured temperature. NOTE: Cable tension tolerance is +3/ -2 lbs of the tension found in Figure 201. (d) Position a cable tensiometer (4, Table 1, 27-00-00) at least three inches from the turnbuckles and measure the cable tension of both cables. Cable diameter is noted in Figure 201. (e) If no adjustment is required, proceed to Step (8). If cable tension requires adjustment, remove safety clips from one or more turnbuckles. It is permissible to adjust any turnbuckle in the elevator trim system. WARNING: If cable tension at any time is below 5 pounds, check all elevator trim system cable drums and pulleys for proper cable engagement. (f) Adjust turnbuckle(s) to tension cables to the proper tension per Figure 201. Cable diameter is noted in Figure 201. (g) Rotate the elevator trim wheel from full up stop to full down stop through three cycles to distribute the cable tension throughout the elevator trim system. (h) Check cable tension to verify that the tension meets required tension per Figure 201. (i) Install safety clips on the turnbuckles. (8) Move the right side actuator upper cable (5) until the distance from the actuator housing (3) to the upper cable terminal end (7) is 5.50 to 5.75 inches (Ref. Figure 206, Detail A). (9) Check the position of the lower cable terminal end (6) of the right side elevator trim tab actuator. The distance between the actuator housing (3) and the outboard end of the cable terminal end (6) must be 12.75 inches minimum for the right side actuator lower cable. If this minimum distance is too short, then repeat Step (8). When the right side actuator upper cable terminal is positioned between 5.50 inches to 5.75 inches, then the lower cable must be 12.75 inches minimum. (Ref. Figure 206, Detail B). (10) With the right side actuator cables positioned per Step (8), check the position of the cable terminals ends of the left side elevator trim tab actuator. (a) The distance between the actuator housing (3) and the outboard end of the upper cable terminal end (6) must be 12.75 inches minimum for the left side actuator upper cable (5) (Ref. Figure 205, Detail B).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (b)

The distance between the actuator housing (3) and the outboard end of the lower cable terminal end (7) must be 5.50 inches minimum for the left side actuator lower cable (5) (Ref. Figure 205, Detail A).

(11) If no adjustment is required, proceed to Step (12). If adjustments are required, adjust the turnbuckles to meet the distance requirements of Step (10). (a) If the left actuator upper cable terminal end is less than 12.75 inches, then tighten the upper cable turnbuckle and loosen the lower turnbuckle. (b) If the left actuator lower cable terminal end is less than 5.50 inches, then loosen the upper cable turnbuckles and tighten the left side lower turnbuckle. (c) The bottom turnbuckle (5) inside the fuselage conduit tubes may need to be adjusted if sufficient adjustment is not available in the horizontal stabilizer turnbuckles to tension the cables and meet the minimum distance between the actuator housing and the cable terminal ends (Ref. Figure 208). (d) Perform Step (7) to tension the cables and then repeat Steps (8), (9), (10) and (11) to position the cables for the following Step (12). (12) The cables must be positioned per Steps (8), (9), (10) and (11). Check the position of the elevator trim tabs for neutral (0° deflection). If no adjustments are required, proceed to Step (13). If adjustment is required, perform the following Steps (a) Adjust the push-pull rod (10) by removing the safety wire and loosening the large jam nut (6) and remove the bolt, washer and nut (13) (Ref. Figure 202). (b) Loosen the small jam nut (8) and rotate the clevis (7) to bring the elevator tab to 0° position. (c) Tighten the small jam nut (8), taking care not to alter push-pull rod (10) adjustment. (d) Install bolt, washer and nut (13) attaching the trim tab to the push-pull rod (10). Tighten the large jam nut (6) to lock the outer clevis (7), bolt, washer and nut (13) in place and safety wire (178, Table 1, Chapter 91-00-00). (13) Position the elevator trim wheel to the full nose up limit. Verify the right side cable stop (1) contacts the cable stop guard (4) (Ref. Figure 204). (14) Using the elevator trim tab travel board or digital protractor (5, Table 1, 27-00-00), check the elevator trim tab for a deflection of 15° to 16° down from neutral (Ref. 27-00-02, READING A TRAVEL BOARD). (15) Position the elevator trim wheel to the full nose down limit. Verify the left side stop (1) contacts the cable stop guard (4). (16) Using the elevator trim tab travel board or digital protractor (5, Table 1, 27-00-00), check the elevator trim tab for a deflection of 5° to 5 1/2° up from neutral. (17) If no adjustment is required, proceed to Step (18). If adjustments are required, perform the following Steps (a) Remove safety wire (3) from the cable stop(s) (1) located in the horizontal stabilizer area (Ref. Figure 204).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (b) Loosen the cable stop(s) (1) and adjust as needed to achieve proper deflection. (c) Tighten cable stop(s) (1) and torque to 40 to 45 inch-pounds and safety wire (3). (d) Perform Steps (13), (14), (15) and (16) to check travel. (18) Rotate the elevator trim wheel on the pedestal so the trim tab trailing edge is positioned at neutral (0° deflection). (19) If using elevator travel board at HSS 35.00 with the elevator trim travel board attached, perform the following: (a) Remove rig pin from the elevator aft bellcrank (Ref. Figure 203). (b) Disconnect the trim tab and align the trim tab with the elevator. (c) Move the elevator surface to the full down position. Measurement should be 14° +1°/ -0° deflection. (d) Connect the trim tab and check measurement on travel board. Subtract the reading of the tab measurement from the elevator measurement. The trim tab must be 2° ± 1/2° (elevator trim tab deflection (servo travel)) up at full down elevator. Example: Elevator surface with trim tab disconnected and aligned with the elevator surface. Elevator deflection reads 14°. Elevator trim tab connected with the elevator in full down position. Tab deflection reads 12°. 14° - 12° = 2° which is acceptable. (e) Disconnect the trim tab and align the trim tab with the elevator. WARNING: Verify elevator movement in the following Step by moving only the cockpit control column. (f) Move the elevator surface to the full up position. Measurement should be 20° +1°/ -0° deflection. (g) Connect the trim tab and check measurement on travel board. Subtract the reading of the tab measurement from the elevator measurement. The trim tab must be 6°± 1° (elevator trim tab deflection (servo travel)) down at full up elevator. Example: Elevator surface with trim tab disconnected and aligned with the elevator surface. Elevator deflection reads 21°. Elevator trim tab connected with the elevator in full up position. Tab deflection reads 15°. 21° - 15° = 6° which is acceptable. NOTE: Maximum allowable trim tab differential (lagging tab) between the left tab and the right tab is to be 1° at full up elevator and 1/2° at full down elevator. (h) Proceed to Step (21). (20) If using elevator travel board at HSS 50.00 or HSS 35.00 with a digital protractor, perform the following: (a) With elevator surface at 0° (neutral) position, attach digital protractor (5, Table 1, 27-00-00) to the elevator trim tab surface.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (b) Zero out the digital protractor by pressing the ALT 0 button. (c) Remove rig pin from the elevator aft bellcrank (Ref. Figure 203). (d) Manually lower the elevator surface to the full down position. Elevator trailing edge down must be 14° +1°/ -0°. (e) Read the digital protractor. Subtract the reading of the digital protractor from the elevator measurement. The trim tab must be 2° ± 1/2° (elevator trim tab deflection (servo travel)) up at full down elevator. Example: Elevator deflection reads 14°. Digital protractor reads 12°. 14° - 12° = 2° which is acceptable. WARNING: Verify elevator movement in the following Step by moving only the cockpit control column. (f) Manually move the elevator surface to the full up position (20° +1°/ -0°). (g) Read the digital protractor. Subtract the reading of the digital protractor from the elevator measurement. The trim tab must be 6° ± 1° (elevator trim tab deflection (servo travel)) down at full up elevator. Example: Elevator deflection reads 21°. Digital protractor reads 15°. 21° - 15° = 6° which is acceptable. NOTE: Maximum allowable trim tab differential (lagging tab) between the left trim tab and the right trim tab is 1° at full up elevator and 1/2° at full down elevator. (h) Proceed to Step (21). (21) If no adjustment is required, proceed to Step (22). If adjustments are required, perform the following Steps: (a) A maximum of one laminated shim (15) (P/N 101-524078-3) may be placed between the adapter (14) and the actuator flange (Ref. Figure 202). (b) Adjustment of the actuator is accomplished by removing laminations from the shim as required to obtain the correct servo tab travel. NOTE: Shims shall not be under both the top and bottom of the actuator flange at the same time. Add or remove shims either at the top or bottom of the actuator flange. (c) To increase servo (lag), install shims under the actuator bottom flange only. No shims under the top flange. (d) To decrease servo (lag), install shims under the actuator top flange only. No shims under the bottom flange. (e) Bond shim (15) to the elevator trim tab actuator (1) using Epibond 104 (170, Table 1, Chapter 91-00-00). (f) If further adjustments are required, perform the ELEVATOR TRIM ACTUATOR REMOVAL/ INSTALLATION procedures (Ref. 27-30-06) and adjust actuators.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (22) Repeat Steps (5), (6), (8) and (12) through (21) for the opposite side. If both sides were rigged simultaneously, then these Steps have been accomplished; proceed to the next Step. (23) Move the elevator surface (1) and install a rig pin (3) (7, Table 1, 27-00-00) in the elevator aft bellcrank through the vertical stabilizer (4). Using minimum force try to manually move the elevators up and down to verify that the rig pin is correctly installed. Ensure that the elevators do not move (Ref. Figure 203). (24) Rotate the elevator trim wheel on the pedestal so the trim tab trailing edge is positioned at neutral (0° deflection). The 0 mark on the trim control wheel indicator must align with the triangle mark on the pedestal edgelighted panel. If required, perform the ELEVATOR TRIM TAB INDICATOR ADJUSTMENT procedure (Ref. 27-30-08). (25) Remove the elevator aft bellcrank rig pin. NOTE: Before attaching cable clamps and cables to the elevator trim cables, ensure that all twist is removed from the elevator trim cables by operating the system at least six (6) times from stop to stop. (26) Perform the ELEVATOR ELECTRIC TRIM TAB CABLE CONNECTION procedure, if installed (Ref. 27-30-07). Do not perform the ELEVATOR ELECTRIC TRIM OPERATIONAL CHECK procedure, since this will be accomplished in Step (28) of this procedure. (27) If installed, connect the Supplemental Type Certificate (STC) Flight Data Recorder (FDR) trim sensor bridle clamp. Refer to the STC holders instructions. (28) Perform the ELEVATOR ELECTRIC TRIM OPERATIONAL CHECK procedure, if installed (Ref. 27-30-03). (29) Perform the ELEVATOR TRIM TAB OPERATIONAL CHECK procedure contained in this section. (30) Remove the elevator travel board, elevator trim travel board, digital protractor and tape. (31) Install the left and right access panels 18, 21 and 22 (Ref. Chapter 6-50-00, STABILIZER ACCESS PANELS). (32) Install access panel on the left and right side of the upper vertical stabilizer 34 (Ref. Chapter 6-50-00, STABILIZER ACCESS PANELS).

B. Operational Check (1) Rotate the elevator trim control wheel (1) counter clockwise (nose down) and make sure that the elevator trim tab moves up smoothly with no unusual noise or binding (Ref. Figure 207). (2) Rotate the elevator trim control wheel (1) clockwise (nose up) and make sure that the elevator trim tab moves down smoothly with no unusual noise or binding. (3) Set the elevator trim control wheel (1) one and a half units in the nose up position. (4) With the elevators resting on the down stops, verify that the elevator trim tabs approximately align with the elevator surface. (5) If requirements are not met, perform the ELEVATOR TRIM TAB RIGGING procedure in this section.

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C. Cable Tension Check (1) Remove aft fuselage panels 7 and 8 (Ref. Chapter 6-50-00, FUSELAGE ACCESS PANELS). (2) Rotate the elevator tab control wheel (1) to the nose up and nose down position three cycles to equalize system tension (Ref. Figure 207). (3) Check the cable tension by performing the following Steps: NOTE: Do not measure the tension when the temperature is varying rapidly or with the airplane located in direct sunlight. (a) Allow the temperature of the airframe to stabilize before measuring and adjusting the cable tension. (b) Measure the temperature in the compartment next to the elevator trim cables. (c) Refer to the Elevator Trim Cable Tension Graph Figure 201 and read the pounds of tension for the measured temperature. (d) Position a cable tensiometer (4, Table 1, 27-00-00) at least three inches from the turnbuckles and measure the cable tension. Cable diameter is noted in Figure 201. NOTE: Cable tension tolerance is +3/ -2 pounds. (e) If cable tension does not require adjustment, proceed to Step (4). If cable tension requires adjustment, perform the ELEVATOR TRIM TAB RIGGING procedure in this section. (4) Install panel 7 and 8 (Ref. Chapter 6-50-00, FUSELAGE ACCESS PANELS). (5) Perform the ELEVATOR TRIM TAB OPERATIONAL CHECK procedure in this section.

D. Functional Check (1) Remove access panel 34 on the left and right side of the upper vertical stabilizer (Ref. Chapter 6-50-00, STABILIZER ACCESS PANELS). (2) Move the elevator surface (1) and install a rig pin (3) (7, Table 1, 27-00-00) in the elevator aft bellcrank through the vertical stabilizer (4). Look at the opposite side of the vertical stabilizer to verify that the rig pin has protruded. Using minimum force try to manually move the elevators up and down to verify that the rig pin is correctly installed and that the elevators will not move (Ref. Figure 203). (3) Perform the ELEVATOR TRAVEL BOARD INSTALLATION procedure (Ref. 27-00-02). Verify that the elevator is at 0°. If the elevator is not at 0°, perform the ELEVATOR CONTROL SYSTEM RIGGING procedure (Ref. 27-30-02). (4) Perform the ELEVATOR TAB TRAVEL BOARD INSTALLATION procedure (Ref. 27-00-02). (5) Rotate the elevator trim control wheel (1) to align the trailing edge of the trim tab with the elevator surface (Ref. Figure 207). (6) The 0 mark on trim position dial (6) must be aligned with the trim indicator mark (3) on the pedestal edgelighted panel (4).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL CAUTION: The elevator trim system must not be forced past the limits that are indicated on the elevator trim position dial (6) by a red mark (2). (7) Rotate the elevator tab control wheel (1) counter clockwise to the full nose down position and make sure that the tab moves to the full up position 5 to 5 1/2° smoothly with no unusual noise or binding. (8) Rotate the elevator tab control wheel (1) clockwise to the full nose up position and make sure that the tab moves to the full down position 15° to 16° smoothly with no unusual noise or binding. (9) If elevator trim tab requires adjustment, perform the ELEVATOR TRIM TAB RIGGING in this section. (10) Remove the elevator tab travel boards. (11) Remove the elevator travel board. (12) Remove rig pin (3) from the elevator aft bellcrank (6) (Ref. Figure 203). (13) Install access panel 34 on the left and right side of the upper vertical stabilizer (Ref. Chapter 6-50-00, STABILIZER ACCESS PANELS).

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Figure 201 Elevator Trim Tab Cable Tension Graph

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A 1. 2. 3. 4. 5. 6. 7. 8. 9. 10. 11. 12. 13. 14. 15.

ELEVATOR TRIM TAB ACTUATOR ACTUATOR HOUSING ACTUATOR ROD WASHER ROD END JAM NUT LARGE JAM NUT OUTER CLEVIS SMALL JAM NUT ROD ENDS PUSH-PULL ROD BOLT, WASHER AND NUT INSPECTION HOLE BOLT, WASHER AND NUT ADAPTER SHIM (LAMINATED)

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Figure 202 Elevator Trim Tab Actuator Rod Adjustment

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1. ELEVATOR 2. RUDDER 3. ELEVATOR AFT BELLCRANK RIG PIN 4. VERTICAL STABILIZER 5. HORIZONTAL STABILIZER

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Figure 203 (Sheet 1 of 2) Elevator Aft Bellcrank Rig Pin Installation

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6. ELEVATOR AFT BELLCRANK 7. ELEVATOR AFT RIG PIN HOLE

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VERTICAL STABILIZER (REF)

DETAIL

B UC27B 040558AB.AI

Figure 203 (Sheet 2 of 2) Elevator Aft Bellcrank Rig Pin Installation

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A 1. CABLE STOP 2. TRIM TAB CABLE 3. SAFETY WIRE 4. CABLE STOP GUARD

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DETAIL

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Figure 204 Elevator Trim Tab Cable Stop

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B A

1. HORIZONTAL STABILIZER 2. ELEVATOR TRIM TAB ACTUATOR 3. ELEVATOR TRIM TAB ACTUATOR HOUSING 4. TAPE MEASURE 5. ELEVATOR TRIM TAB CABLE 6. TERMINAL END 7. TERMINAL END INBD AFT

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Figure 205 Left Elevator Trim Tab Actuator Cable Check

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AB

1. HORIZONTAL STABILIZER 2. ELEVATOR TRIM TAB ACTUATOR 3. ELEVATOR TRIM TAB ACTUATOR HOUSING 4. TAPE MEASURE 5. ELEVATOR TRIM TAB CABLE 6. TERMINAL END 7. TERMINAL END

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Figure 206 Right Elevator Trim Tab Actuator Cable Check

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1. TRIM CONTROL WHEEL 2. RED MARK 3. TRIM INDICATOR MARK 4. EDGELIGHTED PANEL 5. PEDESTAL 6. TRIM POSITION DIAL

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NOTE: RED MARK SHOWN IS FOR REFERENCE ONLY. LOCATION MAY VARY BETWEEN AIRPLANES.

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Figure 207 Elevator Trim Tab Indicator Adjustment

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

1. TIE STRAPS 2. CONDUIT TUBES 3. BRACKET 4. CABLE BLOCK 5. TURNBUCKLE 6. RIGHT - HAND THREADS TERMINAL END

7. TURNBUCKLE 8. LEFT - HAND THREADS TERMINAL END 9. BRACKET 10. FORWARD CABLE 11. LEFT AFT CABLE 12. RIGHT AFT CABLE

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Figure 208 Aft Trim Tab Cable

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FLIGHT CONTROLS 27-30-06

ELEVATOR TRIM TAB ACTUATORS MAINTENANCE PRACTICES

200200

1. PROCEDURES A. Removal (1) Remove access panels 15, 18, 21 and 22 (Ref. Chapter 6-50-00, STABILIZER ACCESS PANELS). (2) Remove both elevators (Ref. 27-30-00, ELEVATOR REMOVAL). (3) Remove screws from access plate around trim actuator housing and remove the access plate. (4) Move the right side actuator upper cable (12) until the distance from the actuator housing (10) to the terminal end is 5.50 to 5.75 inches (Ref. Figure 204). (5) Install cable block (2) on the elevator trim actuator cables (Ref. Figure 201). (6) Identify, tag and disconnect the turnbuckle left-hand threads terminal end (5) and the turnbuckle right-hand threads terminal end (7) on the outboard side of each turnbuckle (4 and 8) (Ref. Figure 202). (7) Loosen the large jamnut (11) on clevis and remove nuts, washers and bolts (12) that attach the left and right push-pull rods (10) to the rod ends (6). Remove the push-pull rods (10) (Ref. Figure 203). (8) Remove the mounting bolts and remove the elevator trim actuator. NOTE: Note the locations of shims, if installed, and lengths of the actuator mounting bolts during actuator removal to ensure proper reassembly.

B. Cable Replacement NOTE: If a used cable is installed, the cable should be dipped in corrosion preventive compound (11, Table 1, Chapter 91-00-00). Excess material should be removed by wiping with a clean cloth. (1) Remove the trim tab actuator in accordance with ELEVATOR TRIM TAB ACTUATOR REMOVAL. (2) Pull the actuator cables until the “up “cable and the “down” cable hanging from the drum are of equal length. (3) Remove the large jam nut and key from the bronze shaft. (4) Remove safety wire from the three drum housing bolts and remove the bolts. (5) Remove the end cover off of the trim actuator retaining the cable drum. (6) With a pencil, mark the position of the drum against the housing and mark the position of the shafts in the end of the screw housing. (7) Lift the drum out of the housing and unwind the cable.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (8) Remove the cable retaining sleeve from the slot and remove the cable. (9) Install the new cable in position and install the cable retaining sleeve. (10) Wind the cable on to the cable drum. (11) Install the drum in the housing as previously assembled, check pencil marks for alignment, reinstall and secure the housing bolts and jam nut (DO NOT OVERTORQUE THE JAM NUT). (12) Install the trim actuator in accordance with ELEVATOR TRIM TAB ACTUATOR INSTALLATION. (13) Install safety clips on the turnbuckles.

C. Installation (1) Prior to actuator installation, with the actuator on bench, align the cable ends together. Check measurement from end of bushing (13) to the end of the rod (3). Measurement should be 1.31 ± 0.05 inches (Ref. Figure 203). (2) If no adjustment is required, proceed to Step (3). If adjustment is required, perform the following Steps: (a) Remove nut, washer and bolt (17) from rod (3). (b) Rotate rod (3) as necessary to obtain measurement. (c) Install bolt, washer and nut (17) in rod (3). NOTE: If necessary, move cables (15 and 16) to allow bolt (17) installation into the rod (3). (3) Align the cable ends together and check measurement from the actuator housing (2) to the center of the rod end (6). Measurement should be 2.44 to 2.50 inches. The rod end (6) threads must be visible through the actuator rod (3) inspection holes (14). (4) If no adjustment is required, proceed to Step (5). If adjustment is required, perform the following Steps: (a) Remove safety wire from the rod end jamnut (5) to the tab washer (4). (b) Loosen jamnut (5) on rod end (6). (c) Adjust rod end (6) until measurement is obtained. (d) Tighten the rod end jamnut (5). (5) Position the actuator in place with the cable ends inboard. Install shims (if installed) between the actuator housing and adapter. Install bolts, washers and nuts. Check for proper bolt length at each corner of the tab actuator. (6) Identify and connect the turnbuckles by moving the actuator cables as necessary. Tension cables to remove cable slack. (7) Remove cable blocks from the actuator, if installed, and from the trim cables.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (8) Move the right side actuator upper cable (12) until the distance from the actuator housing (10) to the terminal end is 5.50 to 5.75 inches (Ref. Figure 204). NOTE: The left side actuator lower cable may not be at the same dimension as the right side upper cable. (9) Check measurement on both actuators from the actuator housing (10) to the center of the rod end (5). Measurement should be 2.69 to 2.75 inches. The rod end (5) threads must be visible through the actuator rod (8) inspection holes (2) (Ref. Figure 204). (10) If no adjustment is required, proceed to Step (11). If adjustment is required, perform the following Steps: (a) Remove safety wire from the rod end jamnut (6) to the tab washer (7). (b) Loosen jamnut (6) on rod end (5). (c) Adjust rod end (5) until measurement is obtained. (d) Tighten the rod end jamnut (6). (e) Safety wire the jamnut (6) to the tab washer (7) and recheck measurement. (11) Check the length of the push-pull rod (10) from the center of the forward clevis to the center of the aft clevis. Measurement should be 10.71 ± 0.06 inches (Ref. Figure 203). (12) If no adjustment is required, proceed to Step (13). If adjustment is required, perform the following Steps: (a) Loosen small jamnut (9) on aft clevis (7). (b) Adjust aft clevis (7) until measurement is obtained. (c) Adjust the small jamnut (9) on aft clevis (7) until the small jamnut makes contact to the aft clevis to preset rod length. (d) Remove aft clevis (7) from rod (10). (13) Install bolts, washers and nuts (12) that attach the left and right push-pull rods (10) to the rod ends (6). Tighten the forward clevis jamnut (11) on forward clevis. Make sure that the head of the bolt (12) is on the outboard side of the rod end (6). (14) Perform the ELEVATOR INSTALLATION procedure (Ref. 27-30-00). (15) Install aft clevis (7) to rod (10) and connect to the trim tab horn. (16) Perform the ELEVATOR TRIM TAB RIGGING procedure (Ref. 27-30-05). (17) Install access panels 15, 18, 21 and 22 (Ref. Chapter 6-50-00, STABILIZER ACCESS PANELS). NOTE: It is permissible to use Sealant Tape (19, Table 2, 27-00-00) as a moisture seal when installing these panels.

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1. ELEVATOR TRIM ACTUATOR 2. CABLE BLOCK 3. ACTUATOR BOTTOM CABLE 4. ACTUATOR TOP CABLE

RH ACTUATOR DETAIL

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Figure 201 Elevator Trim Tab Actuator

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

A 1. HORIZONTAL STABILIZER 2. ELEVATOR TRIM ACTUATOR 3. LOWER TRIM CABLE 4. TURNBUCKLE 5. LEFT-HAND THREADED CABLE END 6. UPPER TRIM CABLE 7. RIGHT-HAND THREADED CABLE END 8. TURNBUCKLE

FWD INBD

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A UC27B 041226AA.AI

Figure 202 Elevator Trim Tab Actuator

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1. ELEVATOR TRIM TAB ACTUATOR 2. ACTUATOR HOUSING 3. ACTUATOR ROD 4. TAB WASHER 5. ROD END JAMNUT 6. ROD END 7. OUTER CLEVIS 8. AFT CLEVIS JAMNUT 9. SMALL JAMNUT 10. PUSH-PULL ROD 11. FORWARD CLEVIS JAMNUT 12. BOLT, WASHER AND NUT 13. BUSHING 14. INSPECTION HOLE 15. BOTTOM CABLE 16. TOP CABLE 17. BOLT, WASHER AND NUT

1

A A

B

WITH CABLE ENDS TOGETHER 2 4 5 6

3

14

13 2.44 TO 2.50 INCHES

DETAIL 15

12

11

A

10 9

8

7

16

1

DETAIL

B

17

Figure 203 Elevator Trim Tab Actuator Adjustments

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1. 2. 3. 4. 5. 6. 7. 8. 9. 10. 11. 12.

A

BOLT, WASHER AND NUT INSPECTION HOLE BUSHING BOLT, WASHER AND NUT ROD END ROD END JAMNUT TAB WASHER ACTUATOR ROD ACTUATOR HOUSING ELEVATOR TRIM TAB ACTUATOR TOP CABLE BOTTOM CABLE

A

RIG DIMENSIONS AT NEUTRAL TAB

2.69 TO 2.75 INCHES 1

4

3 2

8

7 6

5

9 10

5.50 to 5.75 INCHES

11

12

DETAIL

A

UC27B 041850AA.AI

Figure 204 RH Elevator Trim Tab Actuator Dimensions

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27-30-07 200200

FLIGHT CONTROLS ELEVATOR ELECTRIC TRIM TAB SYSTEM MAINTENANCE PRACTICES 1. ELEVATOR ELECTRIC TRIM TAB A. Cable Disconnection (1) Attach a red tag to the elevator electric trim switch with the words “Do Not Operate, Maintenance In Progress”. (2) Open the ELEV TRIM circuit breaker on the circuit breaker panel. (3) Remove the fuselage access panel 7 and 8 (Ref. Chapter 6-50-00, FUSELAGE ACCESS PANELS). (4) Remove the left and right elevator electric trim servo covers. (5) Install a cable block on the left and right cable drums so the cable will not unwind. (6) Remove safety clips from the right electric trim cable turnbuckle (1) and disconnect the right elevator electric trim cable (14) from the right elevator trim system cable (9) (Ref. Figure 202). (7) Remove safety clips from the left electric trim cable turnbuckle (19) and disconnect the left elevator electric trim cable (4) from the left elevator trim system cable (5).

B. Actuator Removal (1) Attach a red tag to the elevator electric trim switch with the words “Do Not Operate, Maintenance In Progress”. (2) Open the ELEV TRIM circuit breaker on the circuit breaker panel. (3) Remove fuselage access panels 7 and 8 (Ref. Chapter 6-50-00, FUSELAGE ACCESS PANELS). (4) Remove the left and right elevator electric trim servo covers. (5) Disconnect the electrical connector (1) from the elevator electric trim control box (2) (Ref. Figure 201). (6) Install a cable block on the left and right cable drums so the cable will not unwind. (7) Remove safety clips from the right servo cable turnbuckle (1) and disconnect the servo cable (14) from the elevator trim cable (9) at the servo cable turnbuckle (1) (Ref. Figure 202). (8) Remove three mounting screws (4) and two nuts (the upper mounting screw has a nutplate) from the elevator electric trim actuator (3) (Ref. Figure 201). (9) Remove the elevator electric trim actuator (3) and control box (2) with the mounting bracket (5) from the airplane by pulling the actuator outboard disengaging the universal joint.

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C. Actuator Installation (1) Position the elevator electric trim actuator (3), control box (2) and mounting bracket (5) (Ref. Figure 201) in place and unwind the right elevator electric servo cable (14) from the right servo cable drum (12) to align the right elevator electric servo cable (14) to the right elevator trim cable (9) at the cable turnbuckle (1) (Ref. Figure 202). (2) Align the elevator electric trim actuator (3), control box (2) and mounting bracket (5) in place ensuring the cable drum shaft is engaged to the universal joint and install three mounting screws (4) and two nuts (the upper mounting screw has a nutplate) (Ref. Figure 201). (3) Lubricate turnbuckle (1) with grease (1, Table 2, 27-00-00) for corrosion protection prior to installation (Ref. Figure 202). (4) Connect the right elevator electric servo cable (14) to the right elevator trim system cable (9) at the cable turnbuckle (1) and tension right elevator trim cable to remove slack. (5) Remove the cable blocks from the left and right elevator electric trim servo cable drums. (6) Connect the electrical connector (1) to the elevator electric trim control box (2) (Ref. Figure 201). (7) Close the ELEV TRIM circuit breaker on the circuit breaker panel. (8) Remove the red tag from the elevator electric trim switch. (9) Perform the ELEVATOR ELECTRIC TRIM TAB CABLE RIGGING procedure in this section. (10) Install the left and right elevator electric trim servo covers. (11) Install fuselage access panels 7 and 8 (Ref. Chapter 6-50-00, FUSELAGE ACCESS PANELS).

D. Cable Connection (1) Lubricate turnbuckles (1 and 19) with grease (1, Table 2, 27-00-00) for corrosion protection prior to installation (Ref. Figure 202). (2) Connect the right elevator electric trim cable (14) to the right elevator trim system cable (9) at turnbuckle (1) and tension the right elevator trim cable to remove slack. (3) Connect the left elevator electric trim cable (4) to the left elevator trim system cable (5) at cable turnbuckle (19) and tension the left elevator trim cable to remove the slack. (4) Remove the cable blocks from the left and right elevator electric trim servo cable drums. (5) Install the left and right elevator electric trim servo covers. (6) Close the ELEV TRIM circuit breaker on the circuit breaker panel. (7) Remove the red tag from the elevator electric trim switch. (8) Perform the ELEVATOR ELECTRIC TRIM TAB CABLE RIGGING procedure contained in this section. (9) Install the left and right elevator electric trim cable turnbuckle safety clips.

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2. LEFT CABLE DRUM AND CABLE A. Removal (1) Perform the ELEVATOR ELECTRIC TRIM TAB ACTUATOR REMOVAL procedure in this section. (2) Using the manual elevator trim wheel, set the elevator trim to 0°. (3) Install cable block to the right side elevator electric servo cable (14) at the right cable drum (12) (Ref. Figure 202). (4) Remove safety clips from the left elevator electric trim cable turnbuckle (19). (5) Disconnect the elevator left electric trim cable (4) terminal end from the turnbuckle (19). (6) Remove safety wire (10) from the left side of the universal joint (13) and remove the roll pin securing the left side of the universal joint (13) to the left cable drum shaft (7). (7) Remove screws, washers and nuts (6) from the elevator electric trim cable drum support (22). (8) Unwind the cable (4) from the cable drum (15) to expose the cable drum retaining pin. (9) Align the small hole in the cable drum shroud (8) with the cable drum retaining pin hole. CAUTION: While removing the cable drum retaining pin from the cable drum (15), do not damage the cable drum grooves. (10) Remove the cable drum retaining pin from the cable drum (15). (11) Remove the retaining ring (16) from the cable drum shaft (7). (12) Hold the cable drum (15) and tap on the inboard side of the shaft (7) toward the outboard side of the left cable drum support (22). The bearing (21) will come out with the cable drum shaft (7). Remove the cable drum shaft (7). (13) Remove the cable drum shroud (8), cable drum (15) and cable (4) from the left cable drum support (22). (14) Remove the cable (4) from the cable drum (15).

B. Installation (1) Position the cable ball in the cable drum (15) and slide the cable retaining sleeve in place (Ref. Figure 202). (2) Slide the cable drum shroud (8) over the cable drum (15) ensuring that the small hole in the shroud (8) and the retaining pin hole in the cable drum (15) are positioned outboard. (3) Position the cable drum shroud (8), cable drum (15) and cable (4) in the left cable drum support (22). (4) Slide the cable drum shaft (7) through the cable drum support (22) and cable drum (15) and press the cable drum bearing (21) into the left cable drum support (22). (5) Install the retaining ring (16) to the cable drum shaft (7).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL CAUTION: While installing the cable drum retaining pin into the cable drum (15), do not damage the cable drum grooves. (6) Align the cable drum shaft (7) hole with the cable drum (15) retaining pin hole and install the cable drum retaining pin in the cable drum (15). (7) Install screws, washers and nuts (6) securing the cable drum shroud (8) in the elevator electric trim cable drum support (22). (8) Wrap the cable (4) around the cable drum (15) fully ensuring the cable (4) wraps off the top of the cable drum (15). (9) Lubricate turnbuckle (19) with grease (1, Table 2, 27-00-00) for corrosion protection prior to installation. (10) Connect the left elevator electric trim cable (4) terminal end to the turnbuckle (19). (11) Install safety clips in the left elevator electric trim cable turnbuckle (19). (12) Align the cable drum shaft (7) with the universal joint (13). (13) Install the roll pin into the universal joint (13) and safety wire (10). (14) Perform the ELEVATOR ELECTRIC TRIM TAB ACTUATOR INSTALLATION procedure in this section.

3. ACTUATOR CABLE DRUM AND CABLE A. Removal (1) Perform the ELEVATOR ELECTRIC TRIM TAB ACTUATOR REMOVAL procedure in this section. (2) Remove screws, washers and nuts (17) from the elevator electric trim cable drum end plate (18) (Ref. Figure 202). (3) Remove the cable drum shroud (11) and the cable drum end plate (18) together from the trim actuator. (4) Unwind the cable (14) from the cable drum (12) to gain access to the cable drum retaining pin. CAUTION: While removing the cable drum retaining pin from the cable drum (12), do not damage the cable drum grooves. (5) Remove the cable drum retaining pin from the cable drum (12). (6) Slide the cable drum (12) and cable (14) off the trim actuator cable drum shaft. (7) Remove cable (14) from the cable drum (12).

B. Installation (1) Position the cable ball in the cable drum (12) and slide the retaining sleeve in place (Ref. Figure 202).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (2) Slide the cable drum (12) on the trim actuator cable drum shaft and align the cable drum retaining pin hole with the cable drum shaft hole. CAUTION: While installing the cable drum retaining pin to the cable drum (12), do not damage the cable drum grooves. (3) Install the cable drum retaining pin in the cable drum (12) ensuring that the retaining pin is recessed below the cable drum grooves. (4) Wrap the cable (14) around the cable drum (12) ensuring the cable (14) wraps off the bottom of the cable drum (12). (5) Position the cable drum shroud (11) and end plate (18) to the trim actuator. (6) Install screws, washers and nuts (17) securing the cable drum shroud (11) and end plate (18) to the trim actuator. (7) Perform the ELEVATOR ELECTRIC TRIM TAB ACTUATOR INSTALLATION procedure in this section.

4. SERVO UNIVERSAL JOINT A. Removal (1) Perform the ELEVATOR ELECTRIC TRIM TAB ACTUATOR REMOVAL procedure in this section. (2) Install cable block at the left cable drum. (3) Remove safety wire (10) from the left side of the universal joint (13) and remove the roll pin securing the left side of the universal joint (13) to the left cable drum shaft (7) (Ref. Figure 202). (4) Remove the universal joint (13) from the airplane.

B. Installation (1) Align the left side universal joint (13) hole with the left cable drum shaft (7) hole and install the roll pin (Ref. Figure 202). (2) Safety wire (10) the roll pin to the universal joint (13). (3) Remove cable block from the left cable drum. (4) Perform the ELEVATOR ELECTRIC TRIM TAB ACTUATOR INSTALLATION procedure in this section.

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1. ELECTRICAL CONNECTOR 2. ELECTRICAL CONTROL BOX 3. ELECTRIC TRIM ACTUATOR 4. MOUNTING SCREWS 5. BRACKET 6. R1 ADJUSTMENT SCREW

A

2 1

6

3

4

5

UC27B 044606AA.AI

Figure 201 Electric Elevator Trim Tab Servo Actuator Installation

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3

2

A

1

1. RIGHT ELECTRIC TRIM CABLE TURNBUCKLE 2. RIGHT ELEVATOR TRIM SYSTEM TURNBUCKLE 3. ELECTRIC TRIM CABLE BRIDLES 4. LEFT ELEVATOR ELECTRIC TRIM CABLE 5. LEFT ELEVATOR TRIM SYSTEM CABLE 6. SCREWS, WASHERS, AND NUTS 7. SHAFT 8. LEFT CABLE DRUM SHROUD 9. RIGHT ELEVATOR TRIM SYSTEM CABLE 10. SAFETY WIRE 11. RIGHT CABLE DRUM SHROUD 12. RIGHT CABLE DRUM 13. UNIVERSAL JOINT 15 14. RIGHT ELEVATOR ELECTRIC TRIM CABLE 14 15. LEFT CABLE DRUM 16. RETAINING RING 13 17. SCREWS, WASHERS, AND NUTS 18. END PLATE 17 19. LEFT ELECTRIC TRIM 18 CABLE TURNBUCKLE 12 20. LEFT ELEVATOR TRIM 11 SYSTEM TURNBUCKLES 21. CABLE DRUM BEARING 10 22. LEFT CABLE DRUM SUPPORT

19 20

14 9 4 5 16

4

5

9 8

22

6 21 7 6 DETAIL

A

UC27B 044607AA.AI

Figure 202 Electric Elevator Trim Tab Cable Installation

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Figure 203 Trim Tab Servo

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C

B

A

G F

D E

PILOT CONTROL WHEEL

N O S E D W N

U P 1

DETAIL

2

3

4

5

SECOND STEP TRIM

Y D D I S C S W

D W N

U P

COPILOT CONTROL WHEEL

DETAIL RIGHT TRIM SWITCH

DISCONNECT SWITCH

N O S E

N O S E

E

DISCONNECT SWITCH

RIGHT TRIM SWITCH

LEFT TRIM SWITCH N O S E

DETAIL

6

FIRST STEP AP & YD

FIRST STEP AP & YD

SECOND STEP TRIM

N O S E

15

22

14

13

14

9 13

9

22

15

6

LEFT TRIM SWITCH N O S E

N O S E

D W N

N O S E

D W N

U P 5

C

4

3

U P 2

1

F 15 14 10 18 11

DETAIL

12

A

ELECTRIC TRIM OFF SIGNAL (LEFT ADVISORY LIGHT TEST & TIME DELAY PCB ASSEMBLY)

16

DETAIL 13 17 RELAY BOARD - A120 (FORWARD LOWER CABIN ELECTRICAL PANEL)

6A H

F

G

E

A

K

C

B

J

D

5A OFFRESET 1

B

5A CIRCUIT BREAKER (CIRCUIT BREAKER PANEL)

2

TRIM MOTOR CLUTCH

3 ON

ELECTRIC TRIM DISCONNECT RELAY

DETAIL

D

ELECTRIC TRIM POWER SWITCH (PEDESTAL PANEL)

M

DETAIL

G

TRIM MOTOR

TRIM MOTOR SPEED CONTROL

TRIM MOTOR CLUTCH POWER CONTROL UC27B 031266AB.AI

TRIM TAB ELECTRIC CONTROL ASSEMBLY

Figure 204 Electric Elevator Trim Tab Wiring Diagram

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5. MAGNETIC CLUTCH A. Removal (1) Remove the lid from the clutch housing (Ref. Figure 203). (2) Loosen the setscrew in the clutch rotor and armature hubs. (3) Remove the motor from the clutch housing. (4) Slide the cable drum and shaft assembly from the clutch housing. (5) Remove the clutch from the clutch housing.

B. Installation (1) Install the clutch in the clutch housing (Ref. Figure 203). (2) Slide the cable drum and shaft assembly into the clutch housing. (3) Tighten the clutch armature setscrews with no visible end play in the cable drum shaft. Slide the clutch rotor on the motor shaft to obtain 0.010 to 0.015 inch clearance between the friction surfaces of the clutch before tightening the setscrews. Stake both setscrews. (4) Install the motor on the clutch housing with the attaching screws. (5) Install the lid on the clutch housing.

C. Torque Test The following check should be performed any time the magnetic clutch is replaced: (1) With a 28 vdc power source, actuate the magnetic clutch. Connect the red lead to ground and the white lead to 28 vdc. With a torque wrench, check to see that the clutch will hold 30 inch-pounds applied at the actuator shaft. (2) If the static torque of the clutch is less than 30 inch-pounds, burn in the clutch as follows: (a) Bolt the actuator to a plate and tighten the plate in a vise. (b) Fabricate a rod that will fit the hole in the end of the drum shaft. Grind the rod so that it has a flat end (like a screwdriver) to fit into the slot in the bottom of the hole. (c) Fit the other end of the rod to a low speed motor (450 rpm) for a source of power to burn in the clutch. (d) Using 14 to 16 vdc power, activate the magnetic clutch and run the motor for 15 seconds. Allow to cool for one minute and check torque. When 30 inch-pounds of torque as checked in Step a. is achieved, reassemble the unit and place it back in service. CAUTION: Exceeding the 15 second burn-in periods may overheat and damage the magnetic clutch. Always allow to cool and blow clean with compressed air to prevent damage.

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6. ELEVATOR ELECTRIC TRIM TAB A. Cable Rigging (All Airplanes Without Collins APS-65H Autopilot System) NOTE: Make sure that the elevator trim system has been rigged prior to performing this procedure. (1) Remove fuselage access panels 7 and 8 (Ref. Chapter 6-50-00, FUSELAGE ACCESS PANELS). (2) Rotate the elevator trim wheel to the full up travel. Verify that the turnbuckles (1 and 19) clear the cable drums (12 and 15) and the bridle clamp (3) does not enter the vertical stabilizer area. Ensure that the elevator electric servo cables (4 and 14) are not twisted around the elevator trim primary cables (5 and 9). Adjust the trim cable bridle clamps (3) as necessary (Ref. Figure 202). (3) Rotate the elevator trim wheel to the full down travel. Verify that the turnbuckles (1 and 19) clear the cable drums (12 and 15) and the bridle clamp (3) does not enter the vertical stabilizer area. Ensure that the elevator electric servo cables (4 and 14) are not twisted around the elevator trim primary cables (5 and 9). Adjust the trim cable bridle clamps (3) as necessary. (4) Verify that the bridle clamp (3) screws are installed with the screw heads on the right side of the primary elevator trim cables (5 and 9). (5) Position a cable tensiometer (4, Table 1, 27-00-00) at least three inches from the turnbuckles (1 and 19) and measure the cable tension of both elevator electric trim cables (4 and 14). Cable tension must be 9 +3/ -2 pounds. Adjust turnbuckles (1 and 19) as required. (6) Perform the ELEVATOR ELECTRIC TRIM OPERATIONAL CHECK procedure (Ref. 27-30-03). (7) Install fuselage access panels 7 and 8 (Ref. Chapter 6-50-00, FUSELAGE ACCESS PANELS).

B. Actuator Speed Adjustment (All Airplanes Without Collins APS-65H Autopilot System) Adjustment of servo speed must be performed any time a complete servo assembly is replaced. The elevator trim system and the electric elevator trim system must be fully installed, rigged and tensioned prior to performing the following Steps: (1) Remove fuselage access panels 7 and 8 (Ref. Chapter 6-50-00, FUSELAGE ACCESS PANELS). (2) Connect a 28 vdc regulated power supply to the airplane. (3) Turn the battery switch ON. (4) Turn the external power switch ON. (5) Attach a piece of tape to either elevator system trim cable and locate a scale next to the tape.

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(6) Starting from just short of the tab full-up position (nose-down), operate the electric elevator trim in the nose-up direction. The time required for the cable to travel 10 inches should be between 25.1 and 26.6 seconds. If not, adjust R1 (6) on the electrical control box (2) to produce the required travel rate. If the required travel rate cannot be produced, check for magnetic clutch slippage or binding in the cable system (Ref. Figure 201). (7) Remove the scale and tape. (8) Install fuselage access panels 7 and 8 (Ref. Chapter 6-50-00, FUSELAGE ACCESS PANELS).

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FLIGHT CONTROLS ELEVATOR TRIM TAB INDICATOR MAINTENANCE PRACTICES

27-30-08 200200

1. PROCEDURES A. Removal (1) Place the elevator trim control wheel so that the 0 mark on trim position dial (6) aligns with the trim indicator mark (3) on the pedestal edgelighted panel (4) (Ref. Figure 203, Detail A). (2) Remove screw (1) and washers (8) from the trim control wheel (3). Pull the trim control wheel (3) outboard and remove from pedestal (Ref. Figure 202). NOTE: The position dial (6) and the mount (9) are bonded together as an assembly. (3) Rotate the position dial (6) to align holes with the screws (10). Remove three screws (10) from the mount (9). (4) Remove the position dial (6) and mount (9) from the pedestal.

B. Installation NOTE: The position dial (6) and the mount (9) are bonded together as an assembly (Ref. Figure 202). (1) Align the position dial (6) and mount (9) to the pedestal and install three screws (10). (2) Align the 0 degree position on the position dial (6) with middle triangular trim indicator mark (3) on the pedestal edgelighted panel (4) (Ref. Figure 203). (3) Align hub (2) and push the control wheel (3) inboard and rotate to engage the hub (2) to the chain sprocket (7) (Ref. Figure 202). (4) Engage the control wheel gear (4) to the reduction gear (5) and install screw (1) and washers (8). (5) Perform the ELEVATOR TRIM TAB INDICATOR ADJUSTMENT procedure in this section.

C. Inspection (1) Remove screw (1) and washers (8) from the trim control wheel (3). Pull the trim control wheel (3) outboard and remove from pedestal (Ref. Figure 202). (2) Inspect the control wheel gear (4) and reduction gear (5) for distortion, wear and missing teeth. (3) Align hub (2) and push the control wheel (3) inboard and rotate to engage the hub (2) to the chain sprocket (7). (4) Engage the control wheel gear (4) to the reduction gear (5) and install screw (1) and washers (8). (5) Perform the ELEVATOR TRIM TAB INDICATOR ADJUSTMENT procedure in this section.

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D. Adjustment (1) Remove the lower aft rig pin access panel 34 on the left and right side of the upper vertical stabilizer (Ref. Chapter 6-50-00, STABILIZER ACCESS PANELS). (2) Move the elevator surface to neutral position (0°) and install rig pin (7, Table 1, 27-00-00) in the aft elevator bellcrank through the vertical stabilizer (4) Look at the opposite side of the vertical stabilizer to verify that the rig pin (3) has protruded. Using minimum force, try to manually move the elevators (1) to verify that the rig pin is correctly installed and that the elevators do not move (Ref. Figure 201). (3) Rotate the elevator trim wheel to align the trailing edge of the trim tab (6) with the trailing edge of the elevator surface (1). (4) The 0 mark on trim position dial (6) must be aligned with the trim indicator mark (3) on the pedestal edge lighted panel (4). If the 0 mark on the trim position dial (6) is not aligned with the trim indicator mark (3) on the pedestal edge lighted panel, proceed to Step (4) (a). If the 0 mark on the trim position dial (6) is aligned with the trim indicator mark (3) on the pedestal edge lighted panel, proceed to Step (5). (Ref. Figure 203, Detail A). (a) Remove screw (1) from the trim control wheel (3). Pull the trim control wheel outboard, until the control wheel gear (4) is disengaged from the reduction gear (5) (Ref. Figure 202). NOTE: The control wheel gear (4) is bonded to the hub inboard of the trim control wheel (3). (b) With the reduction gear (5) (Ref. Figure 202) disengaged, rotate the position dial (6) aligning the 0 degree position on the position dial with middle triangular trim indicator mark (3) on the pedestal edge lighted panel (4) (Ref. Figure 203). (c) Push the control wheel inboard and rotate to engage the hub (2) to the chain sprocket (7) (Ref. Figure 202). (d) Engage the control wheel gear (4) to the reduction gear (5) and install screw (1). (5) Rotate the trim wheel (1) to the full nose up limit. Verify the elevator trim tab trailing edge is down. Place a red mark (2) 0.06 inches wide and 0.30 inches long on the position dial (6) opposite the trim indicator mark (3) on the pedestal edge lighted panel (Ref. Figure 203, Detail B). (6) Rotate the trim wheel (1) to the full nose down limit (Ref. Detail C). Verify the elevator trim tab trailing edge is up. Place a red mark (2) 0.06 inches wide and 0.30 inches long on the position dial (6) opposite the trim indicator mark (3) on the pedestal edge lighted panel. (7) Remove aft elevator bellcrank rig pin (3) (Ref. Figure 201). (8) Install the lower aft rig pin access panel 34 on the left and right side of the upper vertical stabilizer (Ref. Chapter 6-50-00, STABILIZER ACCESS PANELS). (9) Rotate the elevator trim wheel until the 0 mark on position dial (6) is aligned with the trim indicator mark (3) on the pedestal (5) edge lighted panel (4) (Ref. Figure 203, Detail A).

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1. ELEVATOR 2. RUDDER 3. AFT ELEVATOR BELLCRANK RIG PIN 4. VERTICAL STABILIZER 5. HORIZONTAL STABILIZER 6. TRIM TAB

A 5

6 1

2 4

3

VIEW LOOKING UP LEFT HAND SIDE DETAIL

A

UC27B 041403AA.AI

Figure 201 Elevator Aft Bellcrank Rig Pin Installation

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B

DETAIL

A

A NOTE: THE CONTROL WHEEL GEAR (4) IS BONDED TO THE HUB (2) INBOARD OF THE TRIM CONTROL WHEEL (3).

4 2

6

DETAIL

C

5

3 10

7

C 1

9

10

8

DETAIL

B

Figure 202 Elevator Trim Tab Indicator Adjustment

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1. SCREW 2. CONTROL WHEEL HUB 3. TRIM CONTROL WHEEL 4. CONTROL WHEEL GEAR 5. REDUCTION GEAR 6. TRIM POSITION DIAL 7. CHAIN SPROCKET 8. WASHERS 9. MOUNT 10. SCREWS UC27B 041404AB.AI

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

*

*

1. TRIM CONTROL WHEEL 2. RED MARK 3. TRIM INDICATOR MARK 4. EDGELIGHTED PANEL 5. PEDESTAL 6. TRIM POSITION DIAL

A B C

NOTE: RED MARK SHOWN IS FOR REFERENCE ONLY. LOCATION MAY VARY BETWEEN AIRPLANES.

1

1

1

2* 2

3

*

3

3

4

4 4

6

6

6

5 5

5

DETAIL

A

DETAIL

B

DETAIL

C

UC27B 041402AB.AI

Figure 203 Elevator Trim Tab Indicator Adjustment

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FLIGHT CONTROLS STALL WARNING SYSTEM DESCRIPTION AND OPERATION

27-31-00 00

1. GENERAL The nucleus of the stall warning system installed in the airplane is the lift computer which processes information from the lift transducer and the flap downlimit switch assembly. The computer makes use of this information and compensates for changes in airplane attitude, airspeed and flap and gear position to sound the stall warning horn, thereby alerting the flight crew to an impending stall condition. The lift computer is grounded through the left landing gear safety switch which opens and disables the entire system when the airplane is on the ground. Power is supplied to the lift computer through a 5-ampere circuit breaker on the right circuit breaker panel. Power for the stall warning transducer heat is taken from a 15-ampere circuit breaker on the left inboard subpanel.

A. Lift Transducer The lift transducer, a wing mounted electro-mechanical devise, uses AC current supplied by the lift computer to produce a modulated AC current for input to the lift computer, which translates that signal into linear DC response. The transducer makes use of a moveable vane which protrudes into the airstream and senses local airflow and dynamic air pressure. Ice protection is provided by heating elements in the vane and in the mounting plate. The lift transducer heat elements are controlled through a relay mounted on the stall warning heat panel assembly. The heat control relay coil is energized through the left landing gear safety switch when the airplane is on the ground. Current to the heat elements is shunted around the open contacts of the heat control relay through series resistors, which reduce the voltage to the heat elements and prevent the elements from overheating during ground operation. During inflight operations, the ground circuit to the heat element control relay coil is opened by the landing gear safety switch, and bus voltage is shunted around the series resistors and applied directly to the transducer vane and mounting plate heat elements. An internal solenoid in the lift transducer, energized through the self-test circuit of the lift transducer when the self-test switch in the flight compartment is depressed, raises the vane of the transducer and simulates an impending stall condition, thereby sounding the stall warning horn. When the stall warning test switch is depressed, the landing gear safety switch grounding circuit is bypassed and the system is biased to ground at pin 8 of the lift computer. This is a positive test of the stall warning annunciator system only and does not verify the integrity of calibration, the landing gear safety switch circuit or the flap down limit switch circuits. WARNING: A test flight is required after performing maintenance, or replacing either the lift transducer assembly or the stall warning computer. Refer to STALL WARNING SYSTEM CALIBRATION and FLIGHT CALIBRATION AND ADJUSTMENT. The flap down limit switch assembly provides biasing inputs to the lift computer for flap settings of APPROACH and LANDING. These inputs are processed by the lift computer to compensate for changes in angle of attack as a result of changing the configuration of the wing through the application of flaps.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

B. Lift Computer The lift computer is a solid state electronic unit incorporating power distribution, computation, monitoring, test and warning functions. The computer will process signals from the transducer, the flap position sensor and a test switch in the crew compartment. The lift computer will supply a 28vdc stall warning output signal to the audio system to sound the stall warning horn when the speed ratio decreases below the value indicated for flap position.

Page 2 Nov 1/09

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FLIGHT CONTROLS STALL WARNING SYSTEM TROUBLESHOOTING

100100

1. PROCEDURES WARNING: When testing the anti-icing function of the transducer vane and mounting plate, use extreme caution as either surface can cause burns. It should be further noted that prolonged operation of the anti-ice function in still air could damage the transducer; therefore, it is imperative that the transducer heat be turned off immediately after completion of the heat test. Any time the airplane is on jacks or the landing gear squat switch is extended, the heat switch for the stall warning system, if turned on, will produce high heat at the transducer. The troubleshooting diagrams on the following pages provide for a logical sequence of checks to be made should a fault in the stall warning system become apparent. The particular nature of the fault will dictate which of the diagrams should be used. Three basic fault conditions resulting from pilot observations are defined and may be the only information available to the technician prior to troubleshooting the system. These fault conditions are stated at the beginning of each diagram. Following the logical flow of the diagram will provide the technician with essential information which will allow him to properly differentiate the fault and lead him to the proper remedial action. A sensitive precision voltmeter (digital preferred), accurate enough to detect a change of at least 0.1 volt, should be used for making voltage measurements. A breakout box designed for the purpose of gaining access to the various inputs and outputs for the lift computer may be build. Refer to Figure 102 for a diagram. The wires from each connector are appropriately connected to a terminal strip inside the box so that all inputs and outputs for the lift computer are carried through the breakout box. A series of jacks, correspondingly identified with pin locations, are used to tap off the desired signals for measurement with the voltmeter. Polarity must be closely observed, as both positive and negative voltage may be encountered when measuring outputs and inputs. The breakout box is connected in series between the airplane harness and the lift computer and provides for parallel access to the inputs and outputs. Breakout box jack locations (TP's or test points) are identified at each Step on the troubleshooting diagram each time a different set of jacks is to be used. TP 22 will not be found on the wiring schematic in Figure 101; TP 22 is for test purposes only and is a measure of transducer performance. A force applicator (P/N 1952-1) is used to apply forces to the transducer vane for checking the accuracy of the lift computer trip point calibration. Refer to Troubleshooting Charts in Figures 103, 104 and 105.

2. STALL WARNING SYSTEM CALIBRATION CAUTION: This procedure must be performed in its entirety in sequence. All portions of the calibration sections are to be completed. The operator should not try to perform only one calibration function. WARNING: The Air Calibration Section procedure is required. The Ground Calibration Section is a recommended procedure to be performed before the Air Calibration flight.

A. Ground Calibration Section Should any components of the stall warning system be replaced, a calibration check of the system should be performed before the airplane is released for flight check. The following procedure may be used for this purpose: (1) Actuate the squat switch.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (2) Install the force applicator at the stall warning lift transducer; do not apply any force to the vane. (3) Lower the flaps to the LANDING position. (4) Adjust potentiometer “2” on the lift computer until the stall warning horn just begins to sound. Note this position. Adjust the potentiometer until the sound stops. Note this position. Center the potentiometer between the two noted positions. (5) Retract the flaps to the APPROACH position and apply -2 TG of force to the transducer vane. (6) Adjust potentiometer “1” until the horn just begins to sound. Note this position. Adjust the potentiometer until the sound stops. Note this position. Center the potentiometer between the two noted positions. (7) Retract the flaps to the UP position and apply -3 TG of force to the transducer vane. (8) Adjust potentiometer “0” until the stall warning horn just begins to sound. Note this position. Adjust the potentiometer until the sound stops. Note this position. Center the potentiometer between the two noted positions. (9) In the event any of the previous adjustments cannot be made, refer to the appropriate troubleshooting Chart. (10) Recheck the tip gram figures after calibration.

B. Air Calibration Section CAUTION: This procedure must be performed in its entirety in sequence. All portions of the calibration sections are to be completed. The operator should not try to perform only one calibration function. After Calibration Checks (1) through (10) have been made, perform the following flight checks in smooth air at a safe altitude under stabilized conditions: (1) With the flaps and landing gear down and power off (idle power), slowly reduce speed (at a rate not higher than one knot per second) until the stall warning horn sounds. Note the stall airspeed at the time the warning horn actuates and continue to slowly reduce speed until the airplane stalls, then record the stall speed. The stall warning horn should actuate at an airspeed 5 to 10 knots faster than that at which the stall occurred. (2) With the flaps in the approach position and the landing gear down and power off (idle power), the stall warning should actuate within 5 to 10 knots above the stall. (3) With the flaps and landing gear up and power off (idle power), the stall warning should actuate within 5 to 10 knots above the stall. NOTE: If the stall warning does not actuate at the proper airspeed during the above procedures, gain access to the stall warning lift computer. The lift computer is located under the center aisle floorboard adjacent to the airstair door. Adjust pot No. 2 for full-down flaps, pot No. 1 for approach flaps and pot No. 0 for flaps up. Adjust each pot until the stall warning sounds within 5 to 10 knots of the stall. A clockwise adjustment increases the stall warning margin; a counterclockwise adjust decreases the stall warning margin.

Page 102 May 1/11

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Figure 101 Stall Warning System Schematic

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 102 Breakout Box

Page 104 May 1/11

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

STALL WARNING SYSTEM DOES NOT TEST

28 VDC BETWEEN TP 1 & 2

NO

IS STALL WARNING CONT CB IN?

NO

YES

RESET CB FAULT IN COMPUTER

NO

15 4 VDC BETWEEN TP 8 & 2

YES YES

28 VDC AT CONT CB

NO

TROUBLESHOOT POWER DISTRIBUTION

FAULT IN TEST SWITCH CIRCUIT

NO

VOLTAGE DROPS TO ZERO WHEN TEST SWITCH IS PRESSED

YES

YES

CORRECT OPEN BETWEEN CB AND COMPUTER

FAULT IN COMPUTER

NO

12 VDC BETWEEN TP 16 & 2 WHEN TEST SWITCH IS PRESSED

YES

FAULT IN TRANSDUCER OR CIRCUIT

NO

DOES TRANSDUCER VANE MOVE FORWARD WHEN TEST SWITCH IS PRESSED? YES

FAULT IN COMPUTER

NO

28 VDC BETWEEN TP 12 & 2 WHEN TEST SWITCH IS PRESSED YES

FAULT IN HORN OR CIRCUIT UE27B 991652AA.AI

Figure 103 TROUBLESHOOTING-STALL WARNING SYSTEM: STALL WARNING SYSTEM DOES NOT TEST

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STALL WARNING SYSTEM TEST OK, BUT HORN DOES NOT SOUND IN FLIGHT FAULT IN COMPUTER

10 .2 VAC BETWEEN TP 24 & 25

NO

YES FAULT IN TRANSDUCER

6 VDC BETWEEN TP 22 & 2

NO

YES 16 VDC BETWEEN TP 4 & 2

NO FAULT IN COMPUTER

INSTALL FORCE APPLICATOR WITH "0" FORCE ON VANE & LOWER FLAPS TO LANDING. DOES HORN SOUND NOW?

NO

FAULT IN TRANSDUCER

YES

YES

NO

16 VDC BETWEEN TP 6 & 2

NO

YES

DOES HORN SOUND AFTER ADJUSTING POT "2"?

FAULT IN FLAP LIMIT SWITCH

YES

RETRACT FLAPS TO APPROACH & APPLY -2 TG OF FORCE TO VANE. DOES HORN SOUND NOW?

5 VDC BETWEEN TP 22 & 2

NO

NO FAULT IN COMPUTER

FAULT IN TRANSDUCER

NO

4.5 VDC BETWEEN TP 22 & 2

NO

YES

DOES HORN SOUND AFTER ADJUSTING POT "0"?

YES

YES

RETRACT FLAPS TO UP & APPLY -3 TG OF FORCE TO VANE. DOES HORN SOUND NOW?

YES

DOES HORN SOUND AFTER ADJUSTING POT "1"

YES

YES

SYSTEM OPERATING NORMALLY

NO

FAULT IN COMPUTER

NO

FAULT IN COMPUTER *CONDITIONS: 28 VDC APU CONNECTED, EXTERNAL POWER AND BATTERY SWITCHES ON, LANDING GEAR SQUAT SWITCH BYPASSED BY GROUNDING TP 7

Figure 104 TROUBLESHOOTING-STALL WARNING SYSTEM: STALL WARNING TESTS OK BUT HORN DOES NOT SOUND IN FLIGHT Page 106 May 1/11

27-31-00

UE27B 991653AA.AI

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

STALL WARNING TEST OK, BUT HORN SOUNDS TOO EARLY OR TOO LATE FAULT IN COMPUTER

NO

10 .2 VAC BETWEEN TP 24 & 25 YES

INSTALL FORCE APPLICATOR WITH ZERO FORCE ON THE VANE & LOWER FLAPS TO LANDING FAULT IN TRANSDUCER

NO

6 VDC BETWEEN TP 22 & 2 YES

FAULT IN COMPUTER

NO

16 VDC BETWEEN TP 4 & 2 YES

FAULT IN FLAP DOWN LIMIT SWITCH

NO

16 VDC BETWEEN TP 6 & 2 YES

FAULT IN COMPUTER

NO

CAN POT "2" BE ADJUSTED UNTIL HORN JUST STARTS TO SOUND? YES

FAULT IN TRANSDUCER

NO

RETRACT FLAPS TO APPROACH & APPLY -2 TG OF FORCE TO THE VANE CHECK FOR 5 VDC BETWEEN TP 22 & 2 YES

FAULT IN COMPUTER

NO

CAN POT "1" BE ADJUSTED UNTIL HORN JUST STARTS TO SOUND? YES

FAULT IN TRANSDUCER

NO

RETRACT FLAPS TO UP & APPLY -3 TG OF FORCE TO VANE CHECK FOR 4.5 VDC BETWEEN TP 22 & 2 YES

FAULT IN COMPUTER

NO

CAN POT "0" BE ADJUSTED UNTIL HORN JUST STARTS TO SOUND? YES

STALL WARNING SYSTEM SHOULD BE PROPERLY CALIBRATED

*CONDITIONS: 28 VDC APU CONNECTED, EXTERNAL POWER AND BATTERY SWITCHES ON, LANDING GEAR SQUAT SWITCH BYPASSED BY GROUNDING TP 7.

UE27B 991654AA.AI

Figure 105 TROUBLESHOOTING-STALL WARNING SYSTEM: STALL WARNING TESTS OK BUT HORN SOUNDS TOO EARLY OR TOO LATE

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FLIGHT CONTROLS FLAPS DESCRIPTION AND OPERATION

27-50-00 00

1. GENERAL The flaps, two on each wing, are driven by an electric motor through a gearbox mounted on the forward side of the rear spar. The motor incorporates a dynamic brake, which helps to prevent overtravel of the flaps, through the use of two sets of motor windings. The gearbox drives four flexible drive shafts connected to jackscrews at each flap. The flap motor power circuit is protected by a 20-ampere circuit breaker, placarded FLAP MOTOR, located on the right circuit breaker panel. A 5-ampere circuit breaker, placarded FLAP IND & CONTROL, for the control circuit is also located on this panel (Ref. Figures 1 and 2). The flaps are operated by a sliding lever located just below the condition levers on the pedestal. The flap control is used to select UP, TAKEOFF, APPROACH, and LANDING flap positions. Flap position is indicated on the flap position indicator located directly above the engine condition levers in the pedestal. The indicator reads a signal from the flap position transmitter that is driven by the right inboard flap. The warning horn in the cockpit overhead will sound when the flap position is not compatible with engine power lever settings and landing gear positions. The flap travel limit switches, located outboard of the right nacelle, along with the limit switches in the pedestal provide signals to the flap warning circuit. A flap safety switch, located at the junction of the adjacent flaps in each wing, disables power to the flap motor when either of the flaps in that pair travels more than 3° to 6° out of phase with the adjacent flap.

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Figure 1 Flap Control System

Page 2 Nov 1/09

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 2 Flap System Wiring Diagram

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

FLIGHT CONTROLS FLAPS MAINTENANCE PRACTICES

200200

1. OUTBOARD FLAP A. Removal (1) Connect external electrical power to the airplane. (2) Select the BATT switch to the ON position. (3) Select the EXT PWR switch to the EXT PWR position. (4) Lower flaps to the full down position, then open the FLAP IND & CONTROL circuit breaker located on the right circuit breaker panel Zone 246 (Ref. Chapter 6-40-00, ZONE DIAGRAM). (5) Select the EXT PWR switch to the OFF position. (6) Select the BATT switch to the OFF position. (7) Disconnect external electrical power from the airplane. (8) Attach a red tag to the flap control lever with the words “Do Not Operate, Maintenance In Progress”. (9) Remove the bolt (13), washers (12) and nut (11) securing the flap actuator (10) to the bracket (3) on the leading edge of the flap (2). Slide the end of the actuator (10) free of the bracket (3) and remove the bushing (14) from the end of the actuator (10) (Ref. Figure 202). NOTE: Measure the extension of the flap actuator (10) screw so it can be attached in its original position. (10) Remove the patch plates (5) from the top of the flap on each side of the inboard and outboard roller brackets. (11) Disconnect the end of the bonding jumpers (7) from the flap cove by removing screws (8) and washers (9). (12) Disconnect the flap safety switch mechanism (Ref. 27-50-06). (13) Remove the cotter pin (14), nut (13), washer (12) and bolt (11) from the outboard flap outboard track aft slot (Ref. Figure 204). (14) Remove the cotter pin (19), nut (18), washer (17) and bolt (16) from the outboard flap inboard track aft slot. (15) While providing support for the flap, remove the cotter pin (4), nut (3), washer (2) and bolt (1) from the outboard flap outboard track forward slot. (16) While providing support for the flap, remove the cotter pin (9), nut (8), washer (7) and bolt (6) from the outboard flap inboard track forward slot. (17) Remove flap from the wing.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (18) Remove the bearing (5) from the outboard flap outboard track forward slot. (19) Remove the bearing with bonded washer (15) from the outboard flap outboard track aft slot. (20) Remove the bearing (10) from the outboard flap inboard track forward slot. (21) Remove the bearing with bonded washer (20) from the outboard flap inboard track aft slot. CAUTION: Never operate the flaps with the RH inboard flap removed. The limit switches will be inoperative and serious damage to the flaps, wings and actuating system could result. (22) Perform FLAP ROLLER BRACKET CHECK as outlined in this section.

B. Installation NOTE: It is also permissible to lubricate the flap tracks using dry film lubricant (82, Table 1, Chapter 91-00-00). If using this type of lubricant, it will be necessary to allow for proper cure time following the manufacturer’s instructions. (1) Lubricate flap tracks using lubricant (68 and 81, Table 1, Chapter 91-00-00). (2) Position the bearing (5) in the outboard flap outboard track forward slot (Ref. Figure 204). (3) Position the bearing with bonded washer (15) in the outboard flap outboard track aft slot. (4) Position the bearing (10) in the outboard flap inboard track forward slot. (5) Position the bearing with bonded washer (20) in the outboard flap inboard track aft slot. CAUTION: To prevent flap track damage, install the flap rollers so that the flanges are facing each other to the center of the flap. The aft slot rollers have a teflon washer bonded to the roller flange. NOTE: It is permissible to use a bolt one size longer or shorter as necessary to ensure nut is secured. Table 201 OUTBOARD FLAP ATTACH BOLTS Bolt Part No.

Bracket, Doubler, Bushing Combination Installed

NAS1303-12D

standard bolt - no doublers installed

NAS1303-13D

0.063 inch doubler on one bracket

NAS1303-14D

0.063 inch doubler on two brackets or one 0.063-inch doubler and flanged bushing in one bracket

NAS1303-15D

0.063 inch doubler on two brackets and a flanged bushing in one bracket

NAS1303-16D

0.063 inch doubler on two brackets and flanged bushings in both brackets

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(6) With assistance, position the flap to the tracks. Align holes and install the bolt (1), washer (2), nut (3) and cotter pin (4) to the outboard flap outboard track forward slot. (7) With assistance, position the flap to the tracks. Align the holes and install the bolt (6), washer (7), nut (8) and cotter pin (9) to the outboard flap inboard track forward slot. (8) Align holes and install the bolt (16), washer (17), nut (18) and cotter pin (19) to the outboard flap inboard track aft slot. (9) Align holes and install the bolt (11), washer (12), nut (13) and cotter pin (14) to the outboard flap outboard track aft slot. (10) Ensure that the flap actuator (10) is extended to the same length that was measured during removal, then install the bushing (14) in the end of the actuator (10) (Ref. Figure 202). (11) Align the end of the actuator (10) with holes in the bracket (3) on the flap and install the bolt (13), washers (12) and nut (11) securing the end of the actuator (10) in place. (12) Connect the flap safety switch mechanism (Ref. 27-50-06). (13) Connect the end of the bonding jumpers (7) to the flap cove by installing screws (8) and washers (9). NOTE: With the flaps fully retracted and the aileron in the neutral position, the clearances noted in the AILERON CLEARANCES illustration in Chapter 27-10-00 must be maintained. The gap between the aileron and the outboard flap should be constant within ± 0.06 inch from the leading edge to the trailing edge. These dimensions do not apply to the lower forward area where the aileron tapers outboard. (14) Install the patch plates (5) on each side of the flap brackets for the flap rollers. (15) Remove the red tag from the flap control lever. (16) Perform the FLAP SYSTEM RIGGING procedure (Ref. 27-50-05).

C. Flap Roller Bracket Check (1) Perform INBOARD FLAP REMOVAL and/or OUTBOARD FLAP REMOVAL procedures in this section. (2) Inspect the flap roller brackets, roller bearings and attachment hardware for wear. (3) If any wear is noted, repair flap roller bracket as required (Ref. Model 1900 Airliner Series Structural Repair Manual (SRM) Chapter 57-92-03 and/or 57-92-04). (4) Perform INBOARD FLAP INSTALLATION and/or OUTBOARD FLAP INSTALLATION procedures in this section.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

1. FLAP SAFETY SWITCH BRACKET 2. INBOARD FLAP OUTBOARD TRACK 3. INBOARD FLAP 4. PATCH PLATES 5. INBOARD FLAP ACTUATOR BRACKET 6. INBOARD FLAP INBOARD TRACK 7. FLAP LIMIT SWITCH BRACKET 8. BONDING JUMPER 9. SCREW 10. WASHER 11. FLAP LIMIT SWITCH ARM 12. NUT 13. WASHERS 14. BOLT 15. FLAP ACTUATOR 1 16. NUT 17. WASHERS 18. BOLT 19. BUSHING

20. FLAP TRANSMITTER LINKAGE 21. NUT 22. WASHERS 23. BOLT 24. FLAP TRANSMITTER BRACKET 25. COTTER PIN

A A

4

2

4 3

24

5 4 6

4

B E 10 9 DETAIL

A

D

C B

7

8

DETAIL

B

18

17 23

14

17 16

13 13 12

15

22 21

20

11 19 25 RIGHT SIDE ONLY DETAIL

DETAIL

22

D

C

DETAIL

Figure 201 Inboard Flap Installation

Page 204 Nov 1/09

27-50-00

RIGHT SIDE ONLY

E

UC27B 045948AA.AI

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

1. OUTBOARD FLAP OUTBOARD TRACK 2. OUTBOARD FLAP 3. OUTBOARD FLAP ACTUATOR BRACKET 4. OUTBOARD FLAP INBOARD TRACK 5. PATCH PLATES 6. FLAP SAFETY SWITCH BRACKET 7. BONDING JUMPER 8. SCREW 9. WASHER 10. FLAP ACTUATOR 11. NUT 1 12. WASHERS 13. BOLT 14. BUSHING

A A 5 2

3

B

4 5

DETAIL

A

C 6

B

13

9 8

12 12 11 10 7 14 DETAIL DETAIL

B

C UC27B 045946AA.AI

Figure 202 Outboard Flap Installation

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1. BOLT * 2. LARGE PHENOLIC WASHER 3. SMALL WASHER 4. NUT 5. COTTER PIN 6. BEARING 7. BOLT * 8. LARGE PHENOLIC WASHER 9. SMALL WASHER 10. NUT 11. COTTER PIN 12. BEARING 13. LARGE THIN TEFLON WASHER * NOTE: THE FORWARD ROLLERS PHENOLIC WASHERS (2 AND 8) MUST BE 1.25 INCH OUTSIDE DIAMETER.

2 8

3 1

9 7

4

10

11

5

13

6

INBOARD FLAP OUTBOARD TRACK FWD & AFT ROLLERS

12

INBOARD FLAP INBOARD TRACK FWD & AFT ROLLERS

UC27B 045665AA.AI

Figure 203 Inboard Flap Rollers

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1. BOLT 2. WASHER 3. NUT 4. COTTER PIN 5. BEARING 6. BOLT 7. WASHER 8. NUT 9. COTTER PIN 10. BEARING 11. BOLT 12. WASHER 13. NUT 14. COTTER PIN 15. BEARING WITH BONDED WASHER 16. BOLT 17. WASHER 18. NUT 19. COTTER PIN 20. BEARING WITH BONDED WASHER

2

1

7 3

6

8

4

9

10

5

OUTBOARD FLAP OUTBOARD TRACK FWD ROLLER

OUTBOARD FLAP INBOARD TRACK FWD ROLLER

12

11

17

16 13

18

14

15

OUTBOARD FLAP OUTBOARD TRACK AFT ROLLER

19 20

OUTBOARD FLAP INBOARD TRACK AFT ROLLER UC27B 045666AA.AI

Figure 204 Outboard Flap Rollers

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

2. INBOARD FLAP A. Removal (1) Connect external electrical power to the airplane. (2) Select the BATT switch to the ON position. (3) Select the EXT PWR switch to the EXT PWR position. (4) Lower flaps to the full down position, then open the FLAP IND & CONTROL circuit breaker located on the right circuit breaker panel Zone 246 (Ref. Chapter 6-40-00, ZONE DIAGRAM). (5) Select the EXT PWR switch to the OFF position. (6) Select the BATT switch to the OFF position. (7) Disconnect external electrical power from the airplane. (8) Attach a red tag to the flap control lever with the words “Do Not Operate, Maintenance In Progress”. (9) Remove the bolt (18), washers (17) and nut (16) securing the flap actuator (15) to the bracket (5) on the leading edge of the flap. Slide the end of the actuator (15) free of the bracket (5) and remove the bushing (19) from the end of the actuator (15) (Ref. Figure 201). NOTE: Measure the extension of the flap actuator (15) screw so it can be attached in its original position. (10) Remove the patch plates (4) from the top of the flap (3) on each side of the inboard and outboard roller brackets. (11) Disconnect the end of the bonding jumpers (8) from the flap cove by removing screws (9) and washers (10). (12) On the RH inboard flap only, disconnect the linkage (20) to the flap position transmitter from the bracket (24) located on the leading edge of the flap (3) approximately 19 inches outboard of the inboard flap track (6). (13) Disconnect the flap safety switch mechanism (Ref. 27-50-06). (14) Disconnect the flap limit switch arm (11) located on the RH inboard flap by removing cotter pin (25), nut (12) washers (13) and bolt (14). (15) Remove the cotter pin (5), nut (4), washer (3) and bolt (1) from the inboard flap outboard track aft slot (Ref. Figure 203). (16) Remove the cotter pin (11), nut (10), washer (9) and bolt (7) from the inboard flap inboard track aft slot. (17) While providing support for the flap, remove the cotter pin (5), nut (4), washer (3) and bolt (1) from the inboard flap outboard track forward slot. (18) While providing support for the flap, remove the cotter pin (11), nut (10), washer (9) and bolt (7) from the inboard flap inboard track forward slot.

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27-50-00

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (19) Remove flap from the wing. (20) Remove the bearing (6) and large phenolic washer (2) from the forward and aft slots of the inboard flap outboard tracks. NOTE: A steel washer (2) may be installed on the forward roller of the inboard flap outboard track instead of a phenolic washer. It is recommended to discard the steel washer and install a phenolic washer (2) in its place. (21) Remove the bearing (12), large phenolic washer (8) and large thin teflon washer (13) from the forward and aft slots of the inboard flap inboard tracks. CAUTION: Never operate the flaps with the RH inboard flap removed. The limit switches will be inoperative and serious damage to the flaps, wings and actuating system could result. (22) Perform FLAP ROLLER BRACKET CHECK as outlined in this section.

B. Installation NOTE: It is also permissible to lubricate the flap tracks using dry film lubricant (82, Table 1, Chapter 91-00-00). If using this type of lubricant, it will be necessary to allow for proper cure time following the manufacturer’s instructions. (1) Lubricate flap tracks using lubricant (68 and 81, Table 1, Chapter 91-00-00). (2) Position the bearing (6 and 12) and large phenolic washer (2 and 8) in the forward slots of the inboard flap inboard and outboard flap tracks. Install a large thin teflon washer (13) on the inboard flap inboard track only (Ref. Figure 203). NOTE: A steel washer (2) may be installed on the forward roller of the inboard flap outboard track instead of a phenolic washer. It is recommended to discard the steel washer and install a phenolic washer (2) in its place. (3) With assistance, position the flap to the tracks. Align the holes and install the bolts (1 and 7), washers (3 and 9), nuts (4 and 10) and cotter pin (5 and 11) to the forward slots of the inboard flap inboard and outboard flap tracks. (4) Position the bearing (6 and 12) and large phenolic washer (2 and 8) in the aft slots of the inboard flap inboard and outboard flap tracks. Install a large thin teflon washer (13) on the inboard flap inboard track only. (5) With assistance, position the flap to the tracks. Align the holes and install the bolts (1 and 7), washers (3 and 9), nuts (4 and 10) and cotter pin (5 and 11) to the aft slots of the inboard flap inboard and outboard flap tracks. CAUTION: To prevent flap track damage, install the flap rollers with the phenolic washer between each flap roller bracket and the track (facing each other) to the center of the flap. NOTE: It is permissible to use a bolt one size longer or shorter as necessary to ensure nut is secured. It is permissible to add an additional 118-100000-9 teflon washer to the aft slot and a 118-100000-15 teflon washer to the forward slot to provide proper clearance where necessary.

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Table 202 INBOARD FLAP ATTACH BOLTS Bolt Part No.

Bracket Doubler Combination Installed

NAS1305-14D

standard bolt - no doublers installed

NAS1305-15D

0.063 inch doubler on one bracket

NAS1305-16D

0.063 inch doubler on two brackets or one 0.063 inch doubler and flanged bushing on one bracket

NAS1305-17D

0.063 inch doubler on two brackets and a flanged bushing in one bracket

NAS1305-18D

0.063 inch doubler on two brackets and flanged bushings in both brackets

(6) Ensure that the flap actuator (15) is extended to the same length that was measured during removal, then install the bushing (19) in the end of the actuator (15) (Ref. Figure 201). (7) Align the end of the actuator (15) with holes in the bracket (5) on the flap (3) and install the bolt (18), washers (17) and nut (16) securing the end of the actuator (15) in place. (8) Connect the end of the bonding jumpers (8) to the flap cove by installing screws (9) and washers (10). (9) Connect the flap safety switch mechanism (Ref. 27-50-06). (10) On the RH inboard flap only, connect the linkage (20) to the flap position transmitter bracket (24) located on the leading edge of the flap (3) approximately 19 inches outboard of the inboard flap track (6) by installing the bolt (23), washers (22) and nut (21). (11) Connect the flap limit switch arm (11) located on the RH inboard flap by installing bolt (14), washers (13), nut (12) and cotter pin (29). (12) Install the patch plates (4) on each side of the inboard and outboard roller brackets. (13) Remove the red tag from the flap control lever. (14) Perform the FLAP SYSTEM RIGGING procedure (Ref. 27-50-05).

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FLIGHT CONTROLS FLAP CABLES MAINTENANCE PRACTICES

27-50-01 200200

1. PROCEDURES A. Flap Drive Cables When installing the driveshaft assemblies, the direction of internal cable twist (or lay) must be determined. Do this by looking at the diagonal lines on either of the square ends of the driveshaft (Ref. FLAP FLEXIBLE DRIVESHAFT INSTALLATION Figure in Chapter 25-50-03). A right-lay, 52-inch long driveshaft connects the left forward attach point of the flap gearbox to the left inboard flap actuator. A right-lay, 140-inch long driveshaft connects the right aft attach point of the flap gearbox to the right outboard flap actuator. A left-lay, 140-inch long driveshaft connects the left aft attach point of the flap gearbox to the left outboard flap actuator. A left-lay, 52-inch long driveshaft connects the right forward attach point of the flap gearbox to the right inboard flap actuator. Each driveshaft has a ferrule on one end that is 2-inches long and a ferrule on the other end that is 2.5-inches long. Connect the end of the driveshaft with the 2-inch ferrule to the flap gearbox.

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FLIGHT CONTROLS FLAP TRACKS MAINTENANCE PRACTICES

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1. PROCEDURES A. Wear Limits The maximum slot dimensions, including allowable wear, are 0.785 inch for the flap tracks on the outboard flaps and 1.038 inches for the flap tracks on the inboard flaps. The allowable wear into the track side surface is 0.050 inch (Ref. Figure 201). Track wear within the above limitations may be dressed smooth with light emery cloth to prevent roller binding. The teflon chafing washers should also be kept in good condition. To help reduce wear, it is recommended that the tracks and rollers be kept clean and lubricated with Lubriplate 130A or Lubriplate Aero (68, Table 1, Chapter 91-00-00).

THE MAXIMUM FLAP TRACK SLOT DIMENSION WITH ALLOWABLE WEAR IS: OUTBOARD FLAP TRACKS - 0.785 INCH INBOARD FLAP TRACKS - 1.038 INCHES

THE MAXIMUM ALLOWABLE WEAR TO THE FLAP TRACK SIDE SURFACE IS 0.050 INCH.

UC27B 070984AA.AI

Figure 201 Flap Track Wear Limits

B. Repair Reference the Model 1900 Airline Series Structural Repair Manual, Chapter 57-92-01.

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FLIGHT CONTROLS FLAP MOTOR AND GEARBOX MAINTENANCE PRACTICES

27-50-03 200200

1. PROCEDURES A. Removal (1) Connect external electrical power to the airplane. (2) Select the BATT switch to the ON position. (3) Select the EXT PWR switch to the EXT PWR position. (4) Lower flaps to the full down position, then open the FLAP IND & CONTROL circuit breaker located on the right circuit breaker panel Zone 246 (Ref. Chapter 6-40-00, ZONE DIAGRAM). (5) Select the EXT PWR switch to the OFF position. (6) Select the BATT switch to the OFF position. (7) Disconnect external electrical power from the airplane. (8) Attach a red tag to the flap control lever with the words “Do Not Operate, Maintenance In Progress”. (9) Remove the aft center isle main spar ramp. (10) Remove the center aisle passenger compartment carpet (Ref. Chapter 25-20-01, CARPET REMOVAL AND INSTALLATION) as required to access floor access panel 10 (UA-1 and After) or 11 (UB-1 and After and UC-1 and After). (11) Remove the center aisle floorboard 10 (UA-1 and After) or 11 (UB-1 and After and UC-1 and After) (Ref. Chapter 6-50-00, FLOOR ACCESS PANELS). (12) Remove left floorboard panel 16E (Ref. Chapter 6-50-00, FLOOR ACCESS PANELS). (13) Tag and disconnect the flap motor electrical leads from the flap motor relay. (14) If necessary, remove or loosen clamps securing the flexible drive shafts to allow removal of the flap gearbox and motor. (15) Remove the safety wire (4) from the retaining nuts (1) securing the flexible drive shafts (3) to the flap gearbox and motor (2) (Ref. Figure 201). (16) Remove the flexible drive shafts (3) from the flap gearbox and motor (2). (17) Remove the cotter pins (6), nuts (7) and washers (8) from the studs and remove the flap gearbox and motor (2) from the airplane.

B. Installation (1) Prepare the area where the flap motor attaches to the spar at FS 326.00 for electrical bonding (Ref. Chapter 20-00-01, PREPARATION OF SURFACE).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (2) Position the flap gearbox and motor (2) over the studs at FS 326.00 and install washers (8), nuts (7) and cotter pins (6) (Ref. Figure 201). (3) Install the flexible drive shafts (3) to the flap gearbox and motor (2). Tighten the nuts (1) and safety wire (4). NOTE: Do not exceed 75 inch-pounds torque when tightening the flexible drive shaft (3) retaining nuts (1). (4) Connect the electrical leads from the flap motor to the flap motor relay and torque to no more than 19 inch-pounds. Remove the tags from the wiring. (5) Remove the red tag from the flap control lever. (6) Connect external electrical power to the airplane. (7) Select the BATT switch to the ON position. (8) Select the EXT PWR switch to the EXT PWR position. (9) Close the FLAP IND & CONTROL circuit breaker located on the right circuit breaker panel Zone 246 (Ref. Chapter 6-40-00, ZONE DIAGRAM). CAUTION: With assistance, observe flap movement and be prepared to open the FLAP IND & CONTROL circuit breaker to stop the flaps. (10) Perform the FLAP SYSTEM FUNCTIONAL CHECK procedure (Ref. 27-50-05). (11) Install the center aisle floorboard 10 (UA-1 and After) or 11 (UB-1 and After and UC-1 and After) (Ref. Chapter 6-50-00, FLOOR ACCESS PANELS). (12) Install left floorboard panel 16E (Ref. Chapter 6-50-00, FLOOR ACCESS PANELS). (13) Install the center aisle passenger compartment carpet (Ref. Chapter 25-20-01, CARPET REMOVAL AND INSTALLATION).

C. Flap Motor Gearbox Freedom of Movement Check Engage the drive shaft of the gearbox with a tool that matches the size of the spline end of the motor drive shaft, then rotate the gearbox by hand. Any gearbox that fails to rotate freely under hand pressure must be replaced.

2. FLAP MOTOR A. Removal (1) Open the FLAP IND & CONTROL circuit breaker located on the right circuit breaker panel Zone 246 (Ref. Chapter 6-40-00, ZONE DIAGRAM). (2) Attach a red tag to the flap control lever with the words “Do Not Operate, Maintenance In Progress”. (3) Remove the aft center isle main spar ramp.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (4) Remove the center aisle passenger compartment carpet as required to access floor access panel 10 (UA-1 and After) or 11 (UB-1 and After and UC-1 and After) (Ref. Chapter 25-20-01, CARPET REMOVAL AND INSTALLATION). (5) Remove the center aisle floorboard 10 (UA-1 and After) or 11 (UB-1 and After and UC-1 and After) (Ref. Chapter 6-50-00, FLOOR ACCESS PANELS). (6) Tag and disconnect the flap motor electrical leads from the flap motor relay. (7) Remove the cotter pins (6), nuts (7) and washers (8) from the studs of the flap gearbox and motor (2) (Ref. Figure 201). NOTE: Use care to avoid damaging the flap gearbox and motor assembly and/or the aft spar at FS 326.00. (8) Remove the flap gearbox and motor (2) from the aft spar studs and rotate to gain access to the flap motor mount screws (9) located on the bottom of the motor. (9) Remove the mount screws (9) from the flap motor and remove the motor from the flap gearbox.

B. Installation (1) Prepare the area where the flap motor attaches to the spar at FS 326.00 for electrical bonding (Ref. Chapter 20-00-01, PREPARATION OF SURFACE). NOTE: Use care to avoid damaging the flap motor gearbox assembly and/or the aft spar at FS 326.00. (2) Rotate the flap gearbox, position the flap motor onto the flap gearbox and install the two mount screws (9) using thread lock (161, Table 1, Chapter 91-00-00) and tighten. As an option, it is acceptable to stake the flat head of the mount screws at slots for added security. (3) Position the flap gearbox and motor (2) over the studs on the aft spar at FS 326.00 and install washers (8), nuts (7) and cotter pins (6) (Ref. Figure 201). (4) Connect the electrical leads from the flap motor to the flap motor relay and torque to no more than 19 inch-pounds. Remove the tags from the wiring. (5) Remove the red tag from the flap control lever. (6) Connect external electrical power to the airplane. (7) Select the BATT switch to the ON position. (8) Select the EXT PWR switch to the EXT PWR position. (9) Close the FLAP IND & CONTROL circuit breaker located on the right circuit breaker panel Zone 246 (Ref. Chapter 6-40-00, ZONE DIAGRAM). CAUTION: With assistance, observe flap movement and be prepared to open the FLAP IND & CONTROL circuit breaker to stop the flaps. (10) Perform the FLAP SYSTEM FUNCTIONAL CHECK procedure (Ref. 27-50-05).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (11) Install the center aisle floorboard 10 (UA-1 and After) or 11 (UB-1 and After and UC-1 and After) (Ref. Chapter 6-50-00, FLOOR ACCESS PANELS). (12) Install the center aisle passenger compartment carpet (Ref. Chapter 25-20-01, CARPET REMOVAL AND INSTALLATION). (13) Install the aft center isle main spar ramp. (14) Select the EXT PWR switch to the OFF position. (15) Select the BATT switch to the OFF position. (16) Disconnect external electrical power from the airplane.

3. FLAP FLEXIBLE DRIVE SHAFT A. Removal (1) Connect external electrical power to the airplane. (2) Select the BATT switch to the ON position. (3) Select the EXT PWR switch to the EXT PWR position. (4) Lower flaps to the full down position, then open the FLAP IND & CONTROL circuit breaker located on the right circuit breaker panel Zone 246 (Ref. Chapter 6-40-00, ZONE DIAGRAM). (5) Select the EXT PWR switch to the OFF position. (6) Select the BATT switch to the OFF position. (7) Disconnect external electrical power from the airplane. (8) Attach a red tag to the flap control lever with the words “Do Not Operate, Maintenance In Progress”. (9) Remove the left and right passenger seats, located immediately forward of the rear spar (Ref. Chapter 25-20-00, PASSENGER SEAT REMOVAL). (10) Remove the left, right and center aisle passenger compartment carpet (Ref. Chapter 25-20-01, CARPET REMOVAL AND INSTALLATION) as required to access floor access panels 16E, 10 (UA-1 and After) or 11 (UB-1 and After and UC-1 and After) and 17E. (11) Remove floor panels 16E, 10 (UA-1 and After) or 11 (UB-1 and After and UC-1 and After) and 17E (Ref. Chapter 6-50-00, FLOOR ACCESS PANELS). (12) Remove wing access panels 8, 27, 28, 35 and 36 (UA-1 and After; UB-1 and After) or 17 (UC-1 and After) (Ref. Chapter 6-50-00, WING ACCESS PANELS). (13) Remove two Plexiglas panels in the aft main landing gear wheel well area. (14) Remove the seal grommet at the root rib, BL 28.28, and the clamps along the length of the flexible drive shaft (3) (clamps are located at RBL 25.55, WS 41.85, WS 46.25, WS 67.05, WS 104.25, and WS 124.59) (Ref. Figure 201).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL NOTE: The clamp located at RBL 25.55 is accessed through the open floor panel. The remaining clamps located at WS 41.85 and outboard are accessed through the access panels located at the aft lower surface of the wing at the same WS locations. Clamps located at WS 67.05 and outboard are for the outboard flap flexible drive shaft (3) only. If the flexible drive shaft (3) is to be reused, exercise care that the outer plastic housing is not damaged. CAUTION: Ensure that the flap actuator 90° drive assembly union does not loosen while loosening the flexible drive shaft retaining nut. (15) Remove the safety wire (4) from the flexible drive shaft (3) retaining nuts (1) located at both ends of the assembly. (16) Remove the flap flexible drive shaft (3).

B. Installation CAUTION: When installing the flexible drive shafts (3), the direction of internal cable twist, or lay must be determined. Do this by looking at the diagonal lines on either of the square ends of the flexible drive shaft (3) (Ref. Figures 201 and 202). A right-lay, 52 inch long flexible drive shaft (3) connects the left forward attach point of the flap gearbox (2) to the left inboard flap actuator. A right-lay, 140 inch long flexible drive shaft (3) connects the right aft attach point of the flap gearbox (2) to the right outboard flap actuator. A left-lay, 140 inch long flexible drive shaft (3) connects the left aft attach point of the flap gearbox (2) to the left outboard flap actuator. A left-lay, 52 inch long flexible drive shaft (3) connects the right forward attach point of the flap gearbox (3) to the right inboard flap actuator. Each flexible drive shaft (3) has a ferrule on one end that is two inches long and a ferrule on the other end that is 2.5 inches long. Connect the end of the flexible drive shaft (3) with the two inch ferrule to the flap gearbox (2). NOTE: Exercise care during handling and installation of the flexible drive shaft (3) to prevent the outer plastic housing from being nicked or cut through. Such damage can allow moisture to infiltrate the flexible drive shaft (3) and cause premature failure. Damage to the housing during installation may be avoided by temporarily applying a tough vinyl or other heavy tape to any sharp edges of nearby structure and holes through frames not containing permanent grommets. (1) Apply a light coat of grease (1, Table 2, 27-00-00) to the square ends of the flexible drive shaft(s) (3) (Ref. Figure 201). (2) Carefully route the flexible drive shaft (3) from the flap gearbox (2) toward the actuator (5). NOTE: Do not exceed 75 inch-pounds torque when tightening the flexible drive shaft (3) retaining nuts (1). (3) Install the flexible drive shaft (3) to the flap gearbox and motor (2). Tighten the retaining nut (1) and safety wire (4). (4) Install the flexible drive shaft (3) to the flap actuator (5). Tighten the retaining nut and safety wire.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (5) Replace any clamps removed along the length of the flexible drive shaft (3) (clamps are located at RBL 25.55, WS 41.85, WS 46.25, WS 67.05, WS 104.25, and WS 124.59). NOTE: The clamp located at RBL 25.55 is accessed through the open floor panel. The remaining clamps located at WS 41.85 and outboard are accessed through the access panels located at the aft lower surface of the wing at the same WS locations. Clamps located at WS 67.05 and outboard are for the outboard flap flexible drive shaft (3) only. (6) Install a new seal grommet where the flexible drive shaft (3) penetrates the root rib at BL 28.28. To ensure a pressure-tight seal, coat the exposed portions of the grommet with sealant (21, Table 2, Chapter 91-00-00). (7) Remove the red tag from the flap control lever. CAUTION: With assistance, observe flap movement and be prepared to open the FLAP IND & CONTROL circuit breaker to stop the flaps. (8) If the flexible drive shaft (3) for the right-hand inboard flap is replaced, rig the flap system (all flaps) (Ref. 27-50-05, FLAP SYSTEM RIGGING). Replacement of any other flexible drive shaft (3) requires the rigging of that particular flap, only. (9) Install wing access panels 8, 27, 28, 35 and 36 (UA-1 and After; UB-1 and After) or 17 (UC-1 and After) (Ref. Chapter 6-50-00, WING ACCESS PANELS). (10) Install two Plexiglas panels in the aft main landing gear wheel well area. (11) Install floor panels 16E, 10 (UA-1 and After) or 11 (UB-1 and After and UC-1 and After) and 17E (Ref. Chapter 6-50-00, FLOOR ACCESS PANELS). (12) Install the left, right and center aisle passenger compartment carpet (Ref. Chapter 25-20-01, CARPET REMOVAL AND INSTALLATION). (13) Install the left and right passenger seats (Ref. Chapter 25-20-00, PASSENGER SEAT INSTALLATION).

C. Outer Housing Inspection and Repair Cuts or nicks that penetrate the flexible drive shaft (3) outer plastic housing may allow moisture to enter, promoting lubrication failure and internal corrosion. Such flexible drive shaft (3) housing damage within the pressurized compartment may be especially harmful. Any pressurization air leakage through the flexible drive shaft (3) housing, though minor, may carry significant amounts of moisture, condensing and pooling in the colder outboard portion of the flexible drive shaft (3) (Ref. Figure 201). Inspect the flexible drive shaft (3) housings for nicks and cuts that penetrate. Replace the flexible drive shaft (3) or patch with Tedlar tape (13, Table 2, 27-00-00). If patching, clean the area with solvent (14, Table 2, 27-00-00) and wipe dry. Apply tape with a spiral wrap, tight and wrinkle free, overlapping each wrap by two-thirds. The patch should extend a minimum of two inches on each side of the damaged area. The free end of the tape patch may be additionally secured with one or two small nylon tie-wraps obtained locally.

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D. Inner Shaft Inspection Inspect the condition of the ends of the flexible drive shafts (3). Replace the flexible drive shaft (3) if either square shaft end is frayed or worn to less than 0.172-inch as measured between any two adjacent corners. (1) Connect external electrical power to the airplane. (2) Select the BATT switch to the ON position. (3) Select the EXT PWR switch to the EXT PWR position. (4) Lower flaps to the full down position, then open the FLAP IND & CONTROL circuit breaker located on the right circuit breaker panel, Zone 246 (Ref. Chapter 6-40-00, ZONE DIAGRAM). (5) Select the EXT PWR switch to the OFF position. (6) Select the BATT switch to the OFF position. (7) Disconnect external electrical power from the airplane. (8) Attach a red tag to the flap control lever with the words “Do Not Operate, Maintenance In Progress”. (9) Remove the aft center aisle main spar ramp. (10) Remove the center aisle passenger compartment carpet (Ref. Chapter 25-20-01, CARPET REMOVAL AND INSTALLATION) as required to access floor access panel 10 (UA-1 and After) or 11 (UB-1 and After and UC-1 and After). (11) Remove the center aisle floorboard 10 (UA-1 and After) or 11 (UB-1 and After and UC-1 and After) (Ref. Chapter 6-50-00, FLOOR ACCESS PANELS). (12) Remove the left floorboard panel 16E (Ref. Chapter 6-50-00, CABIN FLOORBOARD PANELS). (13) Remove the safety wire (4) from the retaining nuts (1) securing the flexible drive shafts (3) to the flap gearbox and motor (2) (Ref. Figure 201). (14) Remove the flexible drive shafts (3) from the flap gearbox and motor (2). (15) Confirm that the retainer shown in Figure 203 is attached to the inner shaft inside of the ferrule. (16) If the retainer is attached, proceed to Step (17). If the retainer is not attached, remove and replace the flap flexible shaft assembly. (17) Measure the distance from the end of the inner shaft to the closest edge of the retainer (Ref. Figure 203). (18) If the measurement is 0.78 inch to 0.87 inch, proceed to Step (19). If the measurement is less than 0.78 inch or more than 0.87 inch, remove and replace the flap flexible shaft assembly.

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(19) To determine if the retainer is attached at the outboard end of the inner shaft, establish the end play as detailed in the following Steps: (a) Push the inner shaft into the ferrule to the initial point of resistance and hold this position while measuring the distance from the end of the inner shaft to the closest edge of the ferrule. Record the distance. (b) Pull the end of the inner shaft out of the ferrule to the point of resistance, measuring the distance from the end of the inner shaft to the closest edge of the ferrule. Record the distance. NOTE: Do not exceed 75 inch-pounds torque when tightening the flexible drive shaft retaining nuts. (c) If the difference between the measurements is less than 0.50 inch, reengage the flexible casing over the end of the inner shaft. Connect the shaft assembly to the flap gearbox and tighten the retaining nut. After each of the four shaft assemblies has been inspected, proceed to Step (26). (d) If the difference between the measurements is more than 0.50 inch, proceed to Step (20). (20) Remove the applicable access panel(s) from underneath the wing and locate the applicable inboard or outboard flap actuator shaft assembly (Ref. Chapter 6-50-00, WING ACCESS PANELS). CAUTION: Ensure that the flap actuator 90° drive assembly union does not loosen while loosening the flexible drive shaft retaining nut. (21) Remove safety wire from the retaining nut on the flap actuator and disconnect the flexible drive shaft from the flap actuator. (22) If the retainer is attached, proceed to Step (23). If the retainer is not attached, remove and replace the flap flexible shaft assembly. (23) Measure the distance from the end of the inner shaft to the closest edge of the retainer (Ref. Figure 203). (24) If the measurement is 0.78 inch to 0.87 inch, proceed to Step (25). If the measurement is less than 0.78 inch or more than 0.87 inch, remove and replace the flap flexible shaft assembly. NOTE: Do not exceed 75 inch-pounds torque when tightening the flexible drive shaft retaining nuts. (25) Connect the flexible drive shaft to the flap actuator and safety wire the retaining nut. (26) Remove the red tag from the flap control lever. (27) If the flexible drive shaft (3) for the right-hand inboard flap is replaced, rig the flap system (all flaps) (Ref. 27-50-05, FLAP SYSTEM RIGGING). Replacement of any other flexible drive shaft (3) requires the rigging of that particular flap, only (Ref. Figure 201). (28) Install the applicable access panel(s) underneath the wing (Ref. Chapter 6-50-00, WING ACCESS PANELS).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (29) Install the left floorboard panel 16E (Ref. Chapter 6-50-00, FLOOR ACCESS PANELS). (30) Install the center aisle floorboard 10 (UA-1 and After) or 11 (UB-1 and After and UC-1 and After) (Ref. Chapter 6-50-00, FLOOR ACCESS PANELS). (31) Install the center aisle passenger compartment carpet (Ref. Chapter 25-20-01, CARPET REMOVAL AND INSTALLATION).

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D

3

4 1

2

A

3

DETAIL

B

1. NUT 2. FLAP GEAR BOX AND MOTOR 3. FLAP FLEXIBLE DRIVESHAFT 4. SAFETY WIRE 5. FLAP ACTUATOR 6. COTTER PIN 7. NUT 8. WASHER 3 9. MOUNT SCREWS

9

3

CL

2 1

3

4

4

3

5 6 7 8

6 7 8

DETAIL DETAIL

LWS 142.0

C

D

CL

LWS 56.0

RWS 47.83

B

RWS 56.0

RWS 142.0

FS 326.00

C

TOP VIEW

DETAIL

A

Figure 201 Flap Gearbox and Motor Installation

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UC27B 045664AB.AI

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 202 Flap Flexible Drive Shaft Installation

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RETAINING NUT 2 INCHES INBOARD SIDE 2 1/2 INCHES OUTBOARD SIDE

INNER SHAFT

CLOSEST EDGE OF FERRULE

INNER SHAFT RETAINER CLOSEST EDGE OF RETAINER

0.78 INCH TO 0.87 INCH MEASURE THE DISTANCE FROM THE END OF THE INNER SHAFT TO THE CLOSEST EDGE OF THE RETAINER. UC27B 050465AA.AI

Figure 203 Inspection of Retainer

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FLIGHT CONTROLS FLAP ACTUATORS MAINTENANCE PRACTICES

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1. OUTBOARD A. Removal (1) Apply external electrical power to the airplane (Ref. Chapter 24-40-00). (2) Select the BATT switch to the ON position. (3) Select the EXT PWR switch to the EXT PWR position. (4) Lower flaps to the full down position, then open the FLAP IND & CONTROL circuit breaker located on the right circuit breaker panel Zone 246 (Ref. Chapter 6-40-00, ZONE DIAGRAM). (5) Select the EXT PWR switch to the OFF position. (6) Select the BATT switch to the OFF position. (7) Disconnect external electrical power from the airplane (Ref. Chapter 24-40-00). (8) Attach a red tag to the flap control lever with the words “Do Not Operate, Maintenance In Progress”. (9) Remove wing access panels 8, 27, 28, 35 and 36 (UA-1 and After; UB-1 and After) or 17 (UC-1 and After) as necessary to access effected outboard flap actuator (Ref. Chapter 6-50-00, WING ACCESS PANELS). (10) Remove the bolt (2), washers (1) and nut (5) securing the flap actuator (3) to the brackets (6) on the leading edge of the flap (7). Slide the end of the actuator (3) free of the brackets (6) and remove the bushing (8) from the end of the actuator (3) (Ref. Figure 201). CAUTION: Ensure that the flap actuator 90° drive assembly (16) does not loosen while loosening the flexible drive shaft retaining nut (14). (11) Remove safety wire (13) from the retaining nut (14) on the flap actuator (3) and disconnect the flexible drive shaft (9) from the flap actuator (3). (12) Remove nuts (11), washers (10), bolts (12) and bushings (15) from the flap actuator (3). (13) Remove the flap actuator (3) from the airplane. (14) On airplanes serials UC-169 thru UC-174, remove the cable rub block (2) from the end of the flap actuator (1) housing (Ref. Figure 203).

B. Installation NOTE: If a new or overhauled actuator is being installed, make sure that the tape placed over the vent hole for shipping has been removed.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (1) On airplanes serials UC-169 thru UC-174, install the cable rub block (2) to the end of the actuator (1) housing. Position the cable rub block (2) to be parallel to the aileron cable with the flaps full down. If contact occurs, the aileron cable must contact the phenolic surface of the cable rub block (2) and not contact the metal housing of the outboard flap actuator (1) (Ref. Figure 203). (2) Position the actuator (3) with its vent hole up and install bushings (15), bolts (12), washers (10) and nuts (11) (Ref. Figure 201). NOTE: Do not exceed 75 inch-pounds torque when tightening the flexible drive shaft (9) retaining nuts (14). (3) Connect the flexible drive shaft (9) to the flap actuator (3) and safety wire (13) the retaining nut (14). (4) On the outboard actuators (1), loosen the locknut (4) and turn the adjustment nut (3) all the way in (toward the spring (2)), until reaching the end of the threads. Then back the adjustment nut (3) off four turns and tighten the lock nut (4) against it (Ref. Figure 202). (5) Make sure that the flap actuator (3) is extended, then install the bushing (8) in the end of the actuator (3) (Ref. Figure 201). (6) Adjust the left and right outboard flap actuator rod ends so that the flap trailing edges are 0.24 to 0.50 inch above the trailing edge of the adjacent inboard flap. (7) Align the end of the actuator (3) with holes in the brackets (6) on the flap (7) and install the bolt (2), washer (1) and nut (5) securing the end of the actuator (3) in place. (8) Remove the red tag from the flap control lever. CAUTION: With assistance, observe flap movement and be prepared to open the FLAP IND & CONTROL circuit breaker to stop the flaps. (9) Perform the FLAP OPERATIONAL CHECK procedure (Ref. 27-50-05). (10) If a new or overhauled actuator (3) is installed, lift lightly on the trailing edge of the flap (7) while running the flaps through a complete extension-retraction cycle. There should be no roughness or indication of binding in the actuator. (11) Install wing access panels 8, 27, 28, 35 and 36 (UA-1 and After; UB-1 and After) or 17 (UC-1 and After) as necessary (Ref. Chapter 6-50-00, WING ACCESS PANELS).

2. INBOARD A. Removal (1) Apply external electrical power to the airplane (Ref. Chapter 24-40-00). (2) Select the BATT switch to the ON position. (3) Select the EXT PWR switch to the EXT PWR position. (4) Lower flaps to the full down position, then open the FLAP IND & CONTROL circuit breaker located on the right circuit breaker panel Zone 246 (Ref. Chapter 6-40-00, ZONE DIAGRAM). (5) Select the EXT PWR switch to the OFF position. Page 202 Nov 1/09

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (6) Select the BATT switch to the OFF position. (7) Disconnect external electrical power from the airplane (Ref. Chapter 24-40-00). (8) Attach a red tag to the flap control lever with the words “Do Not Operate, Maintenance In Progress”. (9) Remove wing access panels 8, 27, 28, 35 and 36 (UA-1 and After; UB-1 and After) or 17 (UC-1 and After) as necessary to access effected flap actuator (Ref. Chapter 6-50-00, WING ACCESS PANELS). (10) Remove the bolt (2), washers (1) and nut (5) securing the flap actuator (3) to the brackets (6) on the leading edge of the flap (7). Slide the end of the actuator (3) free of the brackets (6) and remove the bushing (8) from the end of the actuator (3) (Ref. Figure 201). CAUTION: Ensure that the flap actuator 90° drive assembly (16) does not loosen while loosening the flexible drive shaft retaining nut (14). (11) Through access panels 8, 27, 28, 35 and 36 (UA-1 and After; UB-1 and After) or 17 (UC-1 and After) as necessary, remove safety wire (13) from the retaining nut (14) on the flap actuator (3) and disconnect the flexible drive shaft (9) from the flap actuator (3). (12) Through access panels 8, 27, 28, 35 and 36 (UA-1 and After; UB-1 and After) or 17 (UC-1 and After) as necessary, remove nuts (11), washers (10), bushings (15) and bolts (12) from the flap actuator (3). (13) Rotate flap actuator 90° to the left so that the flexible drive shaft connector is facing down. (14) Remove the flap actuator (3) from the airplane.

B. Installation NOTE: If a new or overhauled actuator is being installed, make sure that the tape placed over the vent hole for shipping has been removed. (1) Using super glue, glue the nut (11) to the washer (10) and to the flap actuator mount hole bushing (15) (Ref. Figure 201). (2) Install safety wire to the 90° drive assembly (16) for later installation to retaining nut. (3) Carefully position the flap actuator (3) with its vent hole up into the mount bracket and install bolts (12). NOTE: Do not exceed 75 inch-pounds torque when tightening the flexible drive shaft (9) retaining nuts (14). (4) Connect the flexible drive shaft (9) to the flap actuator (3) and safety wire (13) the retaining nut (14). (5) Make sure that the flap actuator (3) is extended, then install the bushing (8) in the end of the actuator (3). (6) Align the end of the actuator (3) with holes in the brackets (6) on the flap (7) and install the bolt (2), washer (1) and nut (5) securing the end of the actuator (3) in place.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (7) Remove the red tag from the flap control lever. (8) Perform the FLAP RIGGING procedure (Ref. 27-50-05). (9) If a new or overhauled actuator (3) is installed, lift lightly on the trailing edge of the flap (7) while running the flaps through a complete extension-retraction cycle. There should be no roughness or indication of binding in the actuator. (10) Install wing access panels 8, 27, 28, 35 and 36 (UA-1 and After; UB-1 and After) or 17 (UC-1 and After) as necessary (Ref. Chapter 6-50-00, WING ACCESS PANELS).

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1. WASHER 2. BOLT 3. FLAP ACTUATOR 4. WING 5. NUT 6. FLAP ACTUATOR BRACKET 7. FLAP 8. BUSHING

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9. FLEXIBLE DRIVESHAFT 10. WASHERS 11. NUTS 12. BOLTS 13. SAFETY WIRE 14. RETAINING NUT 15. BUSHING 16. 90ø DRIVE ASSEMBLY

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Figure 201 Flap Actuator Installation

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Figure 202 Flap Actuator Adjustment

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1. FLAP ACTUATOR 2. RUB BLOCK 3. FLAP ACTUATOR ROD

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Figure 203 Flap Actuator Rub Block

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

FLIGHT CONTROLS FLAP CONTROL SYSTEM MAINTENANCE PRACTICES

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1. PROCEDURES A. Rigging - Using Travel Boards (1) Perform the FLAP TRAVEL BOARD INSTALLATION procedure on the right inboard flap (Ref. 27-00-02). (2) Connect external electrical power to the airplane. (3) Select the BATT switch to the ON position. (4) Lower flaps to the full down position, then open the FLAP IND & CONTROL circuit breaker located on the right circuit breaker panel Zone 246 (Ref. Chapter 6-40-00, ZONE DIAGRAM). (5) Select the BATT switch to the OFF position. (6) Disconnect external electrical power from the airplane. (7) On airplanes equipped with flap bumpers: (a) Loosen the jam nut on each of the four flap bumpers located forward of the flaps at WS 33.00, 108.94, 138.00 and 188.00 on each wing. (b) Turn the bumpers inward to provide clearance from the flaps when they are in the up position. (8) On each of the inboard actuators (1), check that the adjustment nut (3) is all the way in toward the spring (2) at the end of the threads. If adjustment is required, loosen the locknut (4) and turn the adjustment nut (3) all the way in (toward the spring (2)) until reaching the end of the threads. Then tighten the locknut (4) against the adjustment nut (3) (Ref. Figure 201, Sheet 1). (9) On each of the outboard actuators (1), check the position of the adjustment nut (3). There must be 3 to 5 threads showing (6) on the spring side of the adjustment nut. If adjustment is required, loosen the locknut (4) and turn the adjustment nut (3) all the way in (toward the spring (2)) until reaching the end of the threads. Then back the adjustment nut (3) off four turns and tighten the locknut (4) against it (Ref. Figure 201, Sheet 2). (10) Disconnect the flap actuator rod end (3) from each of the four flap attach brackets (6) by removing the nuts (5), washers (1), bushing (8) and bolts (2) (Ref. Figure 207). (11) Retract each flap actuator by rotating the rod end (3) inward by hand until it will not interfere with the flap, when raised by hand, to the full up position. (12) Raise and lower each flap by hand to assure freedom of movement with no binding or unusual noise. (13) Locate the turnbuckle (1) between the flap limit switches and the right inboard flap. Measure the turnbuckle length between the centers of the holes in the clevis (2) on each end of the turnbuckle (1). Measurement must be 3.92 ± 0.06 inches. Remove and adjust the turnbuckle (1) as required (Ref. Figure 206).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (14) Remove the flap travel limit switches access panel 1 (Ref. Chapter 6-50-00, WING ACCESS PANELS). NOTE: Use the RH flap travel board reading as the master when making all adjustments. (15) With assistance and by positioning the right inboard flap by hand, adjust the flap limit switches by performing the following Steps: (a) Raise the right inboard flap slightly above 33° then slowly lower flap to 35 +1/ -2° and check that the S6 switch (7) activates by hearing an audible clicking noise. If the S6 switch (7) is not activated, use an allen-head wrench and adjust the S6 switch cam (6) to activate the S6 switch (7) at the 35 +1/ -2° flap position (Ref. Figure 202). (b) Raise the right inboard flap from 35 +1°/ -2° to 19° to 21° and check that the approach flap limit S4 switch (9) activates by hearing an audible clicking noise. If the S4 switch (9) is not activated, use an allen-head wrench and adjust the S4 switch cam (4) to activate the S4 switch (9) at the 19° to 21° flap position. (c) Raise the right inboard flap to a position less than 17° then lower the flap to 19° to 21° and check that the S5 switch (8) activates by hearing an audible clicking noise. If the S5 switch (8) is not activated, use an allen-head wrench and adjust the S5 switch cam (5) to activate the S5 switch (8) at the 19° to 21° flap position. NOTE: There must be a 0.5° to 2° flap angle (as read on travel board) difference between the activation of the S4 switch (9) and the S5 switch (8) after both switch cams have had their final adjustments. Maintaining this 0.5° to 2° flap angle difference will reduce the possibility of the flaps making small uncommanded movements up and down (hunting) in this flap position. The S4 switch (9) must activate before the S5 switch (8) deactivates when raising the flaps from full down. The S5 switch (8) must activate before the S4 switch (9) deactivates when lowering the flaps from full up. Example: The S5 switch (8) activates at 19°. The S4 switch (9) activates at 20.5°. This is the activation difference of 0.5° to 2°. (d) Raise the right inboard flap to the full UP position and check that the S1 switch (12) activates at the 1° ± 0.5° trailing edge down flap position by hearing an audible clicking noise. If the S1 switch (12) is not activated, use an allen-head wrench and adjust the S1 switch cam (1) to activate the S1 switch (12) at the 1 ± 0.5° trailing edge down flap position. Although the allowable range is 0.5° to 1.5°, set as close to the 1° reading as possible to provide additional flap clearance at the flap cove. (e) Lower the right inboard flap from the full up position to 9° to 11° and check that the S3 switch (10) activates by hearing an audible clicking noise. If the S3 switch (10) is not activated, use an allen-head wrench and adjust the S3 switch cam (3) to activate the S3 switch (10) at the 9° to 11° flap position. (f) Raise the right inboard flap from the full down position to 9° to 11° and check that the S2 switch (11) activates by hearing an audible clicking noise. If the S2 switch (11) is not activated, use an allen-head wrench and adjust the S2 switch cam (2) to activate the S2 switch (11) at the 9° to 11° flap position.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL NOTE: There must be a 0.5° to 2° flap angle (as read on travel board) difference between the activation of the S2 switch (11) and the S3 switch (10) after both switch cams have had their final adjustments. Maintaining this 0.5° to 2° flap angle difference will reduce the possibility of the flaps making small uncommanded movements up and down (hunting) in this flap position. The S2 switch (11) must activate before the S3 switch (10) deactivates when raising the flaps from full down. The S3 switch (10) must activate before the S2 switch (11) deactivates when lowering the flaps from full up. Example: The S3 switch (10) activates at 9°. The S2 switch (11) activates at 10.5°. This is the activation difference of 0.5° to 2°. (g) Lower the right inboard flap to 35 +1°/ -2° and check that the S6 switch (7) activates by hearing an audible clicking noise. If the S6 switch (7) is not activated, use an allen-head wrench and adjust the S6 switch cam (6) to activate the S6 switch (7) at the 35 +1°/ -2° flap position. NOTE: Do not alter any limit switch settings to compensate for a faulty indicating system. The indicating system will be checked and adjusted at a later Step. CAUTION: Ensure the flaps do not bottom out in the flap track at the full down position. Be sure all actuators are attached to their respective flaps before proceeding or structural damage may occur. (16) Position the right inboard flap by hand to the full down position. Adjust the inboard flap actuator rod end (3) to line up with the flap attach bracket (6) holes (Ref. Figure 207). (17) Check the travel board for a reading of 35 +1/ -2°, and adjust as necessary. Record the position of the right inboard flap angle as read on the travel board. (18) Install bushing (8), bolt (2), washers (1) and nut (5) attaching the flap actuator to the right inboard flap. Ensure the bolt head faces outboard. (19) Perform the FLAP TRAVEL BOARD INSTALLATION procedure on the left inboard flap (Ref. 27-00-02). (20) Position the left inboard flap by hand to the full down position. Adjust the inboard flap actuator rod end to line up with the flap attach bracket holes until the left flap is within 0.5° of the right flap and adjust as necessary. (21) Install bushing (8), bolt (2), washers (1) and nut (5) attaching the flap actuator to the left inboard flap. Ensure the bolt head faces outboard (Ref. Figure 207). (22) Adjust the left and right outboard flap (2) actuator rod ends so that the flap trailing edges are 0.24 to 0.50 inch above the trailing edge of the adjacent inboard flap (1). Install bushing, bolt, washers and nut attaching the flap actuator to the outboard flaps (Ref. Figure 204). NOTE: A fine adjustment of the flap can be made by extending the flaps to the full down position and disconnecting the cable at the flap motor gearbox. Rotate the cable 90° in either direction. Access the flap motor gearbox by removing the aft spar ramp and floor access panel 10 (Ref. Chapter 6-50-00, FLOORBOARD ACCESS PANELS).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (23) Adjust the flap safety switch (asymmetric switch) (Ref. Chapter 27-50-06, FLAP SAFETY SWITCH ADJUSTMENT). Do not perform FLAP SYSTEM OPERATIONAL CHECK at the end of the procedure at this time. NOTE: For all remaining Steps, move the flap travel board between the left and right inboard flaps as necessary. (24) Connect external electrical power to the airplane. (25) Select the BATT switch to the ON position. (26) Select the EXT PWR switch to the ON position. CAUTION: With assistance, observe flap movement and be prepared to open the FLAP IND & CONTROL circuit breaker to stop the flaps. (27) Close the FLAP IND & CONTROL circuit breaker. (28) Select the FLAP lever to the UP 0° position. Momentarily bump the FLAP IND & CONTROL circuit breaker to raise the flaps to 0°. Verify the flaps stop at 1° ± 0.5° trailing edge down. Although the allowable range is 0.5° to 1.5°, set as close to the 1° reading as possible to provide additional flap clearance at the flap cove. If required, adjust the S1 switch cam (1) to activate the S1 switch (12) at this position (Ref. Figure 202). Turn off external power and airplane battery before making any adjustments. (29) Verify that the right and left outboard flap (2) trailing edges are even ± 0.15 inch with their respective inboard flap (1) trailing edges when the flaps are in the full up position (Ref. Figure 204). (30) If adjustment is required, extend the flaps to the full down position and adjust per Step (22). Turn off external power and airplane battery before making any adjustments. Then repeat Step (29). (31) With the FLAP lever at the UP position, select the FLAP lever to the TAKEOFF position. Verify the flaps stop at 9° to 11° on the travel board. If required, adjust the S3 switch cam (3) to activate the S3 switch (10) at this position. Record the position of the right inboard flap angle as read on the travel board (Ref. Figure 202). Turn off external power and airplane battery before making any adjustments. (32) Select the FLAP lever to the APPROACH position. Verify the flaps stop at 19° to 21° on the travel board. If required, adjust the S5 switch cam (5) to activate the S5 switch (8) at this position. Record the position of the right inboard flap angle as read on the travel board (Ref. Figure 202). Turn off external power and airplane battery before making any adjustments. CAUTION: Ensure the flaps do not bottom out in the flap track at the full down position. (33) Select the FLAP lever to the LANDING position and verify the flaps stop at 35 +1/ -2° on the travel board. If required, adjust the S6 switch cam (6) to activate the S6 switch (7) at this position. Turn off external power and airplane battery before making any adjustments. Record the position of the right inboard flap angle as read on the travel board. (34) Verify that the right and left outboard flap (2) trailing edges are 0.24 to 0.50 inch above the trailing edge of the adjacent inboard flap (1) (Ref. Figure 204).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL NOTE: There must be a minimum of 0.5° flap angle difference between the positions of the S4 switch cam (4) and S5 switch cam (5) after both switch cams have had their final adjustments (Ref. Figure 202). There must be a minimum of 0.5° flap angle difference between the positions of the S2 switch cam (2) and S3 switch cam (3) after both switch cams have had their final adjustments. (35) Select the FLAP lever to the APPROACH position. Verify the flaps stop at 19° to 21° on the travel board. If required, adjust the S4 switch cam (4) to activate the S4 switch (9) at this position. Turn off external power and airplane battery before making any adjustments. (36) Select the FLAP lever to the TAKEOFF position. Verify the flaps stop at 9° to 11° on the travel board. If required, adjust the S2 switch cam (2) to activate the S2 switch (11) at this position. Turn off external power and airplane battery before making any adjustments. (37) Select the FLAP lever to the UP position. Verify the flaps stop at 1 ± 0.5° trailing edge down. Although the allowable range is 0.5° to 1.5°, set as close to the 1° reading as possible to provide additional flap clearance at the flap cove. NOTE: Do not adjust the flap-travel limit switches for an incorrect flap position indicating system. CAUTION: Do not exceed more than one complete flap duty cycle every 5 minutes. (38) If the flap position indicator on the pedestal shows the flaps being a position other than the reading on the travelboard, perform the following Steps (Ref. Figure 203): NOTE: If a discrepancy exists between the UP and LANDING readings on the flap position indicator, adjust the clevis length (L) to correct approximately one half of the discrepancy. Adjust the transmitter arm for maximum flap position indicator accuracy at the TAKEOFF and APPROACH positions, with reasonable accuracy at UP and LANDING positions. Verify that the flap position indicator is in the Approach arc at the 19° to 21° flap position as indicated on the travel boards. (a) Note the position of the flap position indicator needle at each flap position. (b) Cycle the flaps from UP to TAKEOFF, APPROACH to LANDING then cycle the flaps from LANDING to APPROACH, TAKEOFF to UP. (c) To decrease the difference between the UP and LANDING indications, loosen the upper nut (5) and tighten the lower nut (12) to increase the length (L). (d) To increase the difference between the UP and LANDING indications, loosen the lower nut (12) and tighten the upper nut (5) to decrease the length (L). (e) To rotate the indicator needle reading, loosen nut (3) and bolt (2) on arm (6) and rotate the shaft (4) until the indicator pointer aligns correctly. (f) Cycle flaps to check adjustments. Repeat adjustment procedure as necessary. (39) On airplanes equipped with flap bumpers, the following Steps must be accomplished: (a) Lower the flaps to the full down position.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (b) Loosen the jam nut on each of the four flap bumpers. NOTE: Use marking material on the bumper surface to determine when the flap makes contact. CAUTION: Do not exceed more than one complete flap duty cycle every 5 minutes. (c) Extend each flap bumper until contact with the flap surface can be made with the flaps in the normal (up) position. (d) Extend each flap bumper an additional turn after flap contact. (e) Tighten the jam nuts. (40) Cycle the flaps to check operation. (41) Select EXT PWR switch to the OFF position. (42) Select the BATT switch to the OFF position. (43) Disconnect electrical power from airplane. (44) Remove the travel boards. (45) Install all access panels. (a) Remove the old sealant from the access panel 1 (Ref. Chapter 6-50-00, WING ACCESS PANELS) and from around the flap switch wing opening. (b) Apply sealant (19, Table 1, Chapter 91-00-00) around the flap switch access panel and install access panel 1 (Ref. Chapter 6-50-00, WING ACCESS PANELS) to secure access to the flap position switches.

B. Rigging - Using Protractors (1) Connect external electrical power to the airplane. (2) Select the BATT switch to the ON position. (3) Select EXT PWR switch to the ON position. (4) Lower flaps to the full down position, then disengage the FLAP IND & CONTROL circuit breaker located on the right circuit breaker panel Zone 246 (Ref. Chapter 6-40-00, ZONE DIAGRAM). (5) Select the BATT switch to the OFF position. (6) Disconnect external electrical power from the airplane. (7) On airplanes equipped with flap bumpers: (a) Loosen the jam nut on each of the four flap bumpers located forward of the flaps at WS 33.00, 108.94, 138.00 and 188.00 on each wing. (b) Turn the bumpers inward to provide clearance from the flaps when they are in the up position.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (8) On each of the inboard actuators (1), check that the adjustment nut (3) is all the way in toward the spring (2) at the end of the threads. If adjustment is required, loosen the locknut (4) and turn the adjustment nut (3) all the way in (toward the spring (2)) until reaching the end of the threads. Then tighten the locknut (4) against the adjustment nut (3) (Ref. Figure 201, Sheet 1). (9) On each of the outboard actuators (1), check the position of the adjustment nut (3). There must be 3 to 5 threads showing (6) on the spring side of the adjustment nut. If adjustment is required, loosen the locknut (4) and turn the adjustment nut (3) all the way in (toward the spring (2)) until reaching the end of the threads. Then back the adjustment nut (3) off four turns and tighten the locknut (4) against it (Ref. Figure 201, Sheet 2). (10) Disconnect the flap actuator rod end (3) from each of the four flap attach brackets (6) by removing the nuts (5), washers (1), bushings (8) and bolts (2) (Ref. Figure 207). (11) Retract each flap actuator by rotating the rod end (3) inward by hand until it will not interfere with the flap, when raised by hand, to the full up position. (12) Raise and lower each flap by hand to assure freedom of movement with no binding or unusual noise. (13) Locate the turnbuckle (1) between the flap limit switches and the right inboard flap. Measure the turnbuckle length between the centers of the holes in the clevis (2) on each end of the turnbuckle (1). Measurement must be 3.92 ± 0.06 inches. Remove and adjust the turnbuckle (1) as required (Ref. Figure 206). (14) Remove the flap travel limit switches access panel 1 (Ref. Chapter 6-50-00, WING ACCESS PANELS). (15) Attach protractors (1) (5, Table 1, 27-00-00) to the upper surfaces of the left and right inboard flaps (2) at approximately WS 60.52, perpendicular to the length of the flaps (Ref. Figure 205). NOTE: Use the RH protractor reading as the master when making all adjustments. (16) Perform the FLAP SYSTEM RIGGING - UP POSITION CHECK to determine the up position of the flaps. This check must be performed if the flap, flap track or the fuselage fairing has been changed since the last time this procedure was performed. If this has been performed and the measurement has been recorded, this result may be used. NOTE: The protractor must be adjusted to agree with the travel board measurement. All flap position requirements are from 0° as would be measured by a travel board. (17) Raise the right inboard flap to the full up position. (a) If the flap can be positioned to 0° per the travel board FLAP SYSTEM RIGGING - UP POSITION CHECK, position the flap at 0° and set the protractor to read 0° in agreement with the travel board. (b) If the flap cannot be raised to 0°, but can be raised to a position between 0° and 1.5° per the travel board, then perform either Step (17) (b) 1 or (17) (b) 2: 1 Adjust the protractor to agree with the travel board measurement. 2 Set the protractor to 0° and then add the travel board measurement to the protractor readings while rigging.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (18) With assistance and by positioning the right inboard flap by hand, adjust the flap limit switches by performing the following Steps: (a) Raise the right inboard flap slightly above 33° then slowly lower flap to 35 +1/ -2° and check that the S6 switch (7) activates by hearing an audible clicking noise. If the S6 switch (7) is not activated, use an allen-head wrench and adjust the S6 switch cam (6) to activate the S6 switch (7) at the 35 +1/ -2° flap position (Ref. Figure 202). (b) Raise the right inboard flap from 35 +1/ -2° to 19° to 21° and check that the approach flap limit S4 switch (9) activates by hearing an audible clicking noise. If the S4 switch (9) is not activated, use an allen-head wrench and adjust the S4 switch cam (4) to activate the S4 switch (9) at the 19° to 21° flap position. (c) Raise the right inboard flap to a position less than 17° then lower the flap 19° to 21° and check that the S5 switch (8) activates by hearing an audible clicking noise. If the S5 switch (8) is not activated, use an allen-head wrench and adjust the S5 switch cam (5) to activate the S5 switch (8) at the 19° to 21° flap position. NOTE: There must be a 0.5° to 2° flap angle (as read on the protractor) difference between the activation of the S4 switch (9) and the S5 switch (8) after both switch cams have had their final adjustments. Maintaining this 0.5° to 2° flap angle difference will reduce the possibility of the flaps making small uncommanded movements up and down (hunting) in this flap position. The S4 switch (9) must activate before the S5 switch (8) deactivates when raising the flaps from full down. The S5 switch (8) must activate before the S4 switch (9) deactivates when lowering the flaps from full up. Example: The S5 switch (8) activates at 19°. The S4 switch (9) deactivates at 20.5°. This is the activation difference of 0.5° to 2°. (d) Raise the right inboard flap to the full up position and check that the S1 switch (12) activates at the 0 ± 1.5° trailing edge down flap position by hearing an audible clicking noise. If the S1 switch (12) is not activated, use an allen-head wrench and adjust the S1 switch cam (1) to activate the S1 switch (12) at the 0 ± 1.5° trailing edge down flap position. Although the allowable range is 0.5° to 1.5°, set as close to the 1° reading as possible to provide additional flap clearance at the flap cove. (e) Lower the right inboard flap from the full up position to 9° to 11° and check that the S3 switch (10) activates by hearing an audible clicking noise. If the S3 switch (10) is not activated, use an allen-head wrench and adjust the S3 switch cam (3) to activate the S3 switch (10) at the 9° to 11° flap position. (f) Raise the right inboard flap from the full down position to 9° to 11° and check that the S2 switch (11) activates by hearing an audible clicking noise. If the S2 switch (11) is not activated, use an allen-head wrench and adjust the S2 switch cam (2) to activate the S2 switch (11) at the 9° to 11° flap position. NOTE: There must be a 0.5° to 2° flap angle (as read on protractor) difference between the activation of the S2 switch (11) and the S3 switch (10) after both switch cams have had their final adjustments. Maintaining this 0.5° to 2° flap angle difference will reduce the possibility of the flaps making small uncommanded movements up and down (hunting) in this flap position.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL The S2 switch (11) must activate before the S3 switch (10) deactivates when raising the flaps from full down. The S3 switch (10) must activate before the S2 switch (11) deactivates when lowering the flaps from full up. Example: The S3 switch (10) activates at 9°. The S2 switch (11) deactivates at 10.5°. This is the activation difference of 0.5° to 2°. (g) Lower the right inboard flap to 35 +1/ -2° and check that the S6 switch (7) activates by hearing an audible clicking noise. If the S6 switch (7) is not activated, use an allen-head wrench and adjust the S6 switch cam (6) to activate the S6 switch (7) at the 35 +1/ -2° flap position. NOTE: Do not alter any limit switch settings to compensate for a faulty indicating system. The indicating system will be checked and adjusted at a later Step. CAUTION: Ensure the flaps do not bottom out in the flap track at the full down position. Be sure all actuators are attached to their respective flaps before proceeding or structural damage may occur. (19) Position the right inboard flap by hand to the full down position. Adjust the inboard flap actuator rod end (3) to line up with the flap attach bracket (6) holes (Ref. Figure 207). (20) Check the protractor for a reading of 35 +1/ -2°, and adjust as necessary. Record the position of the right inboard flap angle as read on the protractor. (21) Install bushing (8), bolt (2), washer (1) and nut (5) attaching the flap actuator to the right inboard flap. Ensure the bolt head faces outboard. (22) Using the FLAP SYSTEM RIGGING - UP POSITION CHECK measurement recorded for the left inboard flap, the protractor must be adjusted to agree with the travel board measurement. All flap position requirements are from 0° as would be measured by a travel board. (23) Raise the left inboard flap to the full up position. (a) If the flap can be positioned to 0° per the travel board FLAP SYSTEM RIGGING - UP POSITION CHECK, position the flap at 0° and set the protractor to read 0° in agreement with the travel board. (b) If the flap cannot be raised to 0°, but can be raised to a position between 0° and 1.5° per the travel board, then perform either Step (23) (b) 1 or (23) (b) 2: 1 Adjust the protractor to agree with the travel board measurement. 2 Set the protractor to 0° and then add the travel board measurement to the protractor readings while rigging. (24) Position the left inboard flap by hand to the full down position. Adjust the inboard flap actuator rod end (3) to line up with the flap attach bracket (6) holes until the left flap is within 0.5° of the right flap (Ref. Figure 207). (25) Install bushing (8), bolt (2), washer (1) and nut (5) attaching the flap actuator to the left inboard flap. Ensure the bolt head faces outboard.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (26) Adjust the left and right outboard flap (2) actuator rod ends so that the flap trailing edges are 0.24 to 0.50 inch above the trailing edge of the adjacent inboard flap (1). Install bushing, bolt, washer and nut attaching the flap actuator to the outboard flaps (2) (Ref. Figure 204). NOTE: A fine adjustment of the flap can be made by extending the flaps to the full down position and disconnecting the cable at the flap motor gearbox. Rotate the cable 90° in either direction. Access the flap motor gearbox by removing the aft spar ramp and floor access panel 10 (Ref. Chapter 6-50-00, FLOORBOARD ACCESS PANELS). (27) Adjust the flap safety switch (asymmetric switch) (Ref. Chapter 27-50-06, FLAP SAFETY SWITCH ADJUSTMENT). Do not perform FLAP SYSTEM OPERATIONAL CHECK at the end of the procedure at this time. NOTE: For all remaining Steps, move the protractor between the left and right inboard flaps as necessary. (28) Connect external electrical power to the airplane. (29) Select the BATT switch to the ON position. (30) Select EXT PWR switch to the ON position. CAUTION: With assistance, observe flap movement and be prepared to open the FLAP IND & CONTROL circuit breaker to stop the flaps. (31) Close the FLAP IND & CONTROL circuit breaker. (32) Select the FLAP lever to the UP 0° position. Momentarily bump the FLAP IND & CONTROL circuit breaker to raise the flaps to 0°. Verify the flaps stop at 1 ± 0.5° trailing edge down. Although the allowable range is 0.5° to 1.5°, set as close to the 1° reading as possible to provide additional flap clearance at the flap cove. If required, adjust the S1 switch cam (1) to activate the S1 switch (12) at this position (Ref. Figure 202). Turn off external power and airplane battery before making any adjustments. (33) Verify that the right and left outboard flap (2) trailing edges are even ± 0.15 inch with their respective inboard flap (1) trailing edges when the flaps are in the full up position (Ref. Figure 204). (34) If adjustment is required, extend flaps to the full down position and adjust per Step (26) and then repeat Step (33). (35) With the FLAP lever at the UP position, select the FLAP lever to the TAKEOFF position. Verify the flaps stop at 9° to 11° on the protractor. If required, adjust the S3 switch cam (3) to activate the S3 switch (10) at this position. Turn off external power and airplane battery before making any adjustments. Record the position of the right inboard flap angle as read on the protractor (Ref. Figure 202). (36) Select the FLAP lever to the APPROACH position. Verify the flaps stop at 19° to 21° on the protractor. If required, adjust the S5 switch cam (5) to activate the S5 switch (8) at this position. Turn off external power and airplane battery before making any adjustments. Record the position of the right inboard flap angle as read on the protractor (Ref. Figure 202). CAUTION: Ensure the flaps do not bottom out in the flap track at the full down position.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (37) Select the FLAP lever to the LANDING position and verify the flaps stop at 35 +1/ -2° on the protractor. If required, adjust the S6 switch cam (6) to activate the S6 switch (7) at this position. Turn off external power and airplane battery before making any adjustments (38) Verify that the right and left outboard flap (2) trailing edges are 0.24 to 0.50 inch above the trailing edge of the adjacent inboard flap (1) (Ref. Figure 204). NOTE: There must be a minimum of 0.5° flap angle difference between the positions of the S4 switch cam (4) and S5 switch cam (5) after both switch cams have had their final adjustments (Ref. Figure 202). There must be a minimum of 0.5° flap angle difference between the positions of the S2 switch cam (2) and S3 switch cam (3) after both switch cams have had their final adjustments. (39) Select the FLAP lever to the APPROACH position. Verify the flaps stop at 19° to 21° on the protractor. If required, adjust the S4 switch cam (4) to activate the S4 switch (9) at this position. Turn off external power and airplane battery before making any adjustments. (40) Select the FLAP lever to the TAKEOFF position. Verify the flaps stop at 9° to 11° on the protractor. If required, adjust the S2 switch cam (2) to activate the S2 switch (11) at this position. Turn off external power and airplane battery before making any adjustments. (41) Select the FLAP lever to the UP position. Verify the flaps stop at 1 ± 0.5° trailing edge down. Although the allowable range is 0.5° to 1.5°, set as close to the 1° reading as possible to provide additional flap clearance at the flap cove. Turn off external power and airplane battery before making any adjustments. NOTE: Do not adjust the flap-travel limit switches for an incorrect flap position indicating system. CAUTION: Do not exceed more than one complete flap duty cycle every 5 minutes. (42) If the flap position indicator on the pedestal shows the flaps being a position other than the reading on the protractor, perform the following Steps (Ref. Figure 203): NOTE: If a discrepancy exists between the UP and LANDING readings on the flap position indicator, adjust the clevis length (L) to correct approximately one half of the discrepancy. Adjust the transmitter arm for maximum flap position indicator accuracy at the TAKEOFF and APPROACH positions, with reasonable accuracy at UP and LANDING positions. Verify that the flap position indicator is in the Approach arc at the 19° to 21° flap position as indicated on the protractor. (a) Note the position of the flap position indicator needle at each flap position. (b) Cycle the flaps from UP to TAKEOFF to APPROACH to LANDING then cycle the flaps from LANDING to APPROACH to TAKEOFF to UP. (c) To decrease the difference between the UP and LANDING indications, loosen the upper nut (5) and tighten the lower nut (12) to increase the length (L). (d) To increase the difference between the UP and LANDING indications, loosen the lower nut (12) and tighten the upper nut (5) to decrease the length (L).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (e) To rotate the indicator needle reading, loosen nut (3) and bolt (2) on arm (6) and rotate the shaft (4) until the indicator pointer aligns correctly. (f) Cycle flaps to check adjustments. Repeat adjustment procedure as necessary. (43) On airplanes equipped with flap bumpers, the following Steps must be accomplished: (a) Lower the flaps to the full down position. (b) Loosen the jam nut on each of the four flap bumpers. NOTE: Use marking material on the bumper surface to determine when the flap makes contact. CAUTION: Do not exceed more than one complete flap duty cycle every 5 minutes. (c) Extend each flap bumper until contact with the flap surface can be made with the flaps in the normal (up) position. (d) Extend each flap bumper an additional turn after flap contact. (e) Tighten the jam nuts. (44) Cycle the flaps to check operation. (45) Select EXT PWR switch to the OFF position. (46) Select the BATT switch to the OFF position. (47) Disconnect electrical power from airplane. (48) Remove the protractors. (49) Install all access plates. (a) Remove the old sealant from the access panel 1 (Ref. Chapter 6-50-00, WING ACCESS PANELS) and from around the flap switch wing opening. (b) Apply sealant (19, Table 1, Chapter 91-00-00) around the flap switch panel and install the panel 1 (Ref. Chapter 6-50-00, WING ACCESS PANELS) to secure access to the flap position switches.

C. Rigging - Up Position Check WARNING: This procedure is intended to measure the up position of the inboard flaps and is used when performing FLAP SYSTEM RIGGING - USING PROTRACTORS. The flaps must be rigged if this check is performed. (1) Connect external electrical power to the airplane. (2) Select the BATT switch to the ON position. (3) Lower flaps to the full down position, then open the FLAP IND & CONTROL circuit breaker located on the right circuit breaker panel Zone 246 (Ref. Chapter 6-40-00, ZONE DIAGRAM). (4) Select the BATT switch to the OFF position.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (5) Disconnect external electrical power from the airplane. (6) On airplanes equipped with flap bumpers: (a) Loosen the jam nut on each of the two flap bumpers located forward of the inboard flaps at WS 33.00 and 108.94 on each wing. (b) Turn the bumpers inward to provide clearance from the inboard flaps when the flaps are in the up position. (7) Disconnect the flap actuator rod end (3) from right inboard flap attach bracket (6) by removing the nut (5), washers (1), bushing (8) and bolt (2) (Ref. Figure 207). (8) Retract the flap actuator by rotating the rod end (3) inward by hand until it will not interfere with the flap, when raised by hand, to the full up position. (9) Raise and lower the flap by hand to assure freedom of movement with no binding or unusual noise. (10) Perform the FLAP TRAVEL BOARD INSTALLATION procedure on the right inboard flap (Ref. 27-00-02). (11) Raise the right inboard flap to the full up position. Measure the position of the flap using the travelboard. (a) If the flap can be raised to a position between 0° and 1.5° per the travel board, record the results for use when performing FLAP SYSTEM RIGGING - USING PROTRACTORS procedure. (b) If the flap cannot be raised to 1.5°, identify the interference and repair as required to allow the flap to be raised to an up position from 0° to 1.5° per the travel board. (12) Remove the travel board from the right inboard flap. (13) Disconnect the flap actuator rod end (3) from left inboard flap attach bracket (6) by removing the nut (5), washers (1), bushing (8) and bolt (2) (Ref. Figure 207). (14) Retract the flap actuator by rotating the rod end (3) inward by hand until it will not interfere with the flap, when raised by hand, to the full up position. (15) Raise and lower the flap by hand to assure freedom of movement with no binding or unusual noise. (16) Perform the FLAP TRAVEL BOARD INSTALLATION procedure on the left inboard flap (Ref. 27-00-02). (17) Raise the left inboard flap to the full up position. Measure the position of the flap using the travelboard. (a) If the flap can be raised to a position between 0° and 1.5° per the travel board, record the results for use when performing FLAP SYSTEM RIGGING - USING PROTRACTORS procedures. (b) If the flap cannot be raised to 1.5°, identify the interference and repair as required to allow the flap to be raised to an up position from 0° to 1.5° per the travel board.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (18) Perform either the FLAP SYSTEM RIGGING WITH TRAVEL BOARD procedure or the FLAP SYSTEM RIGGING WITH PROTRACTOR procedure in this section.

D. Flap Adjustment to Correct a Wing Heavy Condition NOTE: Most wing heavy conditions can be traced to improperly rigged ailerons or flaps. If further adjustment is needed, one outboard flap may be rigged slightly down. This will result in additional drag on the airplane. Use the AILERON GROUND ADJUSTABLE TRIM TAB ADJUSTMENT procedure (Ref. 27-10-00) before using this procedure as it will make a larger correction. Rigging one outboard flap slightly down will make only a small correction. (1) Flight Check for Roll. (a) Set the rudder tab to eliminate any skidding. (b) Set the aileron tab to zero (ignore aileron forces). (c) Match trailing edge of ailerons to trailing edge of flaps on both wings (ailerons neutral). (d) Check for rolling to right or left and note conditions. (2) Maintenance action to correct a wing heavy condition. (a) Perform the AILERON FUNCTIONAL CHECK procedure to ensure proper aileron system operation prior to performing the following procedure (Ref. 27-10-03). (b) Perform the FLAP SYSTEM FUNCTIONAL CHECK procedure to ensure proper flap alignment prior to performing the following procedure (Ref. 27-50-05). NOTE: The following procedure could cause an out-of-tolerance split flap indication. (c) Select the BATT switch to the ON position. (d) Lower the flaps to the full down position. (e) If slight and further adjustment is needed, the outboard flap actuator rod end may be extended 1/2 to 1 turn maximum on the wing that is heavy. 1 Disconnect the flap actuator rod end (3) from the flap attach bracket (6) by removing the nut (5), washer (1), bushing (8) and bolt (2) (Ref. Figure 207). CAUTION: Do not exceed the 1 turn maximum extension on the outboard flap actuator. 2 Rotate the rod end (3) outward 1/2 to 1 turn Maximum. 3 Install bushing (8), bolt (2), washers (1) and nut (5) attaching the flap actuator to the flap. Ensure the bolt head faces outboard. (f) Loosen the jam nut on the two flap bumpers for the flap that was adjusted in Step (2) (e). NOTE: Use marking material on the bumper surface to determine when the flap makes contact.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (g) Raise the flaps to the normal (UP) position. (h) Extend each flap bumper until it makes contact with the flap surface. CAUTION: Do not exceed more than one complete flap duty cycle every 5 minutes. (i) Cycle the flaps and check for contact with the flap bumpers. Repeat Steps (2) (h) as necessary. (j) Extend each flap bumper an additional turn after flap contact is made. (k) Tighten the jam nuts. CAUTION: Do not exceed more than one complete flap duty cycle every 5 minutes. (l) Cycle the flaps to check for proper operation. (m) Select the BATT switch to the OFF position.

E. Functional Check (1) Connect external electrical power to the airplane. (2) Perform the FLAP TRAVEL BOARD INSTALLATION procedure (Ref. 27-00-02) or attach protractors (1) (5, Table 1, 27-00-00) to the upper surfaces of the left and right inboard flaps (2) at WS 60.52, perpendicular to the length of the flaps (Ref. Figure 205). (3) Select the BATT switch to the ON position. CAUTION: Do not exceed more than one complete flap duty cycle every 5 minutes. (4) Perform the following Steps to check the travel of the flaps: (a) With the flaps at the full up position, move the flap control lever down to the TAKEOFF position. Verify the flaps move down and stop at 9° to 11° on the travel board or protractors with no unusual noise or binding. Verify the flap position indicator needle stops within the white TAKEOFF arc. (b) Move the flap control lever down to the APPROACH position. Verify the flaps move down and stop at 19° to 21° on the travel board or protractors with no unusual noise or binding. Verify the flap position indicator needle stops within the white APPROACH arc. (c) Move the flap control lever to the LANDING position. Verify the flaps move down and stop at 35 +1/ -2° on the travel board or protractors with no unusual noise or binding. Verify the flap position indicator needle stops within the LANDING position. (d) Confirm that the left and right outboard flap trailing edges are 0.24 to 0.50 inch above the trailing edge of the adjacent inboard flap (Ref. Figure 204). (e) Move the flap control lever up to the APPROACH position. Verify the flaps move up and stop at 19° to 21° on the travel board or protractors with no unusual noise or binding. Verify the flap position indicator needle stops within the white APPROACH arc.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (f) Move the flap control lever up to the TAKEOFF position. Verify the flaps move up and stop at 9° to 11° on the travel board or protractors with no unusual noise or binding. Verify the flap position indicator needle stops within the white TAKEOFF arc. (g) Move the flap control lever to the UP position. Verify the flaps move up and stop at 1 ± 0.5° trailing edge down with no unusual noise or binding. Verify the flap position indicator needle stops within the UP position. (h) Verify that the right and left outboard flap trailing edges are even ± 0.15 inch with their respective inboard flap trailing edges when the flaps are in the full up position (Ref. Figure 204). (5) If the flaps are not within tolerance perform the FLAP SYSTEM RIGGING procedure in this section. (6) Remove the flap travel board or protractors. (7) Select the BATT switch to the OFF position. (8) Select the BATT switch to the OFF position.

F. Operational Check (1) Connect external electrical power to the airplane. (2) Select the BATT switch to the ON position. CAUTION: Do not exceed more than one complete flap duty cycle every 5 minutes. (3) Perform the following Steps to check the travel of the flaps: (a) Move the flap control lever down to the TAKEOFF position. Verify the flaps move down and stop with no unusual noise or binding. Verify the flap position indicator needle stops within the white TAKEOFF arc. (b) Move the flap control lever down to the APPROACH position. Verify the flaps move down and stop with no unusual noise or binding. Verify the flap position indicator needle stops within the white APPROACH arc. (c) Move the flap control lever to the LANDING position. Verify the flaps move down and stop with no unusual noise or binding. Verify the flap position indicator needle stops within the LANDING position. (d) Move the flap control lever up to the APPROACH position. Verify the flaps move up and stop with no unusual noise or binding. Verify the flap position indicator needle stops within the white APPROACH arc. (e) Move the flap control lever up to the TAKEOFF position. Verify the flaps move up and stop with no unusual noise or binding. Verify the flap position indicator needle stops within the white TAKEOFF arc. (f) Move the flap control lever to the UP position. Verify the flaps move up and stop with no unusual noise or binding. Verify the flap position indicator needle stops within the UP position.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (4) Select the BATT switch to the OFF position. (5) Remove electrical power from the airplane.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 201 (Sheet 1 of 2) Flap Actuator Adjustment

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2. SPRING 3. ADJUSTMENT NUT 4. LOCK NUT 5. FLAP ACTUATOR ROD 6. THREADS

5 2

4

3

6

UC27B 060636AA.AI

Figure 201 (Sheet 2 of 2) Flap Actuator Adjustment

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1

2

3

6

5

4

A

VIEW LOOKING DOWN THROUGH OPEN PANEL FWD INBD

DETAIL

B 1. S1 CAM (FLAP UP LIMIT) 2. S2 CAM (TAKEOFF UP LIMIT) 3. S3 CAM (TAKEOFF DOWN LIMIT) 4. S4 CAM (APPROACH UP LIMIT) 5. S5 CAM (APPROACH DOWN LIMIT AND LANDING GEAR WARNING) 6. S6 CAM (DOWN LIMIT) 7. S6 SWITCH (FLAP DOWN LIMIT)

1

2

3

12

B

4

11

5

8. S5 SWITCH (FLAP APPROACH DOWN LIMIT AND LANDING GEAR WARNING) 9. S4 SWITCH (FLAP APPROACH UP LIMIT AND LANDING GEAR WARNING) 10. S3 SWITCH (FLAP TAKEOFF DOWN LIMIT) 11. S2 SWITCH (FLAP TAKEOFF UP LIMIT) 12. S1 SWITCH (FLAP UP LIMIT)

6

10

9

8

7

VIEW LOOKING UP DETAIL

A

Figure 202 Flap Limit Adjustment Cams

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1

2 3

A

4 5

B

6 12 11 L 10 9 8

1. FLAP POSITION TRANSMITTER 2. BOLT 3. NUT 4. SHAFT 5. UPPER NUT 6. ARM 7. LINK 8. COTTER PIN 9. CLEVIS 10. CLEVIS PIN 11. ADJUSTMENT SHAFT 12. LOWER NUT

7

DETAIL

A

UP

UP FLAPS

FLAPS APPROACH

UP

UP FLAPS

APPROACH

FLAPS APPROACH

APPROACH

LANDING

LANDING

LANDING

LANDING

BEFORE

AFTER

BEFORE

AFTER

ADJUST L DIMENSION

ADJUST SHAFT

DETAIL

B UC27B 044972AB.AI

Figure 203 Flap Position Transmitter

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A

1. INBOARD FLAP 2. OUTBOARD FLAP 3. WING

1 2

3

2 1 DETAIL

A

Figure 204 Outboard Flap Adjustment

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

A 1. DIGITAL PROTRACTOR 2. INBOARD FLAP

1

NO STEP

2

DETAIL

A UC27B 044841AA.AI

Figure 205 Digital Protractor Installation

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1. TURNBUCKLE 2. CLEVIS 3. SAFETY CLIPS

A

2

1

3

DETAIL

3

2

A UC27B 063118AA.AI

Figure 206 Turnbuckle Adjustment 4

5

3

6

7

1

A

2

8

OUTBOARD

RIGHT SIDE SHOWN LEFT SIDE OPPOSITE DETAIL

A

1. WASHER 2. BOLT 3. FLAP ACTUATOR ROD END 4. WING 5. NUT 6. FLAP ATTACH BRACKET 7. FLAP 8. BUSHING

Figure 207 Flap Attach Bracket

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

FLIGHT CONTROLS FLAP SAFETY SYSTEM MAINTENANCE PRACTICES

27-50-06 200200

1. FLAP SAFETY SWITCH (ASYMMETRIC) A. Removal (1) Apply external electrical power to the airplane (Ref. Chapter 24-40-00). (2) Select the BATT switch to the ON position. (3) Select the EXT PWR switch to the EXT PWR position. (4) Lower flaps to the full down position, then open the FLAP IND & CONTROL circuit breaker located on the right circuit breaker panel Zone 246 (Ref. Chapter 6-40-00, ZONE DIAGRAM). (5) Select the EXT PWR switch to the OFF position. (6) Select the BATT switch to the OFF position. (7) Disconnect external electrical power from the airplane (Ref. Chapter 24-40-00). (8) Attach a red tag to the flap control lever with the words “Do Not Operate, Maintenance In Progress”. NOTE: Before removing the parts in the next Step, note the number and position of the washers so that they will be properly installed (Ref. Figure 201 or 202). (9) Remove cover over the flap safety switch, if installed. (10) Remove screws (1), washers (2), spacers (4) and nuts (7) from the safety switch (5) and switch enclosure (6). (11) Remove switch enclosure (6). (12) Identify, tag and disconnect the wires from the safety switch (5). (13) Remove the safety switch (5) from the airplane and discard.

B. Installation NOTE: Ensure that the wires have enough length to allow for movement of the switch. (1) Remove tags and connect the wires to the safety switch. The red wire attaches to the “common terminal” and the white wire attaches to the “normally open” terminal. (2) Install switch enclosure (6) on the safety switch (5) (Ref. Figure 201 or 202). (3) Position safety switch (5) in place and install screws (1), washers (2), spacers (4) and nuts (7). NOTE: For initial adjustment, the safety switch (5) should be pushed to the upper edge of the elongated mounting holes. (4) Remove the red tag from the flap control lever.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (5) Close the FLAP IND & CONTROL circuit breaker located on the right circuit breaker panel Zone 246 (Ref. Chapter 6-40-00, ZONE DIAGRAM). (6) Perform the FLAP SAFETY SWITCH TEST procedure in this section.

C. Removal (With Kit No. 114-5057 Installed) (1) Apply external electrical power to the airplane (Ref. Chapter 24-40-00). (2) Select the BATT switch to the ON position. (3) Select the EXT PWR switch to the EXT PWR position. (4) Lower flaps to the full down position, then open the FLAP IND & CONTROL circuit breaker located on the right circuit breaker panel Zone 246 (Ref. Chapter 6-40-00, ZONE DIAGRAMS). (5) Select the EXT PWR switch to the OFF position. (6) Select the BATT switch to the OFF position. (7) Disconnect external electrical power from the airplane (Ref. Chapter 24-40-00). (8) Attach a red tag to the flap control lever with the words “Do Not Operate, Maintenance In Progress”. NOTE: Before removing the parts, note the number and position of the washers so that they will be properly installed. (9) Remove screws (7) and washers (8) from the flap safety switch cover (14) and remove cover (14) (Ref. Figure 203). (10) Remove screws (28), washers (27), from the safety switch assembly (13). (11) Identify, tag and disconnect the wires at the splice from the safety switch assembly (13). (12) Remove the safety switch assembly (13) from the airplane and discard.

D. Installation (With Kit No. 114-5057 Installed) NOTE: For initial adjustment, the safety switch (13) should be pushed to the upper edge of the elongated mounting holes. (1) Position safety switch assembly (13) in place. Apply loctite (161, Table 1, Chapter 91-00-00) to the threads of the screws and install screws (28) and washers (27) (Ref. Figure 203). NOTE: Ensure that the wires have enough length to allow for movement of the switch. (2) Route wires through the grommet (15) on the safety switch assembly cover (14) and splice wires to the harness per wiring diagram. (3) Close the FLAP IND & CONTROL circuit breaker located on the right circuit breaker panel Zone 246 (Ref. Chapter 6-40-00, ZONE DIAGRAMS). (4) Perform the FLAP SAFETY SWITCH TEST procedure in this section.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (5) Position the flap safety switch cover (14) over the flap safety switch assembly (13) and install screws (7) and washers (8). (6) Remove the red tag from the flap control lever.

E. Test (UA-1 and After; UB-1 and After; UC-1 thru UC-53) (1) Apply external electrical power to the airplane (Ref. Chapter 24-40-00). (2) Select the BATT switch to the ON position. (3) Select the EXT PWR switch to the EXT PWR position. (4) Lower flaps to the full down position, then open the FLAP IND & CONTROL circuit breaker located on the right circuit breaker panel Zone 246 (Ref. Chapter 6-40-00, ZONE DIAGRAM). (5) Disconnect either end of the RH outboard flap adjustable rod end (11) by removing arm attachment nut (8) and washer (9) (Ref. Figure 201). (6) Reposition arm (10) until roller cam (3) has allowed safety switch (5) roller to move off the cam lobe. WARNING: Should the flaps start to retract when performing the following Step, be prepared to immediately open the FLAP IND & CONTROL circuit breaker. (7) Move flap position switch to APPROACH, then reset the FLAP IND & CONTROL circuit breaker located on the right circuit breaker panel Zone 246 (Ref. Chapter 6-40-00, ZONE DIAGRAMS). The flaps should not retract. (8) If the flaps retract, check wiring and replace the flap safety switch (5) as required. Perform the FLAP SAFETY SWITCH REMOVAL and INSTALLATION procedures in this section and then repeat this test. (9) Ensure that the flaps are extended to the full down position. (10) Open the FLAP IND & CONTROL circuit breaker located on the right circuit breaker panel Zone 246 (Ref. Chapter 6-40-00, ZONE DIAGRAMS). (11) Connect RH outboard flap adjustable rod end (11) by installing arm attachment nut (8) and washer (9). (12) Disconnect either end of the LH outboard flap adjustable rod end (11) by removing arm attachment nut (8) and washer (9). (13) Position arm (10) until roller cam (3) has allowed safety switch (5) roller to move off the cam lobe. WARNING: Should the flaps start to retract when performing the following Step, be prepared to immediately open the FLAP IND & CONTROL circuit breaker. (14) Move the flap position switch to APPROACH, then reset the FLAP IND & CONTROL circuit breaker located on the right circuit breaker panel Zone 246 (Ref. Chapter 6-40-00, ZONE DIAGRAMS). The flaps should not retract. (15) If the flaps retract, check wiring and replace the flap safety switch (5) as required.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (16) Perform the FLAP SAFETY SWITCH REMOVAL and INSTALLATION procedures in this section and then repeat this test. (17) Ensure that the flaps are extended to the full down position. (18) Open the FLAP IND & CONTROL circuit breaker located on the right circuit breaker panel Zone 246 (Ref. Chapter 6-40-00, ZONE DIAGRAMS). (19) Connect LH outboard flap adjustable rod end (11) by installing arm attachment nut (8) and washer (9). (20) Perform the FLAP SAFETY SWITCH ADJUSTMENT procedure in the section.

F. Test (UC-54 and After) (1) Apply external electrical power to the airplane (Ref. Chapter 24-40-00). (2) Select the BATT switch to the ON position. (3) Select the EXT PWR switch to the EXT PWR position. (4) Lower flaps to the full down position, then open the FLAP IND & CONTROL circuit breaker located on the right circuit breaker panel Zone 246 (Ref. Chapter 6-40-00, ZONE DIAGRAMS). (5) Disconnect either end of the RH outboard flap adjustable rod end (13) by removing the arm attachment screw (14), spacer (12), washer (11), light washer (9) and nut (8) (Ref. Figure 202). (6) Position the arm (10) until the roller cam (3) has allowed the safety switch (5) roller to move off the cam lobe. WARNING: Should the flaps start to retract when performing the following Step, be prepared to immediately open the FLAP IND & CONTROL circuit breaker. (7) Move the flap position control to APPROACH, then reset the FLAP IND & CONTROL circuit breaker. The flaps should not retract. (8) If the flaps retract, check wiring and replace the flap safety switch (5) as required. Perform the FLAP SAFETY SWITCH REMOVAL and INSTALLATION procedures in this section and then repeat this test. (9) Ensure that the flaps are extended to the full down position. (10) Open the FLAP IND & CONTROL circuit breaker. (11) Connect the RH outboard flap adjustable rod end (13) by installing the arm attachment screw (14), spacer (12), washer (11), light washer (9) and nut (8). (12) Disconnect either end of the LH outboard flap adjustable rod end (13) by removing the arm attachment screw (14), spacer (12), washer (11), light washer (9) and nut (8). (13) Position the arm (10) until the roller cam (3) has allowed the safety switch (5) roller to move off the cam lobe. WARNING: Should the flaps start to retract when performing the following Step, be prepared to immediately open the FLAP IND & CONTROL circuit breaker.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (14) Move the flap position control to APPROACH, then reset the FLAP IND & CONTROL circuit breaker. The flaps should not retract. (15) If the flaps retract, check wiring and replace the flap safety switch (5) as required. (16) Perform the FLAP SAFETY SWITCH REMOVAL and INSTALLATION procedures in this section and then repeat this test. (17) Ensure that the flaps are extended to the full down position. (18) Connect the LH outboard flap adjustable rod end (13) by installing the arm attachment screw (14), spacer (12), washer (11), light washer (9) and nut (8). (19) Perform the FLAP SAFETY SWITCH ADJUSTMENT procedure in this section.

G. Adjustment (Without Kit No. 114-5057 Installed) NOTE: The safety switches are located in each wing between the inboard and outboard flaps (Ref. Figures 201 and 202). (1) Apply external electrical power to the airplane (Ref. Chapter 24-40-00). (2) Select the BATT switch to the ON position. (3) Select the EXT PWR switch to the EXT PWR position. (4) Extend the flaps to the full down position. (5) Select the EXT PWR switch to the OFF position. (6) Select the BATT switch to the OFF position. (7) Disconnect external electrical power from the airplane (Ref. Chapter 24-40-00). (8) Attach a red tag to the flap control lever with the words “Do Not Operate, Maintenance In Progress”. (9) The inboard flap link (without Kit No. 118-4013) (19) is a non adjustable solid link. If Kit No 118-4013 is installed, check that the inboard flap rod end linkage is a length of 2.75 inches between the centers of the rod end bores. (10) The centerline of the pointed end of the cam must align with the switch roller. The cam may have a groove line to aid in this inspection. (11) If the cam does not align with the switch roller, adjust the outboard flap rod end linkage (13). Begin with a length of 2.50 inches between the centers of the rod end bores, then adjust the length to align the cam with the switch roller. (12) Install a digital protractor (5, Table 1, 27-00-00) to the inboard edge of the outboard flap. CAUTION: Do not run flaps or change flap actuator adjustment with the actuator disconnected from the flap. If the flap system is operated with the actuator disconnected or the actuator length is changed, the flap system must be rigged.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (13) Zero the protractor and disconnect the flap actuator from the left outboard flap. The outboard flap will move down to the end of the flap track. CAUTION: Ensure that the actuators will not strike or damage the wing or flaps while moving the outboard flaps during adjustment procedures. (14) Slowly raise the outboard flap by hand while reading the protractor. When the switch opens (audible click), note the protractor reading. This may need to be repeated as necessary to get an accurate measurement. (15) The correct safety switch setting stops all flap travel when the positions of the inboard and outboard flaps have a difference of 3° to 6°. (16) If the difference is less than 3°, move the safety switch up in the slots of the inboard arm. If the difference is more than 6°, move the safety switch down in the slots of the arm. (17) Connect the flap actuator to the left outboard flap. (18) Repeat Steps (10) thru (17) for the right side flaps. (19) After all flap actuators are connected, perform the FLAP SYSTEM OPERATIONAL CHECK procedure (Ref. 27-50-05). (20) Remove the red tag from the flap lever.

H. Adjustment (With Kit No. 114-5057 Installed) The safety switch assemblies are located in each wing between the inboard and outboard flaps. (1) Apply external electrical power to the airplane (Ref. Chapter 24-40-00). (2) Select the BATT switch to the ON position. (3) Select the EXT PWR switch to the EXT PWR position. (4) Extend the flaps to the full down position. (5) Attach a red tag to the flap control lever with the words “Do Not Operate, Maintenance In Progress”. (6) Remove screws (7) and washers (8) from the flap safety switch cover (14) and remove cover (14) (Ref. Figure 203). (7) The inboard flap link (without Kit No. 118-4013) (20) is a non adjustable solid link. If Kit No. 118-4013 is installed, check that the inboard flap rod end linkage is a length of 2.75 inches between the centers of the rod end bores. (8) The centerline of the pointed end of the cam must align with the switch roller. The cam may have a groove line to aid in this inspection. (9) If the cam does not align with the switch roller, adjust the outboard flap rod end linkage (2). Begin with a length of 2.50 inches between the centers of the rod end bores, then adjust the length to align the cam with the switch roller (Ref. Figure 203). (10) Install a digital protractor (5, Table 1, 27-00-00) to the inboard edge of the outboard flap.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (11) Raise flaps to the APPROACH position. (12) Select the EXT PWR switch to the OFF position. (13) Select the BATT switch to the OFF position. (14) Disconnect external electrical power from the airplane (Ref. Chapter 24-40-00). CAUTION: Do not run flaps or change flap actuator adjustment with the actuator disconnected from the flap. If the flap system is operated with the actuator disconnected or the actuator length is changed, the flap system must be rigged. (15) Zero the protractor and disconnect the flap actuator from the left outboard flap. The outboard flap will move down to the end of the flap track. CAUTION: Ensure that the actuators will not strike or damage the wing or flaps while moving the outboard flaps during adjustment procedures. (16) Raise the outboard flap by hand until the protractor reads approximately 0° and listen for an audible clicking noise. The switch roller must be on the pointed end of the cam. (17) Slowly lower the outboard flap by hand while reading the protractor. When the switch opens (audible click), note the protractor reading. This may need to be repeated as necessary to get an accurate measurement. (18) Raise the outboard flap by hand until the protractor reads approximately 0° and listen for an audible clicking noise. The switch roller must be on the pointed end of the cam. (19) Slowly raise the outboard flap by hand while reading the protractor. When the switch opens (audible click), note the protractor reading. This may need to be repeated as necessary to get an accurate measurement. (20) The correct safety switch setting stops all flap travel when the positions of the inboard and outboard flaps have a difference of 5° to 9° in either direction. A higher setting is preferred to prevent nuisance tripping of the switch. (21) If the safety switch setting is 5° to 9°, proceed to Step (25). If not, perform the following Steps: (22) Add the protractor readings. If the number of degrees while lowering the flap plus the number of degrees while raising the flap is not 10° to 18° then adjust the switch in the slots and repeat Steps (15) through (22). NOTE: Lowering the switch in the slots decreases the total degrees. Raising the switch in the slots increases the total degrees. (23) If the number of degrees while lowering the flap does not equal the number of degrees while raising the flap, the outboard rod end length must be adjusted and Steps (15) through (22) must be repeated. NOTE: Lengthening the outboard rod end link decreases flap up travel and increases flap down travel. Shortening the outboard rod end link increases flap up travel and decreases flap down travel.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (24) Example: 10.8° was measured while lowering the outboard flap. 7.4° was measured while raising the outboard flap. Total = 10.8 + 7.4 = 18.2°. The switch must be moved down in the slots until the total is less than 18°. After moving the switch, 10.0° is measured while lowering, 5.8° is measured while raising. Total = 10.0 + 5.8 = 15.8°. The switch is correctly positioned since the total is 18° or less, but the 10° exceeds the 9° requirement. Adjust the outboard rod end linkage until both readings are within 5° to 9°. (25) Connect the flap actuator to the left outboard flap. (26) Apply external electrical power to the airplane (Ref. Chapter 24-40-00). (27) Select the BATT switch to the ON position. (28) Select the EXT PWR switch to the EXT PWR position. (29) Extend the flaps to the full down position. (30) Select the EXT PWR switch to the OFF position. (31) Select the BATT switch to the OFF position. (32) Disconnect external electrical power from the airplane (Ref. Chapter 24-40-00). (33) Position the flap safety switch cover (14) over the flap safety switch assembly (13) and install screws (7) and washers (8) (Ref. Figure 203). (34) Repeat Steps (6) thru (25) for the right side flaps. (35) After all flap actuators are connected, perform the FLAP SYSTEM OPERATIONAL CHECK procedure (Ref. 27-50-05). (36) Remove red tag from flap lever.

I. Hub Lubrication (Without Kit No. 129-5046 Installed) The following procedure is for the disassembly, lubrication and assembly of the Flap Safety Switch Hub Assembly without Kit No. 129-5046 installed. (1) Apply external electrical power to the airplane (Ref. Chapter 24-40-00). (2) Select the BATT switch to the ON position. (3) Select the EXT PWR switch to the EXT PWR position. (4) Extend the flaps to the full down position. (5) Select the EXT PWR switch to the OFF position. (6) Select the BATT switch to the OFF position. (7) Attach a red tag to the flap control lever with the words “Do Not Operate, Maintenance In Progress”. (8) Remove nut (3), washers (2) and bolt (1) attaching the outboard link assembly (4) to the outboard arm (7) (Ref. Figure 204).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (9) Remove nut (17), washers (16) and bolt (15) attaching the inboard link assembly (18) to the inboard arm (14). (10) Remove nut (5), washer (6) and bolt (8) from the outboard arm (7). NOTE: Ensure that the wires have enough length to allow for movement of the switch. CAUTION: Secure loose switch to airplane to avoid damage to the wires during maintenance. (11) Remove washers (9) and slide switch shaft (13) from the inboard side of the hub (11) keeping washers (12) with the switch shaft (13). (12) Remove washers (12) from the switch shaft (13). (13) Clean the switch shaft (13) and the inner part of the hub (11) with alcohol, solvent or equivalent. (14) Inspect the switch shaft (13) and the inner parts of the hub (11) for corrosion or damage and replace as necessary. (15) Apply grease (23 or 61, Table 1, Chapter 91-00-00) to the switch shaft (13) and fill the inside of the hub (11) with grease (23 or 61, Table 1, Chapter 91-00-00). (16) Install washers (12) to the switch shaft (13). (17) Position the switch shaft (13) on the inboard side of the hub (11) and slide the switch shaft (13) into the hub (11). (18) Install washers (9) to the outboard end of the switch shaft (13). (19) Install the outboard arm (7) to the switch shaft (13) and secure using bolt (8), washer (6) and nut (5). (20) Install bolt (1), washers (2) and nut (3) attaching the outboard link assembly (4) to the outboard arm (7). (21) Install bolt (15), washers (16) and nut (17) attaching the inboard link assembly (18) to the inboard arm (14). (22) Remove the red tag from the flap lever. (23) Perform the appropriate FLAP SAFETY SWITCH ADJUSTMENT procedure in this section.

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Figure 201 Flap Safety Switch Installation (LH Wing Shown, RH Wing Opposite (UA-3; UB-1 and After; UC-1 thru UC-53)

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Figure 202 Flap Safety Switch Installation (LH Wing Shown, RH Wing Opposite) (UC-54 and After)

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A

1

5 2

1. SCREWS 2. LINK 3. SPACERS 4. WASHERS 5. ARM 6. NUTS 7. SCREWS 8. WASHERS 9. COVER ARM ASSEMBLY 10. WAVE WASHER 11. BOLT 12. CAM 13. SWITCH ASSEMBLY 14. COVER 15. GROMMET

3 4 4 6 30 7 8

9

29 7 10

16. WASHER 8 17. NUT 18. BUSHINGS 19. NUTS 20. LINK * 21. WAVE WASHER 22. WASHER 23. INBOARD FLAP BRACKET 24. BOLT 25. WAVE WASHER 26. BOLT 27. WASHER 28. SCREWS 29. LIGHT WASHERS 30. OUTBOARD FLAP BRACKET

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14

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18 21

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LINK MAY HAVE AN ADJUSTABLE ROD END (WITH KIT NO. 118-4013 INSTALLED).

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TYPICAL CONFIGURATION (LEFT INSTALLATION SHOWN) ON SOME SERIALS HARDWARE MAY VARY

Figure 203 Flap Safety Switch Installation (LH Wing Shown, RH Wing Opposite) Page 212 Nov 1/09

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1. BOLT 2. WASHERS 3. NUT 4. OUTBOARD LINKAGE 5. NUT 6. WASHER 7. OUTBOARD ARM 8. BOLT 9. WASHERS 10. BUSHINGS 11. HUB 12. WASHERS 13. SWITCH SHAFT 14. INBOARD ARM

15. BOLT 16. WASHERS 17. NUT 18. INBOARD LINKAGE 19. SPACER * 20. LINKAGE *

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* UA-1 THRU UC-174 WITH KIT 118-4013 INSTALLED

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Figure 204 Flap Safety Switch Hub Lubrication (RH Wing Shown, LH Wing Opposite)

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FLIGHT CONTROLS FLAP POSITION SWITCHES MAINTENANCE PRACTICES

27-50-07 200200

1. PROCEDURES A. Removal (1) Remove pedestal covers to gain access to the switches. (2) Remove and mark electrical connector(s) from the switch or switches (3 and 5 or 7 and 8) to be removed. (3) Remove two screws (4 or 9) and nuts (6) holding switches (3 and 5 or 7 and 8) in place (Ref. Figure 202). (4) Remove the switches (3 and 5 or 7 and 8).

B. Installation (1) Put switch or switches (3 and 5 or 7 and 8) in place and insert two screws and nuts then tighten (Ref. Figure 202). (2) Perform FLAP POSITION SWITCH CHECK OUT AND RIGGING procedure below. (3) Install electrical connector(s) removed previously. Ensure the plugs are connected to the same switches they were removed from. (4) Install pedestal covers removed previously.

C. Check Out and Rigging (1) Place the flap handle in each position (Up, Take-off, Approach and Land) starting with the up position. (2) Using a good quality ohmmeter check continuity at each switch (3, 5, 7 and 8) for each flap handle position (Ref. Figure 202). Ensure ohmmeter readings match the readings in the Switch Logic Chart in Figure 201. (3) If the readings do not match, adjustment of the cam (2, 10 or 11) for the offending switch is necessary. NOTE: There are two set screws on each of the three adjustable cams. Ensure both set screws are loose before attempting to adjust the cam. (4) Loosen the set screws (1) on the cam (2, 10 or 11) then rotate the cam to achieve the desired result. (5) Tighten the set screws (1) then test the switch (3, 5, 7 or 8) opening and closing to ensure proper operation in both directions. Adjust the cam (2, 10 or 11) as necessary between tests. NOTE: When all adjustments and tests are completed apply a small amount of Loctite to the set screws (1) loosened during adjustment.

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D. Rigging Requirements NOTE: For initial adjustment, the safety switch (13) should be pushed to the upper edge of the elongated mounting holes. There is no rigging required for the Take-off Flap Position Switch (5), since this is accomplished by the correct assembly of parts as shown in Section B-B (Ref. Figure 202). The Landing Gear Warning Flap Position Switch (3) must make contact in the normally closed position with the detent roller (13) on the crest between the DOWN and APPROACH position detents on the cam (12) as shown in Section A-A. The Approach Flap Position Switch (7) must make contact in the normally closed position with the detent roller (13) on the crest between the APPROACH and TAKE-OFF position detents on the cam (12) as shown in Section C-C. The Full Down Flap Position Switch (8) must make contact in the normally closed position with the detent roller (13) on the crest between the DOWN and APPROACH position detents on the cam (12) as shown in Section D-D.

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Figure 201 Switch Logic Chart

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12 1. SET SCREWS (2 PLACES EACH CAM, EXCEPT TAKE-OFF CAM) 2. CAM (LANDING GEAR WARNING) 3. SWITCH, FLAP POSITION (LANDING GEAR WARNING) 4. SCREWS, SWITCH MOUNTING (2 PLACES) 5. SWITCH, FLAP POSITION (TAKE-OFF) 6. NUTS, SWITCH MOUNTING (4 PLACES) 7. SWITCH, FLAP POSITION (APPROACH) 8. SWITCH, FLAP POSITION (FULL DOWN) 9. SCREWS, SWITCH MOUNTING (2 PLACES) 10. CAM (FULL DOWN) 11. CAM (APPROACH) 12. CAM (TAKE-OFF) 13. DETENT ROLLER

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2 NC

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Figure 202 Flap Position Switches

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

FLIGHT CONTROLS GUST LOCKS AND DAMPENERS MAINTENANCE PRACTICES

27-70-00 200200

1. CONTROL LOCKS WARNING: The flight control gust locks provided by Hawker Beechcraft Corporation for its products are in compliance with federal regulations to provide an unmistakable warning to the pilot when the lock is engaged. When necessary or desirable to use flight control gust locks, use only the flight control gust lock assembly specified by Hawker Beechcraft Corporation for that particular airplane. When a flight control gust lock assembly is used, the lock must be correctly and fully installed, including the rudder pedal lock and throttle control lock. The control lock consists of a U-shaped clamp and two pins connected by a chain. The pins lock the primary flight controls and the U-shaped clamp fits around the engine power control levers and serves to warn the pilot not to start the engines with the control locks installed. It is important that the locks be installed or removed together to preclude the possibility of an attempt to taxi or fly the airplane with the power levers released and the pins still installed in the flight controls.

A. Inspection (1) Inspect the flight control gust lock to determine if it is the correct part number. (a) Measure the length of the chain between point A at the Rudder Lock Pin (3) that installs into the floor and point B at the U-clamp (2) that installs over the Engine Controls (1) (Ref. Figure 201). (b) If the dimension is 43 inches ± one chain link, the correct part number is installed. If the length does not check, replace with the correct part number control lock (29, Table 7, Chapter 91-00-00 or refer to the IPC). (2) Perform CONTROL LOCK INSTALLATION procedure. (3) Check the alignment of the Rudder Lock Pin (3) that installs into the floor. Try to move the rudder pedals to verify the rudder pin locks the rudder systems. (4) Check the alignment of the Control Column Lock Pin (4) that installs into the control column. Try to move the pilot’s control wheel to verify the pin locks the elevator and aileron control systems. (5) If the control lock must be removed at this time, perform CONTROL LOCK REMOVAL procedure.

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B. Removal WARNING: Before starting the engines, remove the control locks. Remove the control locks before towing the airplane. If towed with a tug while the rudder lock is installed, serious damage to the steering linkage can result. (1) Remove the Rudder Lock Pin (3). (2) Remove the Control Column Lock pin (4). (3) Remove the U-clamp (2) from the Engine Controls (1).

C. Installation (1) Position the U-clamp (2) around the engine controls (1) (Ref. Figure 201). NOTE: The holes are aligned when the control wheel is fully forward and rotated approximately 15° to the left. (2) Move the control column as necessary to align the holes, then insert the Control Column Lock Pin (4). (3) Insert Rudder Lock Pin (3) through the hole provided in the floor aft of the rudder pedals. The rudder pedals must be centered to align the hole in the rudder bellcrank with the hole in the floor. The Rudder Lock Pin (3) is then inserted until the flange is resting against the floor. This will prevent any rudder movement.

1. ENGINE CONTROLS 2. U-CLAMP 3. RUDDER LOCK PIN 4. CONTROL COLUMN LOCK PIN

1

4

GUST CONTROL SURFACE LOCK

POINT B

2

4 POINT A 3

3

Figure 201 Control Lock Installation

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CHAPTER 28 - FUEL TABLE OF CONTENTS SUBJECT

PAGE

FUEL SYSTEM (UA-1 AND AFTER; UB-1 AND AFTER) 28-00-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Fuel Storage and Distribution . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Fuel Quantity Indicating . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2 Fuel Flow Indicating . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2 Low Fuel Quantity Warning System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Servicing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Air-Maze Fuel Filter Cleaning - Primary Method . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Initial Cleaning - Filter Pack Element . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Final Cleaning - Filter Pack Element . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Air-Maze Fuel Filter Cleaning - Secondary Method . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202 Initial Cleaning - Secondary Method - Filter Pack Element . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202 Final Cleaning - Secondary Method - Filter Pack Element . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202

FUEL SYSTEM (UC-1 AND AFTER) 28-01-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Fuel Storage and Distribution . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Fuel Purge System (EPA) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Fuel Quantity Indicating . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2 Fuel Flow Indicating . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2 Low Fuel Quantity Warning System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Servicing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201

FUEL STORAGE (UA-1 AND AFTER; UB-1 AND AFTER) 28-10-00 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Fuel Cells . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Wing Outboard Leading Edge Fuel Cell . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Wing Inboard Leading Edge Fuel Cell . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Aft Outboard Wing Center Fuel Cell . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Aft Inboard Wing Fuel Cell . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Center Section Fuel Cell . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

28-CONTENTS

201 201 201 201 202 203 203 204 205 205 206 206 206 207 209 209 209

Page 1 Nov 1/13

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

CHAPTER 28 - FUEL TABLE OF CONTENTS (CONTINUED) SUBJECT PAGE Fuel Supply Collector Tank . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 211 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 211 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 212 Leakage Checks and Repairs . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 214 Leakage Checks and Repairs (Integral Fuel Cells) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 214 Fuel Cell Leakage Test - Bladder-Type Fuel Cells . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 216 Fuel Cell Repair . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 219 Fuel Cell Storage . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 219

ANTISIPHON VALVE (UA-1 AND AFTER; UB-1 AND AFTER) 28-10-01 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 201 201

FUEL STORAGE (UC-1 AND AFTER) 28-11-00 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Leakage Checks and Repairs (Integral Fuel Cells) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Fuel System Tank . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Collector Tank Area . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Main Fuel Tanks, WS 124 thru 130 (L&R) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Fuel Drain Valve . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Maintenance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 201 204 204 204 205 207 207

ANTISIPHON VALVE (UC-1 AND AFTER) 28-11-01 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 201 201

ALTERNATE TANK INSPECTION METHOD (UC-1 AND AFTER) 28-12-00 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Radiographic Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Preparation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Personnel & Safety . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Test Equipment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Film . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Film/X-ray Tube Placement and Exposure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Indication/Evaluation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Completion . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

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201 201 201 202 202 202 203 203 203

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FUEL DISTRIBUTION (UA-1 AND AFTER; UB-1 AND AFTER) 28-20-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1

FUEL FILTERS AND SCREENS (UA-1 AND AFTER; UB-1 AND AFTER) 28-20-02 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Servicing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Cleaning Fuel Filters . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Engine-Driven Fuel Pump Screens and Filters . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 201 201 201

FUEL PUMPS (UA-1 AND AFTER; UB-1 AND AFTER) 28-20-03 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Electric Fuel Boost Pump . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Jet Transfer Pump . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Primary Fuel Jet Pump . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 201 201 201 201 202 202 202 202

FUEL VALVES (UA-1 AND AFTER; UB-1 AND AFTER) 28-20-05 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Cross-Transfer Valve . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Firewall Shutoff Valves . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Functional Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 201 201 201 201 202 202

FUEL DISTRIBUTION (UC-1 AND AFTER) 28-21-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Access Panel . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201

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FUEL FITTINGS (UC-1 AND AFTER) 28-21-01 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Fuel Fittings . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201

FUEL FILTERS AND SCREENS (UC-1 AND AFTER) 28-21-02 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Servicing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Fuel Filter Element . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal and Cleaning (Aircraft Porous Media) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation (Aircraft Porous Media) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal and Cleaning (Air Maze) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation (Air Maze) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Engine-Driven Fuel Pump Screens and Filters . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Air-Maze Fuel Filter . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Cleaning - Primary Method . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Initial Cleaning - Filter Pack Element . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Final Cleaning - Filter Pack Element . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Cleaning - Secondary Method . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Initial Cleaning - Secondary Method - Filter Pack Element . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Final Cleaning - Secondary Method - Filter Pack Element . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 201 201 201 201 202 202 202 205 205 205 205 206 206 206

FUEL PUMPS (UC-1 AND AFTER) 28-21-03 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Electric Fuel Boost Pump . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Transfer Jet Pump . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Main Tank Jet Pump . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Auxiliary Fuel Transfer Pump . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 201 201 202 202 202 202 202 203 203 203 203

FUEL MANIFOLDS (UC-1 AND AFTER) 28-21-04 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

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PAGE

FUEL VALVES (UC-1 AND AFTER) 28-21-05 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Filter Fuel Shutoff Valve . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Firewall Fuel Shutoff Valve . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Functional Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Cross-Transfer Valve . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 201 201 201 201 202 202 203 203 203

LH AND RH FUEL LINES (UC-1 AND AFTER) 28-21-06 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Corrosion Prevention For Chafed Fuel Lines . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 201 202

FUEL QUANTITY INDICATING (UA-1 AND AFTER; UB-1 AND AFTER) 28-40-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Capacitance Fuel Quantity Indicating System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Fuel Quantity Capacitance Probes . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Low Fuel Quantity Warning System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2 Low Fuel Quantity Sensors . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3 Troubleshooting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 101 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 101 System Insulation Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 101 Probe Insulation Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 102 System and Probe Capacitance Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 103 Setting Capacitance Simulation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 105 Indicator Linearity Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 106 Indicator Calibration . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 107 Empty Tanks Calibration (Preferred Procedure) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 107 Full Tanks Calibration (Alternate Procedure Only) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 108 Fuel Quantity Probe Bench Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 109 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Fuel Probe - Inboard Leading Edge and Inboard Aft Fuel Cells . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Installation - Collector Tank . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Fuel Probe - Integral (Wet Wing) Fuel Cell . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202 Fuel Probe - Center Section . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 203 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 203

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CHAPTER 28 - FUEL TABLE OF CONTENTS (CONTINUED) SUBJECT PAGE Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 203 Fuel Probe - Collector Tank . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 204 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 204 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 204

FUEL LEVEL SENSORS 28-40-01 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Functional Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Fuel Quantity Indicator . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal and Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 201 201 202 202 202

FUEL QUANTITY INDICATING (UC-1 AND AFTER) 28-41-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Capacitance Fuel Quantity Indicating System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Low Fuel Quantity Warning System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Troubleshooting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 101 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 101 System Insulation Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 101 Probe Insulation Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 102 System Capacitance Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 106 Probe Capacitance Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 107 Setting Capacitance Simulation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 110 Indicator Linearity Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 110 Indicator Calibration . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 111 Empty Tanks Calibration (Preferred Procedure) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 111 Full Tanks Calibration (Alternate Procedure Only) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 113 Fuel Quantity Probe Bench Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 114 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Fuel Probe . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201

FUEL LEVEL AND LOW FUEL QUANTITY SENSORS (UC-1 AND AFTER) 28-41-01 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Low Fuel Quantity Sensor . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Fuel Level Sensor . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Functional Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Fuel Quantity Indicator . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal and Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

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List of Effective Pages CH-SE-SU

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Page 1 Nov 1/13

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

FUEL FUEL SYSTEM (UA-1 AND AFTER; UB-1 AND AFTER) DESCRIPTION AND OPERATION

28-00-00 00

1. GENERAL A. Fuel Storage and Distribution The fuel system consists of a series of rubber bladder type cells and an integral (wet wing) tank in each wing connected by a cross-transfer line controlled by a valve. The separate fuel system for each engine has a total fuel capacity of 216 gallons (Ref. Figure 1). The fuel system in each wing consists of two wing leading edge tanks, two box section tanks, an integral (wet wing) tank, and one center section tank, all interconnected to flow into the fuel collector tank by gravity. The fuel collector tank is located within the center section fuel tank, adjacent to the fuselage. The collector tank is equipped with an electric fuel pump, primary jet pump, and two transfer jet pumps. The total usable fuel capacity of each fuel system is 212.5 gallons. The filler cap for this system of tanks is located on the outboard leading edge of each wing near the wing tip. The collector tank has a drain located in the center section, adjacent to the fuselage. The inboard leading edge tank has a drain on the underside of the wing just outboard of the nacelle. The integral (wet wing) fuel tank has a sump drain located approximately midway on the underside of the wing aft of the main spar. The fuel system is vented through a flush vent near the wing tip and a recessed vent coupled to a static vent on the underside of the wing adjacent to the nacelle. The static vent is heated to minimize the possibility of icing and also serves as a backup should the other vents become obstructed. The wing tanks are cross vented with one another and then vented through a float operated vent valve installed on the forward outboard side of the integral fuel tank. A line just aft of the float-operated vent valve extends from the integral fuel tank through a suction relief valve and aft to the flush vent on the underside of the wing. The line from the float-operated vent valve in the integral fuel tank is routed forward along the leading edge of the wing inboard to the nacelle and aft through a check valve to the recessed vent just outboard of the nacelle. Another line tees off from the vent line and extends through a flame arrester to a heated ram vent immediately outboard and aft of the recessed vent. The fuel pressure required to operate the engine is provided by an engine-driven fuel pump immediately upstream of the fuel control unit on the accessory case. An engine-driven boost pump, immediately upstream of the engine-driven fuel pump, provides the motive flow for operation of the primary jet pump which is located in the collector tank. The primary jet pump assists the engine-driven pumps in removing fuel from the collector tank. The supply line from the collector tank is routed from the aft side of the center section tank forward to the engine-driven boost pump through a normally open firewall shutoff valve installed in the fuel line immediately aft of the engine firewall. A cross-transfer line connects the two collector tanks. A switch controlled cross-transfer valve in the line is located at the forward outboard corner of the center section fuel cell. When the valve is in its normally closed position, each engine draws fuel from its respective fuel tanks system. A manually operated cross-transfer control switch is mounted on the upper fuel control panel, just above the fuel quantity gages. When the cross-transfer control switch is actuated, power is drawn from a circuit breaker on the lower fuel control panel to the solenoid of the cross-transfer valve. The cross-transfer valve then opens to allow the electrically- driven fuel pump to transfer fuel to either the left or right fuel system.

28-00-00

Page 1 Nov 1/09

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL The electrically-driven pump located in the bottom of each collector tank is provided as a backup pump should the engine-driven pumps fail; it is also used for all cross-transfer operations. Should one of the electric pumps fail, cross-transfer can only be accomplished from the side of the operative pump. The fuel purge system is designed to assure that any residual fuel in the fuel manifolds is consumed during engine shutdown. During engine operation, engine bleed air (P3) is routed through a filter and check valve and pressurizes a small air tank mounted on the engine truss mount. On engine shutdown the pressure differential between the air tank and fuel manifold causes air to be discharged from the air tank, through a poppet valve, into the fuel manifold system. The air forces the residual fuel, remaining in the fuel manifold, out through the nozzles and into the combustion chamber.

B. Fuel Quantity Indicating Fuel system operation and performance is monitored through various indicators (gages and annunciator lights) found in the flight compartment. Left and right fuel quantity gages indicate total fuel in each system in pounds of fuel remaining. Eight capacitance probes in each wing fuel system provide the signals required to drive the fuel quantity indicators which are located on the fuel control panel.

C. Fuel Flow Indicating Fuel flow gages indicate the actual amount of fuel, in pounds per hour, being metered to each engine at any given time. The fuel flow transmitters are located immediately downstream of the fuel control units. This system is discussed in more detail in Chapter 77, ENGINE INDICATING.

D. Low Fuel Quantity Warning System A low fuel warning system provides bi-level warning to the flight crew when the fuel level in each wing reaches a certain level and again when the fuel in the collector tank reaches a certain level. R FUEL QTY and L FUEL QTY annunciators illuminate when approximately 30 minutes of fuel remains in each wing. The 30-minute flight time provides for fuel in each wing sufficient to feed both engines at maximum continuous power. R FUEL FEED and L FUEL FEED annunciators illuminate when approximately 2 minutes of fuel remains in each collector tank. The 2-minute flight time provides for fuel in each collector tank sufficient to feed both engines at maximum continuous power. The low fuel quantity and low fuel feed annunciators are located in the caution advisory panel, mounted in the center subpanel. When the fuel pressure switch, in the plumbing from the main tank, reads 5 psi or less (1 psi for the UA and UB models), the R FUEL PRESS or L FUEL PRESS warning annunciators will illuminate. The PRESS TO TEST switch on the warning annunciator panel provides energizing current to the test circuit of the low level sensors which are mounted in the fuel tanks and provides for functionally testing the sensors and all associated circuitry. Refer to 28-40-00 for the Low Fuel Quantity Warning Block Diagram.

Page 2 Nov 1/09

28-00-00

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

FUEL NOZZLE MANFOLD

COLECTOR TANK

FUEL PURGE

FUEL CROSSTRANSFER LINE

FUEL AT STRAINER OR FILTER FUEL UNDER PUMP PRESSURE

FLAPPER VALVE

FUEL CONTROL UNIT

DRAIN VALVE

FUEL VENT

STRAINER

CHECK VALVES

ENGINE DRIVEN BOOST PUMP

CHECK VALVE MOTIVE FLOW LINE

LOW LEVEL SENSOR PRESSURE RELIEF VALVE

TRANSFER JET PUMP

FUEL FILLER CAP

DETAIL

A

PROBE

B

CHECK VALVE

FROM FUEL NOZZLE MANIFOLD

FUEL HEATER PRIMARY JET PUMP

SUCTION RELIEF VALVE

L

CHECK VALVE DETAIL

ENGINE FUEL PUMP

FILLER PROBES

P3 AIRLINE

FUEL FLOW TRANSMITTER

ELECTRIC PUMP

FUEL SUPPLY LINE

MOTIVE FLOW

AIR FILTER

DRAIN VALVE

FUEL CROSS-TRANSFER

F

PURGE TANK

TRANSFER JET PUMP

FUEL SUPPLY

FUEL PURGE TANK

B FIREWALL

CHECK VALVES DRAIN VALVE PROBE

FUEL PRESSURE SWITCH FIREWALL SHUTOFF VALVE FUEL FILTER

40 GALLONS

13 GALLONS

F

CROSS-TRANSFER VALVE

DRAIN VALVE

80 GALLONS 35 GALLONS INTEGRAL (WET WING) FUEL TANK

L

23 GALLONS

25 GALLONS

A

PROBE VACUUM RELIEF

COLLECTOR TANK

L

HEATED FUEL VENT FLUSH VENT

VENT FLOAT VALVE

PROBE

FLAME ARRESTOR PROBE

PROBE DRAIN

CHECK VALVE

RECESSED VENT

SHUTOFF VALVE (MANUAL)

LOW LEVEL SENSOR

UC28B 023609AA.AI

Figure 1 Fuel System Schematic (UA-1 and After; UB-1 and After)

28-00-00

Page 3 Nov 1/09

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

200200

FUEL FUEL SYSTEM (UA-1 AND AFTER; UB-1 AND AFTER) MAINTENANCE PRACTICES 1. PROCEDURES A. Servicing Refer to Chapter 12-10-00 for fuel filling and draining procedures, for types and brands of approved fuels, and for routine maintenance practices.

2. AIR-MAZE FUEL FILTER CLEANING - PRIMARY METHOD The primary method of cleaning the Air-Maze fuel filters used on this airplane is by use of an ultrasonic cleaner. The following data provides information on the methods to be used for this purpose.

A. Initial Cleaning - Filter Pack Element (1) Carefully remove the filter pack element from the housing and remove all packings. (2) Do not disassemble the filter pack element for this initial cleaning. (3) Plug open end(s) of filter(s) packs with rubber plugs and immerse in a recommended cleaner (51 or 2, Table 1, Chapter 91-00-00). (4) Allow items to soak for 30 to 60 minutes to remove sludge and baked on contaminates. (5) Agitate parts for several minutes then remove the parts from cleaner. (6) Allow parts to drain and then rinse the parts in new cleaner. (7) Use filtered low pressure air (not exceeding 30 psi) to remove the excess cleaner and to facilitate drying. (8) Preform the FINAL CLEANING - FILTER PACK ELEMENT procedure. NOTE: The ultrasonic cleaning tank used in this procedure must have a minimum of 3 watts of power per square inch of tank bottom area.

B. Final Cleaning - Filter Pack Element (1) Fill tank with a recommended ultrasonic cleaner (51 or 2, Table 1, Chapter 91-00-00). (2) Unplug the filters for this cleaning. (3) Disassemble and remove the filter discs from the center tube of the filter pack element. Slide the discs onto a rod rack which is supported horizontally in the tank. Provide a 1/4-inch space between the filter discs. (4) Metal valve parts, etc., may be cleaned simultaneously in a separate basket suspended in the solution. (5) Clean the filters for 20 minutes.

28-00-00

Page 201 Nov 1/09

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (6) Remove the filter discs from the tank and drain or use filtered low pressure air (not exceeding 30 psi) to remove the excess cleaner and to facilitate drying. (7) Examine filter discs for cleanliness and screen damage. If the screen is damaged, replace the filter disc. Examine parts for damage and replace if required. Inspect the filter discs with a 6X magnifier to ensure debris removal. (8) Repeat Steps (3) thru (7) if item is not clean. (9) Reassemble the filter pack element. (10) Replace packings and gaskets (if any) and reassemble valve parts and filter pack element to housing.

3. AIR-MAZE FUEL FILTER CLEANING - SECONDARY METHOD The secondary method of cleaning the Air-Maze fuel filters used on this airplane is by use of an chemical cleaners. The filter pack elements are cleaned twice when chemicals are used. The following data provides information on the methods to be used for this purpose.

A. Initial Cleaning - Secondary Method - Filter Pack Element (1) Carefully remove the filter pack element from the housing and remove all packings. (2) Do not disassemble the filter pack element for this initial cleaning. (3) Plug open end(s) of filter(s) packs with rubber plugs and immerse in a recommended cleaner (51 or 2, Table 1, Chapter 91-00-00). (4) Allow items to soak for 30 to 60 minutes to remove sludge and baked on contaminates. (5) Remove the filter pack element and parts from cleaner. (6) Allow items to air dry or use filtered low pressure air (not exceeding 30 psi) to remove the excess cleaner and to facilitate drying. (7) Preform the FINAL CLEANING - SECONDARY METHOD - FILTER PACK ELEMENT procedure.

B. Final Cleaning - Secondary Method - Filter Pack Element (1) Immerse the filter pack element in a recommended cleaner (51 or 2, Table 1, Chapter 91-00-00). (2) Agitate solution for twenty minutes. (3) Use a stiff bristled non-metallic brush to remove any remaining dirt or deposits. (4) Unplug the filter and then drain or use filtered low pressure air (not exceeding 30 psi) to remove the excess cleaner and to facilitate drying. (5) Examine filter discs for cleanliness and screen damage. If the screen is damaged, replace the filter disc.

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28-00-00

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

(6) Repeat Steps (1) thru (5) if item is not clean. (7) Replace packings and gaskets (if any) and reassemble valve parts and filter pack element to housing.

28-00-00

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

FUEL FUEL SYSTEM (UC-1 AND AFTER) DESCRIPTION AND OPERATION

28-01-00 00

1. GENERAL A. Fuel Storage and Distribution The fuel system consists of two independent systems, one for each engine. The systems are connected by a solenoid valve-controlled cross-transfer line. Each independent system consists of two integral tanks (the main and auxiliary tank). The main tank in each wing holds a maximum of 241 gallons of usable fuel (Ref. Figure 1). The auxiliary tanks, located in each side of the center section, contain 92.2 gallons of usable fuel each. The ribs, which form part of the structure of the main tank, also function as baffles that prevent the fuel from sloshing. Fuel from the main tank is gravity-fed into the collector tank (the inboard aft area of the outboard wing). The collector tank is equipped with an electric fuel boost pump, transfer jet pump, main tank jet pump and a manifold. Fuel is transferred from the auxiliary to the main tank by activating the electric fuel pump in the auxiliary tank. The switch for the transfer pump is located on the fuel control panel on the pilot's side of the crew compartment. A pressure switch located in the auxiliary transfer line, in conjunction with a float switch located in the auxiliary fuel tank, automatically cuts power to the auxiliary fuel pump when the usable fuel in the tank is exhausted. The system is fueled by filling the main tank first through ports located on the outboard leading edge adjacent to the wing tip, then filling the auxiliary tanks through ports located near the inboard side of the nacelle and aft of the main spar. The vent system for the fuel tanks is located in the wing tips and vents the system through a float valve on each side. The valve is designed to shut off when activated by the presence of fuel. However, a pressure activated backup vent prevents an overpressure condition by opening if pressure exceeds approximately 1-1/2 psi. The system is equipped with flame arrestors and protected from freeze-up by a switch-controlled heated vent. The switch which controls the heated vent is located on the pilots RH subpanel. The auxiliary tank vents into the main tank. The cross-transfer line connecting the two systems is controlled by a normally closed, switch-operated solenoid valve. When the valve is in the closed position, each engine draws fuel from its individual system. If cross-transfer is desired, the manually activated switch must be moved to the TRANSFER FLOW position which will activate the standby pump. In addition to the cross-transfer function, the electric boost pump can provide fuel to the engine should the engine-driven fuel pump fail. Power for the switches is drawn through the appropriately placarded circuit breakers at the bottom of the fuel control panel. Fuel for the engine is provided by an engine-driven fuel pump. The engine-driven fuel boost pump (located upstream of the engine-driven fuel pump) provides motive flow for operation of the jet pumps which aid in moving fuel from the main tank.

B. Fuel Purge System (EPA) The fuel purge system is designed to assure that any residual fuel in the fuel manifolds is consumed during engine shutdown. During engine operation, engine bleed air (P3) is routed through a filter and check valve and pressurizes a small air tank mounted on the engine truss mount. On engine shutdown the pressure differential between the air tank and the fuel manifold causes air to be discharged from the air tank, through a poppet valve, and into the fuel manifold system. The air forces the residual fuel, remaining in the fuel manifold, out through the nozzles and into the combustion chamber.

28-01-00

Page 1 Nov 1/09

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

C. Fuel Quantity Indicating Fuel system operation and performance is monitored through various indicators (gages and annunciator lights) found in the flight compartment. Left and right fuel quantity gages indicate the amount of fuel in each tank in pounds of fuel remaining. Eight capacitance probes in each wing fuel system provide the signals required to drive the fuel quantity indicators, located on the fuel control panel.

D. Fuel Flow Indicating Fuel flow gages indicate the actual amount of fuel, in pounds per hour, being metered to each engine at any given time. The fuel flow transmitters are located immediately downstream of the fuel control units. This system is detailed in Chapter 77-00-00.

E. Low Fuel Quantity Warning System The low fuel warning system provides bi-level warning to the flight crew when the fuel level in each tank reaches a predetermined level and again when the fuel in the collector tank area is nearly exhausted. When the fuel in the main tank or tanks reaches a level that equals approximately 30 minutes of flight time, R FUEL QTY, L FUEL QTY or both annunciators will be actuated by liquid level sensors mounted on the inboard forward side of the main tank. If the fuel in the collector tank area drops to a level sufficient to provide approximately two minutes flight time, R FUEL FEED, L FUEL FEED or both annunciator lights will illuminate. Flight time fuel means that sufficient fuel remains to feed the engine at maximum continuous power. When the fuel pressure switch, in the plumbing from the main tank, reads 5 psi or less, the R FUEL PRESS or L FUEL PRESS warning annunciators will illuminate. The PRESS-TO-TEST switch on the warning annunciator panel provides current to the test circuit of the low level sensors and all associated circuitry. Refer to 28-41-00 for the Low Fuel Quantity Warning Block Diagram.

Page 2 Nov 1/09

28-01-00

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

C93UC28B1388.AI

Figure 1 Fuel System Schematic (UC-1 and After)

28-01-00

Page 3 Nov 1/09

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

FUEL FUEL SYSTEM (UC-1 AND AFTER) MAINTENANCE PRACTICES

200200

1. PROCEDURES A. Servicing Refer to Chapter 12-10-00 for fuel filling and draining procedures, for types and brands of approved fuels, and for routine maintenance practices.

28-01-00

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

28-10-00 200200

FUEL FUEL STORAGE (UA-1 AND AFTER; UB-1 AND AFTER) MAINTENANCE PRACTICES 1. FUEL CELLS Rubber bladder type fuel cells are equipped with nipples for attaching plumbing lines to the fuel cell. In order to avoid damaging the fuel cell or its nipple fittings, the following torque values should be applied to the nipple fitting clamps during installation: Table 201 FUEL CELL NIPPLE TORQUE (UA-1 and After; UB-1 and After) NIPPLE FITTING I.D. (IN)

TORQUE INCH-POUNDS

0.25 thru 0.62 0.75 thru 1.00 1.50 2.00 3.00

12 to 15 15 to 20 25 to 30 30 to 35 35 to 40

Bladder type fuel cells having yellow nipples shall be torqued to 25 ± 5 inch-pounds; all others shall be torqued according to Table 201. Always position hose clamps on internal nipple fittings so that the screw body travels freely as the screw is tightened. Never allow the screw body to wedge between nipple fitting O.D. and tank wall. This can result in false torque readings, leaks and/or nipple fitting damage. NOTE: Bladder type fuel cells returned to the manufacturer for repair or replacement will not be repaired if foreign material has been applied to the fittings of the fuel tank wall, either inside or outside. Foreign material is considered to be any material, such as sealants, coatings, cements, etc. not approved by the manufacturer for the installation, operation or repair of a bladder type fuel cell. To facilitate installation of bladder type fuel cells, allow the fuel cell to set at room temperature (above 60°F) for a period of time sufficient to ensure flexibility.

2. WING OUTBOARD LEADING EDGE FUEL CELL A. Removal The fuel cell is located in the leading edge outboard wing panel (Ref. Figure 201). (1) Make sure all electrical power to the airplane is disconnected. (2) Drain the wing fuel cells (Ref. Chapter 12-10-00 DRAINING THE FUEL SYSTEM). (3) Remove the twelve screws securing the fuel filler adapter to the upper wing skin. (4) Remove the access plates on the lower surface of the wing leading edge. (5) Working through the inboard access opening, remove the clamps from the interconnect tubes (two inboard and one aft) and pull the nipples off the tubes.

28-10-00

Page 201 Nov 1/10

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (6) Working through the lower aft outboard access opening, remove the clamp from the outboard aft interconnect tube and pull the nipple off the tube behind the filler neck. (7) Working through the two access openings on the underside of the wing leading edge, release the fuel cell fasteners. Collapse the cell and remove it through the inboard access opening.

B. Installation NOTE: Expose the fuel cell to temperatures of at least 60°F for a period sufficient to ensure flexibility during installation. (1) Check the fuel cell cavity for foreign objects or debris. Collapse the cell and install it through the inboard access opening in the bottom of the wing forward of the main spar (Ref. Figure 201). (2) Secure the snap fasteners that hold the fuel cell against the upper wing skin. (3) Working through the lower aft outboard access opening behind the filler neck, install the interconnect tube and clamp. Torque the Wittec Worm Drive (WWD) clamps (Ref. Table 201). (4) Working through the inboard access opening, install the interconnect tubes (two inboard and one aft). Secure with WWD clamps and torque (Ref. Table 201). (5) Install the access plates on the lower surface of the wing leading edge.

Figure 201 Wing Outboard Leading Edge Fuel Cell (UA-1 and After; UB-1 and After)

Page 202 Nov 1/10

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

3. WING INBOARD LEADING EDGE FUEL CELL A. Removal The fuel cell is located in the leading edge inboard wing panel (Ref. Figure 202). (1) Make sure all electrical power to the airplane is disconnected. (2) Drain the wing fuel cells (Ref. Chapter 12-10-00 DRAINING THE FUEL SYSTEM). (3) Unclamp and remove the drain valve spacer and washer from the inboard bottom of the cell adjacent to the nacelle. (4) Remove the outboard access cover on top of the wing to gain access to the fuel quantity probe. (5) Remove the jumper wire and detach the lead wires to the fuel quantity probe. CAUTION: Handle the fuel quantity probe carefully; damage to the surface of the probe tubes will destroy the accuracy of the probe as a measuring device. (6) Unsafety and remove the five mounting screws and remove the fuel quantity probe from the fuel cell. (7) Remove the attaching bolts and the cover plate to the fuel cell access fitting. (8) Working through the fuel access opening, disconnect the large interconnect in the aft side of the fuel cell inboard of the opening by removing the Wittec Worm Drive (WWD) clamp securing the interconnect line to the nipple cap of the fuel cell. (9) Disconnect the two cell interconnect lines at the outboard end of the cell by removing the WWD clamps securing the interconnect lines to the nipple caps. (10) Remove the inboard access cover on top of the wing to gain access to the fuel quantity probe and fuel cell cover plate. (11) Remove the jumper wire and detach the three lead wires to the fuel quantity probe. CAUTION: Handle the fuel quantity probe carefully; damage to the surface of the probe tubes will destroy the accuracy of the probe as a measuring device. (12) Unsafety and remove the five mounting screws and remove the fuel quantity probe from the fuel cell. (13) Remove the attaching bolts and the cover plate to the fuel cell access fitting. (14) Working through the fuel access opening, disconnect the large interconnect line directly aft of the access opening by disconnecting the WWD clamp securing the interconnect line to the fuel cell. (15) Remove the WWD clamp that connects the small interconnect line to the fuel cell nipple just outboard of the large interconnect line removed in the preceding Step. (16) Remove the access panels adjacent to the outboard side of the nacelle. (17) Disconnect the purge line running from the aft inboard side of the leading edge fuel cell to the forward inboard side of the wing inboard aft fuel cell.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (18) Release the snap fasteners securing the cell to the upper wing skin. (19) Collapse the cell and remove it through the outboard access opening on top of the wing leading edge.

B. Installation The fuel cell is located in the leading edge wing panel (Ref. Figure 202). NOTE: Expose the fuel cell to temperatures of at least 60°F for a period sufficient to ensure flexibility during installation. (1) Collapse the cell and install it through the outboard access opening on top of the wing leading edge. (2) Secure the snap fasteners that hold the fuel cell against the upper wing skin. (3) Install the washer, spacer, and drain valve on the inboard aft end of the fuel cell and secure them in place with the retaining clamp. The washer is to be installed between the spacer and cell liner. (4) Connect the purge line at the aft inboard end of the fuel cell. Torque the Wittec Worm Drive (WWD) clamps (Ref. Table 201). (5) Working through the inboard access opening on top of the wing, connect the large and small interconnect lines at the aft side of the cell. Torque WWD clamps (Ref. Table 201). (6) Install the cover plate and gasket on the fuel cell with the attaching bolts, then lock the bolts in place with safety wire. Apply gasket cement (33, Table 1, Chapter 91-00-00) between the adapter and skin when installing the cover plate. CAUTION: Handle the fuel quantity probe carefully; damage to the surface of the probe tubes will destroy the accuracy of the fuel quantity probe as a measuring device. (7) Insert the fuel quantity probe into its opening in the fuel cell. NOTE: Install the free end of the bonding jumper under one of the cover mounting bolts. (8) Secure the fuel quantity probe in place with the attaching bolts. Torque the bolts to 25 ± 5 inch-pounds, then safety wire them. Install the free end of the bond jumper under one of the fuel quantity probe mounting bolts. (9) Attach the three lead wires to the fuel quantity probe and reattach the jumper wire to the probe. (10) Working through the outboard access opening on top of the wing, connect the two cell interconnect lines at the outboard end of the fuel cell. Torque the WWD clamps (Ref. Table 201). (11) Working through the outboard access opening on top of the wing, connect the large interconnect line at the aft side of the cell inboard of the access opening. Torque the WWD clamps (Ref. Table 201). (12) Install the cover plate and gasket on the fuel cell with the attaching bolts, then lock the bolts in place with safety wire. Apply gasket cement (33, Table 1, Chapter 91-00-00) between the adapter and skin when installing the cover plate. NOTE: Install the free end of the bonding jumper under one of the cover mounting bolts.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (13) Insert the fuel quantity probe into its opening in the fuel cell. CAUTION: Handle the fuel quantity probe carefully; damage to the surface of the probe tubes will destroy the accuracy of the fuel quantity probe as a measuring device. (14) Secure the fuel quantity probe in place with the attaching bolts. Torque the bolts to 25 ± 5 inch-pounds, then safety wire them. NOTE: Install the free end of the bonding jumper under one of the mounting bolts of the fuel quantity probe. (15) Attach the lead wires to the fuel quantity probe and attach the jumper wire to the probe. (16) Install all access plates on the upper and lower surfaces of the wing.

Figure 202 Wing Inboard Leading Edge Fuel Cell (UA-1 and After; UB-1 and After)

4. AFT OUTBOARD WING CENTER FUEL CELL A. Removal (1) Make sure all electrical power to the airplane is disconnected. (2) Drain the main fuel system (Ref. Chapter 12-10-00 DRAINING THE FUEL SYSTEM). (3) Remove the access cover on the underside of the wing (Ref. Figure 203). (4) Working through this access opening, disconnect the large interconnect line forward of the opening by removing the Wittec Worm Drive (WWD) clamp securing the line to the cell nipple.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (5) Working through the same opening, disconnect the large interconnect line and the hose at the forward outboard side of the cell by removing the WWD clamps securing them in place. (6) Disconnect the small interconnect line at the outboard aft end of the fuel cell by removing the WWD clamp holding it in place. (7) Disconnect the small and large interconnect lines at the inboard end of the fuel cell by removing the WWD clamps holding them in place. (8) Release the snap fasteners securing the cell to the upper wing skin. (9) Collapse the cell and remove it through the access opening on the underside of the wing.

B. Installation NOTE: Expose the fuel cell to temperatures of at least 60°F for a period sufficient to ensure flexibility during installation. (1) Check the fuel cell cavity for foreign objects or debris. Collapse the fuel cell and install it through the outboard access cover on the underside of the wing (Ref. Figure 203). (2) Working through this access opening, secure the snap fasteners on the hangers that hold the fuel cell against the upper skin of the wing. (3) Connect the small and large interconnect lines at the inboard end of the fuel cell. Torque the WWD clamps. (Ref. Table 201). (4) Connect the small interconnect line at the outboard end of the fuel cell. Torque the WWD clamps (Ref. Table 201). (5) Connect the large interconnect line and hose at the forward outboard end of the fuel cell. Torque the WWD clamps (Ref. Table 201). (6) Connect the large interconnect line at the forward side of the fuel cell directly in front of the access opening. Torque the WWD clamp (Ref. Table 201). (7) Install the gasket and access cover.

5. AFT INBOARD WING FUEL CELL A. Removal (1) Make sure all electrical power to the airplane is disconnected. (2) Drain the main fuel system (Ref. Chapter 12-10-00 DRAINING THE FUEL SYSTEM). (3) Remove the access cover over the fuel quantity probe (Ref. Figure 204). (4) Remove the jumper wire and detach the three lead wires to the fuel quantity probe. CAUTION: Handle the fuel quantity probe carefully; damage to the surface of the probe tubes will destroy the accuracy of the probe as a measuring device. (5) Unsafety and remove the five mounting screws, then remove the fuel quantity probe from the fuel cell.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (6) Remove the access cover, located on the inboard, aft underside of the wing. (7) Working through this access opening, remove the Wittec Worm Drive (WWD) clamps and disconnect the large interconnect tube on the inboard, aft side of the fuel cell. (8) Disconnect the two smaller tubes at the inboard side of the cell by removing the WWD clamps. (9) Remove the WWD clamps and disconnect the small and large interconnect tubes at the forward end of the cell. (10) Disconnect the two interconnect tubes at the outboard side of the cell by removing the WWD clamps. (11) Release the snap fasteners securing the cell to the upper wing skin. (12) Collapse the cell and remove it through the inboard, aft access opening on the underside of the wing.

Figure 203 Aft Outboard Wing Center Fuel Cell (UA-1 and After; UB-1 and After)

B. Installation NOTE: Expose the fuel cell to temperatures of at least 60°F for a period sufficient to ensure flexibility during installation. (1) Check the fuel cell cavity for foreign objects or debris. Collapse the fuel cell and install it through the inboard, aft access opening on the underside of the wing (Ref. Figure 204). (2) Working through this access opening, secure the six snap fasteners that hold the fuel cell against the upper skin of the wing.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (3) Connect the small and large interconnect lines at the outboard end of the fuel cell. Torque the Wittec Worm Drive (WWD) nipple clamps (Ref. Table 201). (4) Connect the small and large interconnect lines at the forward side of the fuel cell. Torque the WWD nipple clamps (Ref. Table 201). (5) Connect the two small tubes at the inboard side of the fuel cell. Torque the WWD nipple clamps (Ref. Table 201). (6) Connect the large interconnect tube at the inboard, aft side of the fuel cell. Torque the WWD nipple clamps (Ref. Table 201). (7) Install the access cover, located on the inboard, aft underside of the wing. CAUTION: Handle the fuel quantity probe carefully; damage to the surface of the probe tubes will destroy the accuracy of the probe as a measuring device. (8) Insert the fuel quantity probe into its opening in the fuel cell. NOTE: Install the free end of the bonding jumper under one of the mounting bolts of the fuel quantity probe. (9) Secure the fuel quantity probe in place with the attaching bolts. Torque the bolts to 25 ± 5 inch-pounds, then safety wire. (10) Attach the three lead wires to the fuel quantity probe. (11) Install the access cover over the fuel quantity probe and secure in place with the attaching bolts.

Figure 204 Aft Inboard Wing Fuel Cell (UA-1 and After; UB-1 and After)

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6. CENTER SECTION FUEL CELL A. Removal NOTE: Before removing the center section fuel cell, remove the fuel collector tank (Ref. FUEL SUPPLY COLLECTOR TANK REMOVAL). (1) Make sure all electrical power to the airplane is disconnected. (2) Drain the center section fuel system (Ref. Chapter 12-10-00 DRAINING THE FUEL SYSTEM). (3) Remove the small round cover adjacent to the fuselage to gain access to the fuel quantity probe (Ref. Figure 205). (4) Disconnect the three electrical leads of the fuel quantity probe. CAUTION: Handle the fuel quantity probe carefully; damage to the surface of the probe tubes will destroy the accuracy of the probe as a measuring device. (5) Unsafety and remove the five mounting screws securing the jumper wire and fuel quantity probe in place, then remove the probe from the fuel cell. (6) Remove the large oblong cover aft of the main spar on top of the center section to gain access to the fuel cell and fuel quantity probe. (7) Disconnect the three electrical leads of the fuel quantity probe. (8) Remove the five bolts securing the low level sensor in place and remove the sensor from the fuel cell. CAUTION: Handle the fuel quantity probe carefully, for damage to the surface of the probe tubes will destroy the accuracy of the probe as a measuring device. (9) Unsafety and remove the five mounting screws securing the jumper wire and fuel quantity probe in place, then remove the probe from the fuel cell. (10) Remove the large access cover on the bottom of the center section. (11) Working through the access opening on the top of the center section, disconnect the three fuel lines from the fittings at the aft end of the center section fuel cell. (12) Working through the cover opening on top of the center section fuel cell, remove the Wittec Worm Drive (WWD) clamps to disconnect the interconnect tube and vent line at the outboard end of the fuel cell. (13) Unsnap the fasteners securing the cell to the upper wing skin. (14) Collapse the fuel cell and remove it through the large access opening on bottom of the center section.

B. Installation NOTE: Expose the fuel cell to temperatures of at least 60°F for a period sufficient to ensure flexibility during installation.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (1) Check the fuel cell cavity for foreign objects and debris. Collapse the fuel cell and install it through the large opening on the underside of the center section (Ref. Figure 205). (2) Snap the fasteners that secure the fuel cell to the upper center section. (3) Working through the cover opening on top of the center section fuel cell, connect the interconnect tube at the outboard aft side and the vent tube at the outboard forward side of the fuel cell. Install the Wittec Worm Drive (WWD) clamps and torque (Ref. Table 201). (4) Working through the access opening on the top of the center section, connect the three fuel lines to the fittings at the aft end of the center section fuel cell. (5) Insert the low level sensor and gasket into its opening in the fuel cell. Install the plain washer under the bolt head. Apply Oakite solution (35, Table 1, Chapter 91-00-00) to assemble thread seal washer. Install the washers and bolts. NOTE: Do not use an air wrench or speed equipment to install these bolts. (6) Torque the bolts 25 ± 5 inch-pounds. CAUTION: Handle the fuel quantity probe carefully, for damage to the surface of the probe tubes will destroy the accuracy of the probe as a measuring device. (7) Insert the fuel quantity probe into its opening in the fuel cell cover and secure in place with the attaching bolts. Torque the bolts to 25 ± 5 inch-pounds, then safety the bolts. Install the free end of the bonding jumper under one of the probe mounting bolts. (8) Attach the three lead wires to the fuel quantity probe. CAUTION: Handle the fuel quantity probe carefully, for damage to the surface of the probe tubes will destroy the accuracy of the probe as a measuring device. (9) Insert the inboard fuel quantity probe into its opening on top of the center section aft of the main spar and adjacent to the fuselage. Torque the bolts to 25 ± 5 inch-pounds, then safety the bolts. Install the free end of the bonding jumper under one of the probe mounting bolts. (10) Install the access cover over the fuel quantity probe and secure in place with the attaching bolts. (11) Install the fuel collector tank (Ref. FUEL SUPPLY COLLECTOR TANK INSTALLATION).

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Figure 205 Center Section Fuel Cell (UA-1 and After; UB-1 and After)

7. FUEL SUPPLY COLLECTOR TANK A. Removal (1) Make sure all electrical power to the airplane is disconnected. (2) Drain the fuel system (Ref. Chapter 12-10-00 DRAINING THE FUEL SYSTEM). (3) To gain access to the collector tank plumbing, remove the large oblong and small round access covers adjacent to the fuselage on the underside of the wing. The small round access cover is the fourth cover outboard of the fuselage (Ref. Figure 206). (4) Remove the large access cover on top of the center section, aft of the main spar. Disconnect the three plumbing lines and unions at the outboard side of the collector tank. (5) Remove the access plate from the bottom of the collector tank. (6) Disconnect the electrical lead from the boost pump. Disconnect the attaching washers, bolts and plumbing and remove the boost pump from the collector tank. (7) Disconnect the fuel lines from each jet transfer pump. (8) Remove the pump retaining bolts and washers, disconnect the plumbing and remove the jet transfer pumps from the collector tank.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (9) Disconnect the electrical leads from the fuel quantity sensor and fuel level probe. (10) Working through the large oblong hole adjacent to the fuselage on the underside of the wing, support the fuel collector tank and remove the attaching bolts and washers. Remove the fuel collector tank from the airplane.

B. Installation (1) Make sure that all power to the airplane is disconnected. (2) Check that the fuel collector tank cavity is free from foreign objects and debris. (3) Working through the large oblong access opening adjacent to the fuselage on the underside of the wing, position the fuel collector tank and install the gasket and attaching washers and bolts. Torque the bolts 45 to 50 inch-pounds. Safety wire the bolts (Ref. Figure 206). (4) Connect the electrical connector to the fuel quantity sensor and fuel level probe. (5) Position the jet transfer pumps, located at the forward and aft end of the collector tank. Connect the plumbing to each pump and install the attaching washers and bolts. (6) Position the boost pump and install the attaching washers and bolts. Torque the attaching bolts 45 to 50 inch-pounds. Safety wire the bolts. (7) Install the fuel plumbing to the boost pump. (8) Position the access plate to the bottom of the fuel collector tank to install the attaching washers and bolts. Torque the bolts 20 to 30 inch-pounds. Safety wire the bolts. (9) Working through the large access opening on the top of the center section aft of the main spar, connect the three plumbing lines and unions at the outboard side of the fuel collector tank. (10) Install the large access cover on top of the center section. (11) Install the small round and large oblong access covers on the underside of the wing, adjacent to the fuselage.

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Figure 206 Fuel Supply Collector Tank (UA-1 and After; UB-1 and After)

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8. LEAKAGE CHECKS AND REPAIRS A. Leakage Checks and Repairs (Integral Fuel Cells) To classify the degree of leakage in an integral (wet wing) fuel cell, measure the size of the wet area around the leak. A more accurate measurement may be obtained by wiping the leakage clean and applying talcum powder in the area of the leak. After 30 minutes, check the area to determine if the leak classifies as a stain, seep, heavy seep, or running leak (Ref. Figure 207). Fuel leaks must also be classified as to whether they occur in an open area or in an enclosed area to differentiate between those that require immediate repair and those not considered potential flight hazards. WARNING: Any leakage in an enclosed area, such as the wheel well, or in an area where the fuel will blow into the fuselage, requires grounding until repair is made. (1) Heavy seeps or smaller leaks in an open area may be temporarily repaired any time the airplane is grounded for other maintenance. No action is required at any repair station other than the maintenance base of operations. An external sealer (120, Table 1, Chapter 91-00-00) may be used. (2) If leaks are to be repaired from inside the tank, drain all fuel from the affected tank (Ref. Chapter 12-10-00 DRAINING THE FUEL SYSTEM). Remove all damaged sealer from the vicinity of the leak with a wooden or plastic scraper. Chamfer the ends of the remaining sealant so that the new sealant forms a smooth continuous flow. Refer to Table 202 for recommended sealers. (3) If the leakage is around a rivet, strike the rivet. This can only be done once. If the leak persists, replace the rivet. An alternate method for repair of rivet leaks is to install a fuel repair patch (122, Table 1, Chapter 91-00-00). The patch is a thin aluminum disc that is applied directly over the exterior side of the leak and bonded in place with epoxy. Patches are available for flush or protruding head rivets. Because the adhesive used with these patches will cure in the presence of fuel, it is not necessary to drain the tank prior to patch installation. (4) If the leakage is around a bolt with a gasket type seal, torque the bolt. If the leak persists, replace the seal or the bolt. Refer to the appropriate Model 1900 Parts Catalog for component part numbers. (5) If the leakage is at the gasket around an access opening or fitting, torque the attachments. Torque access plate bolts 45 to 50 inch-pounds. If the leak persists, replace the gasket. (6) If the leakage is around anchor nuts of access plates on the underside of the wing, seal around the nuts. NOTE: If the drain valve was removed during the foregoing checks and repairs, install the valve retaining nut with the castellated portion down, then torque the nut 75 to 100 inch-pounds and safety. Refer to the Model 1900 Airliner Series Structural Repair Manual for wing skin and structural member repairs. Seal repairs as necessary with the appropriate sealer (Ref. Table 202).

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Table 202 FUEL CELL SEALERS MFG’S NO. & CLASS

APPLICATION/ POT LIFE TIME (HOURS)*

TACK FREE/ REFUELING TIME (HOURS)*

CURE/ DISPATCH TIME (HOURS)*

TYPE OF SEAL

PR-1440 B-1/2

1/2

10

30

Access Plates

PR-1440 B-2

2

36

72

PR-1440 B-4

4

48

96

All Types

MIL-S-8802

PR-1440 A-1/2

1/2

10

30

Seal

PR-1826

PR-1826 B-1/4**

1/4

1 @ 77°F 4 @ 40°F 8 @ 20°F

1 @ 77°F 4 @ 40°F 8 @ 20°F

Spot Sealing

PR-1440

PR-1826 B-1/2**

1/2

2 @ 77°F 8 @ 40°F 16 @ 20°F

2 @ 77°F 8 @ 40°F 16 @ 20°F

Medium Area Sealing

PR-1440

PR-1826 B-2**

2

9 @ 77°F 36 @ 40°F 72 @ 20°F

9 @ 77°F 36 @ 40°F 72 @ 20°F

Large Areas at Low Temperatures

PR-1440

Pro-seal 860

10 min

1.5

4

All types

EC 776 SR 1 Part Material

15 min to 1 hr

24

Coating of Rivets

PR-1005-L Buna-N Slosh Coating 1 Part Material

20 min to 1 hr

48

Coating

ALTERNATE MATERIALS

MIL-S-4383 EC-164

* Application time is stated as minimum requirement, tack free and cure times are maximum requirements. All time requirements are figured at 77° ± 2°F with 50% ± 5% relative humidity. As a general precaution, for every 10°F rise in temperature from 77°F, reduce application time, tack free times and cure times by one half (1/2) and for every 10°F below 77°F, double the stated time. ** To ensure the bond between metal parts and/or existing sealant, use primer where required (supplied with sealer).

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Figure 207 Leakage of Integral (Wet Wing) Fuel Cells (UC-1 and After)

B. Fuel Cell Leakage Test - Bladder-Type Fuel Cells Rubber bladder type fuel cells may be bench tested for leakage by sealing off all openings and inflating the empty cell to 1/4 psi with a mixture of shop air and ammonia gas, then checking for visible indications of leakage on a cloth saturated with phenolphthalein solution. To set up and conduct the leakage test, proceed as follows: (1) The following equipment is required and should be hooked up as indicated (Ref. Figure 208). (a) Closure plates for the fuel cell openings. Such plates may be fabricated of aluminum sheet cut to a size sufficient to cover the cell openings. Drill holes in the closure plate to match the hole pattern around the opening in the fuel cell. (b) Rubber stoppers to plug the fitting openings in the fuel cell. One of the stoppers should have a hole for insertion of the plastic tubing used to connect the fuel cell into the test setup. (c) A manometer for measuring 6 inches of water differential. The manometer can be fabricated from glass or clear plastic tubing; frame and scale similar to the illustration shown. (d) A regulator that can be set to provide 1/4 psi (6 inches of water) from a supply of shop air. (e) Two flasks (or bottles) approximately one liter (or quart) in capacity. A third container may be hooked into the test setup to provide an optional overflow collector if desired. The two containers should be provided with rubber stoppers that have holes for the insertion of 1/4 inch tubes (glass or metal) (Ref. Figure 208). (f) Plastic tubing of a size to provide a leak free fit over the tubes and of a length sufficient to interconnect the test components (Ref. Figure 208). (g) Make up a solution of phenolphthalein as follows: Add 1/3 ounce phenolphthalein crystals to 1/2 gallon ethyl alcohol, mix, then add 1/2 gallon water. Page 216 Nov 1/10

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (h) Make up an ammonia solution by adding 100cc (3 fluid ounces) of concentrated ammonium hydroxide (NH4OH) per gallon of water. (2) Place the fuel cell and test equipment on a clean work bench. CAUTION: Make sure the work area is clean of metal shavings or other debris that could damage the fuel cell. (3) Install the closure plates over the fuel cell openings and torque the retaining screws as specified in the installation section of the maintenance manual for the openings, then insert the rubber stoppers into the open fittings. (4) The flask (or bottle) containing the ammonium hydroxide solution should be 1/3 to 1/2 full (Ref. Figure 208). (5) Connect a shop air supply to the regulator and interconnect the regulator, beakers, fuel cell, and manometer (Ref. Figure 208). (6) Inflate the fuel cell to 1/4 psi with a mixture of shop air and ammonium gas. A 6 inch difference in the two water levels of the manometer will indicate that the fuel cell is inflated to 1/4 psi. It is not necessary to restrain the cell other than to keep it from rolling off the bench. The filling of the cell will be rather slow at the 1/4 psi, but should not be rushed as overpressure of the cell could result. CAUTION: Wear rubber gloves to protect against skin irritation when handling the cloth. As a further protection against possible penetration of the phenolphthalein solution through the gloves, wash your hands thoroughly after finishing the test. (7) Saturate a large, clean cloth with phenolphthalein. Immerse it in a container and squeeze out excess liquid. NOTE: Continued use of the testing cloth will require repeated saturations with phenolphthalein since rapid evaporation of the alcohol from the cloth progressively reduces the sensitivity of the test unless the solution in the cloth is frequently renewed. (8) Lay the cloth over the various portions of the fuel cell until the entire exterior of the cell has been covered. With each application of the cloth, watch for the formation of a reddish pink stain on the cloth to indicate the presence of a leak. Encircle the area on the fuel cell beneath such stains with a chalk mark to pinpoint the locations of leaks.

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Figure 208 Setup for Leakage Test (UA-1 and After; UB-1 and After)

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C. Fuel Cell Repair Bladder Type fuel cells have a repair manual number stencilled on the fuel cell exterior surface. The appropriate repair manual should be referred to for the correct repair procedures and repair materials to be used with a particular fuel cell construction. All fuel cells received at the manufacturer having foreign material applied to the fittings or the fuel cell wall, either inside or outside, will not be repaired and all warranty is null and void. Foreign material is defined as any material such as sealants, coatings, cements, etc. not approved by the manufacturer for the installation, operation or repair of a bladder type fuel cell. For information regarding the repair, handling, and storage of BTC-85 fuel cell constructions, refer to the manual entitled Repair and Maintenance Manual for Vithane Fuel Cells in the Model 1900 Airliner Series Component Maintenance Manual. BTC-39 and BTC-39-1 type fuel cells should be replaced, rather than repaired, when damaged.

D. Fuel Cell Storage CAUTION: Never store fuel cells in the vicinity of electrical equipment, such as generators and motors. The movement of brushes across the commutators of these units results in minute sparks that cause the formation of ozone gas. Ozone has a highly destructive effect on fuel cells. Although it leaves no visible indication, ozone makes the material of which the cells are formed brittle so that the cells will disintegrate upon the application of stress. Bladder type fuel cells are constructed of polyethylene, they need not be treated for storage. The only stipulation on storage of bladder type cells is that they be carefully wrapped for protection against dust, then be stored at room temperatures in a location shielded from sunlight.

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FUEL ANTISIPHON VALVE (UA-1 AND AFTER; UB-1 AND AFTER) MAINTENANCE PRACTICES 1. PROCEDURES A. Removal The antisiphon valves are located between the fuel cell bladders and the fuel filler adapters. (1) Make sure all electrical power to the airplane is disconnected. (2) Drain the wing fuel cells (Ref. Chapter 12-10-00 DRAINING THE FUEL SYSTEM). (3) Remove the access plates on the lower surface of the wing leading edge. (4) Working through the inboard access opening, remove the clamps from the aft interconnect tube and pull the nipple off the tube. (5) Working through the aft outboard access opening, remove the clamp from the outboard aft interconnect tube and pull the nipple off the tube behind the filler neck. (6) Working through the two access openings on the underside of the wing leading edge, release the fuel cell fasteners. (7) Remove the twelve screws on the top of the wing securing the fuel filler cap assembly. (8) Collapse the cell and move it toward the inboard access opening until access to the fuel filler adapter can be attained through the aft outboard access opening. CAUTION: Do not allow debris from the old gaskets to fall into the fuel bladder. (9) Remove the safety wire and twelve bolts securing the bladder to the fuel filler adapter and antisiphon valve. (10) Break apart the fuel filler adapter and antisiphon valve from the fuel bladder. (11) Remove the fuel filler adapter. (12) Remove the antisiphon valve from the bladder. Remove and discard any gasket material that may adhere to the fuel bladder. (13) Remove the fuel filler adapter and antisiphon valve.

B. Installation (1) Using solvent (30, Table 1, Chapter 91-00-00), clean the sealer from the mating surfaces of the fuel filler adapter and the interior wing skin. NOTE: Insure that the arrow, stamped on the upper surface of the adapter assembly, points forward in the installed position. (2) Check the fuel cell cavity for foreign objects or debris.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (3) Install the two gaskets on the antisiphon valve. Set the antisiphon valve inside the fuel bladder. (4) Set the fuel filler adapter on the antisiphon valve and bladder. Line up the gaskets between the antisiphon valve and fuel cell bladder. Install the twelve bolts and washers. Torque the bolts to 45 to 50 inch-pounds and safety wire. (5) Apply gasket cement (33, Table 1, Chapter 91-00-00) to the flange of the fuel filler adapter where it mounts against the wing skin. (6) Working through the two access openings on the underside of the wing leading edge, secure the fuel filler adapter to the wing skin with the twelve screws. (7) Secure the snap fasteners that hold the fuel cell against the upper wing skin. (8) Working through the lower aft outboard access opening behind the filler neck, install the interconnect tube and clamp. Torque the Wittec Worm Drive (WWD) clamps (Ref. 28-10-00, Table 201). (9) Working through the inboard access opening, install the interconnect tubes (two inboard and one aft). Secure with WWD clamps and torque (Ref. 28-10-00, Table 201). (10) Install the access plates on the lower surface of the wing leading edge.

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28-10-01

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

FUEL FUEL STORAGE (UC-1 AND AFTER) MAINTENANCE PRACTICES

28-11-00 200200

1. PROCEDURES A. Leakage Checks and Repairs (Integral Fuel Cells) To classify the degree of leakage in an integral (wet wing) fuel cell, measure the size of the wet area around the leak. A more accurate measurement may be obtained by wiping the leakage clean and applying talcum powder in the area of the leak. After 30 minutes, check the area to determine if the leak classifies as a stain, seep, heavy seep, or running leak (Ref. Figure 201). Fuel leaks must also be classified as to whether they occur in an open area or in an enclosed area to differentiate between those that require immediate repair and those not considered potential flight hazards. WARNING: Any leakage in an enclosed area, such as the wheel well, or in an area where the fuel will blow into the fuselage, requires grounding until repair is made. (1) Heavy seeps or smaller leaks in an open area may be temporarily repaired any time the airplane is grounded for other maintenance. No action is required at any repair station other than the maintenance base of operations. An external sealant (120, Table 1, Chapter 91-00-00) may be used. NOTE: For ease of sealant removal it is permissible to use sealant emulsifier, PolyGone 300-AG. Do not allow PolyGone 300-AG to come into contact with non-metal parts. Any areas where primer has been removed or dissolved during sealant removal must be cleaned and primed. (2) If leaks are to be repaired from inside the tank, drain all fuel from the affected tank (Ref. Chapter 12-10-00). Remove all damaged sealant from the vicinity of the leak with a wooden or plastic scraper. Chamfer the ends of the remaining sealant so that the new sealant forms a smooth continuous flow. Refer to Table 201 for recommended sealers. (3) If the leakage is around a rivet, strike the rivet. This can only be done once. If the leak persists, replace the rivet. An alternate method for repair of rivet leaks is to install a fuel repair patch (122, Table 1, Chapter 91-00-00). The patch is a thin aluminum disc that is applied directly over the exterior side of the leak and bonded in place with epoxy. Patches are available for flush or protruding head rivets. Because the adhesive used with these patches will cure in the presence of fuel, it is not necessary to drain the tank prior to patch installation. (4) If the leakage is around a bolt with a gasket-type seal, torque the bolt. If the leak persists, replace the seal or the bolt. (5) If the leakage is at the gasket around an access opening or fitting, torque the attachments. Torque access plate bolts 45 to 50 inch-pounds. If the leak persists, replace the gasket. (6) If the leakage is around anchor nuts of access plates on the underside of the wing, seal around the nuts.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL NOTE: If the drain valve was removed during the foregoing checks and repairs, install the valve retaining nut with the castellated portion down, then torque the nut 75 to 100 inch-pounds and safety. Refer to the Model 1900 Airliner Series Structural Repair Manual for wing skin and structural member repairs. Seal repairs as necessary with the appropriate sealer (Ref. Table 201). Table 201 WING FUEL CELL SEALERS TACK FREE/ REFUELING TIME (HOURS)*

CURE/ DISPATCH TIME (HOURS)*

TYPE OF SEAL

ALTERNATE MATERIALS

EC 776 SR 1 Part Material

15 min to 1 hr

24

Coating of Rivets

MIL-S-4383 EC-164

PR-1005-L Buna-N Slosh Coating 1 Part Material

20 min to 1 hr

48

Coating Access Plates

MFG’S NO. & CLASS

APPLICATION/ POT LIFE TIME (HOURS)*

PR-1440 B-1/2

1/2

10

30

PR-1440 B-2

2

36

72

PR-1440 B-4

4

48

96

All Types

MIL-S-8802

PR-1440 A-1/2

1/2

10

30

Seal

PR-1826

PR-1826 B-1/4**

1/4

1 @ 77°F 4 @ 40°F 8 @ 20°F

1 @ 77°F 4 @ 40°F 8 @ 20°F

Spot Sealing

PR-1440

PR-1826 B-1/2**

1/2

2 @ 77°F 8 @ 40°F 16 @ 20°F

2 @ 77°F 8 @ 40°F 16 @ 20°F

Medium Area Sealing

PR-1440

Pro-seal 860

10 min

1.5

4

All types

PR-1826 B-2**

2

9 @ 77°F 36 @ 40°F 72 @ 20°F

9 @ 77°F 36 @ 40°F 72 @ 20°F

Large Areas at Low Temperatures

PR-1440

* Application time is stated as minimum requirement, tack free and cure times are maximum requirements. All time requirements are figured at 77 ± 2°F with 50 ± 5% relative humidly. As a general precaution, for every 10°F rise in temperature from 77°F, reduce application time, tack-free times and cure times by one-half (1/2) and for every 10°F below 77°F, double the stated time. ** To ensure the bond between metal parts and/or existing sealant use primer (supplied with sealant) where required.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 201 Leakage of Integral (Wet Wing) Fuel Cells (UC-1 and After)

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2. FUEL SYSTEM TANK A. Inspection NOTE: It is permissible to use the FUEL TANK INTERNAL INSPECTION procedure in the Model 1900 Airliner Series Corrosion Control Manual (Ref. Chapter 28-10-01). It is also permissible to use the (ALTERNATE TANK INSPECTION METHOD) (UC-1 and After) RADIOGRAPHIC INSPECTION (Ref. Chapter 28-12-00). Microbial or fungal contamination is an increasing threat to fuel tanks. Hormoconis resinae is a virulent acid producing fungus in jet fuel. Unlike typical microbial build-ups, Hormoconis resinae will build up a matte of fibrous material rapidly in areas of condensation, concentrating the acid generating wastes, especially in areas protected from fuel sloshing. The Model 1900C has three such areas in each wing. Two areas are immediately adjacent to both sides (forward and Aft) of the main spar lower cap from WS 124 to WS 130 and one area is in the collector tank. NOTE: Hormoconis resinae may not be detectable in many contamination detector kits available in the United States today. Conidia Bio Sciences has developed and markets a ten minute test kit specifically designed for Hormoconis resinae. For more information go to www.conidia.com on the web. (1) Drain fuel from the collector tanks, main tanks and main tank sumps (Ref. Chapter 12-10-00, DRAINING THE FUEL SYSTEM (UC-1 and After)). (2) Gain access to the collector tanks by removing wing access panels located on top of wing (Ref. Figure 202). (3) Gain access to main tank sump areas by removing wing access panels on underside of wing.

B. Collector Tank Area (1) Inspect collector tanks for evidence of past or present microbiological contamination or sludge buildup. (2) If evidence is found thoroughly clean the collector tank areas with solvent (30, Table 1, Chapter 91-00-00). NOTE: For ease of sealant removal it is permissible to use sealant emulsifier, PolyGone 300-AG. Do not allow PolyGone 300-AG to come into contact with non-metal parts. Any areas where primer has been removed or dissolved during sealant removal must be cleaned and primed. (3) Inspect sealant for condition. Remove any that appears deteriorated. (4) Inspect the collector tank area for evidence of microbiological attack. (5) If evidence of microbiological attack is discovered, contact Hawker Beechcraft Technical Support at 316.676.2000 before taking further action. NOTE: If evidence of microbiological attack is not detected, seal area. Refer to the Model 1900 Airliner Series Structural Repair Manual Chapter 51-30-05 for preparation and sealing instructions.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

C. Main Fuel Tanks, WS 124 thru 130 (L&R) (1) Clean sump areas forward of main spar from WS124 to WS130 with solvent (30, Table 1, Chapter 91-00-00). (2) Inspect main tank sump areas for: (a) Evidence of sealant deterioration. (b) Evidence of past or present microbiological contamination or sludge buildup. NOTE: The initial removal of the sealant from the main spar forward flange, lower cap and the bulkhead at WS 124 thru 130 is required. The sealant removal may be skipped during the 12 month inspection for up to 36 months if the fuel system is sterilized using BIOBOR JF at concentrations of 270 PPM every six months, or Kathon FP 1.5 at concentrations of 135 PPH every six months and is documented in the airplane maintenance records. For application of BIOBOR JF or Kathon FP 1.5 (Ref. Chapter 12-10-00). For ease of sealant removal it is permissible to use sealant emulsifier, PolyGone 300-AG. Do not allow PolyGone 300-AG to come into contact with non-metal parts. Any areas where primer has been removed or dissolved during sealant removal must be cleaned and primed. (3) Remove sealant from the forward flange of the main spar, from the inboard bulkhead from WS 124 to WS 130. (4) Inspect the main spar for evidence of microbiological attack. (5) If evidence of microbiological attack is discovered, contact Hawker Beechcraft Technical Support at 316.676.2000 before taking further action. NOTE: If evidence of microbiological attack is not detected, seal area. Refer to the Model 1900 Airliner Series Structural Repair Manual Chapter 51-30-05 for preparation and sealing instructions. (6) Install all panels removed, fill fuel system (Ref. Chapter 12-10-00, FUEL TANK FILLING) and check for leaks. (7) Sterilize fuel system (Ref. Chapter 12-10-00, FUEL ADDITIVES - BIOCIDAL AGENT). NOTE: To prevent contamination from occurring or returning: •

Drain fuel sumps daily.

•

Periodically sterilize the fuel system every 90 days if evidence of microbiological contamination is detected; annually if no evidence is detected.

•

Clean and inspect collector tanks and main tank sump areas for evidence of microbiological attack or sealant deterioration as required by the Chapter 28 - Fuel System Table in Chapter 5-10-00.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 202 Wing Fuel Tank Access Panels

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

3. FUEL DRAIN VALVE A. Maintenance (1) To service the fuel drain valve (P/N 100-0140-01) push up on the poppet valve with a flat blade screw driver and turn the poppet to the SERVICE position and lower the poppet (Ref. Figure 203). (2) Remove the used packing and install a new packing. (3) Push the poppet up and turn to the closed position.

UE28B 070066AA.AI

Figure 203 P/N 100-0140-01 Fuel Drain Valve

28-11-00

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

FUEL ANTISIPHON VALVE (UC-1 AND AFTER) MAINTENANCE PRACTICES

28-11-01 200200

1. PROCEDURES A. Removal The antisiphon valves are located within each fuel filler adapter. (1) Make sure all electrical power to the airplane is disconnected. (2) Drain the wing fuel cells (Ref. Chapter 12-10-00, DRAINING THE FUEL SYSTEM (UC-1 and After)). (3) Remove the wing access panel which is located below and inboard of the fuel filler adapter (Ref. 28-21-00, ACCESS PANEL REMOVAL). (4) Remove the twelve screws securing the fuel filler adapter assembly. (5) Remove the fuel filler adapter and antisiphon valve through the lower wing access opening.

B. Installation NOTE: For ease of sealant removal it is permissible to use sealant emulsifier, PolyGone 300-AG. Do not allow PolyGone 300-AG to come into contact with non-metal parts. Any areas where primer has been removed or dissolved during sealant removal must be cleaned and primed. (1) Using a plastic scraper, clean the sealant from the mating surfaces of the fuel filler adapter and the interior wing skin. (2) Apply sealant (38, Table 1, Chapter 91-00-00) to the mating surface of the fuel filler adapter. NOTE: Insure that the arrow, stamped on the upper surface of the adapter assembly, points forward in the installed position. (3) Install the fuel filler adapter and antisiphon valve with the twelve screws into the upper wing skin. (4) Install the access cover with a new packing (Ref. 28-21-00, ACCESS PANEL INSTALLATION).

28-11-01

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

28-12-00 200200

FUEL ALTERNATE TANK INSPECTION METHOD (UC-1 AND AFTER) MAINTENANCE PRACTICES

1. RADIOGRAPHIC INSPECTION A. Preparation This inspection is an alternate procedure to the FUEL SYSTEM TANK INSPECTION (Ref. 28-11-00). This alternate procedure allows the inspection to be carried out without removing the sealant along the main spar lower cap from WS 124 to WS 130 in the main tanks. In this procedure, the inspection areas are x-rayed to inspect for corrosion. Only deteriorated sealant must be removed and replaced. Three areas in each wing will be inspected for corrosion caused by fungus growth. These areas are: •

The entire collector tank (Ref. Figure 203).

•

The main tank sump area along the main wing spar from WS 124.58 to WS 135.86 that extends forward to the fuel pickup strainer access panel (Ref. Figure 204).

•

The main tank sump area along the main wing spar from WS 124.58 to WS 135.86 that extends aft to the main tank sump access panel.

(1) Drain fuel from the collector tanks, main tanks and main tank sumps (Ref. Chapter 12-10-00, DRAINING THE FUEL SYSTEM (UC-1 and After)). (2) Gain access to the collector tanks by removing wing access panels (Ref. Figure 204). (3) Gain access to main tank sump areas by removing wing access panels. (4) Thoroughly clean all tank surfaces in the three inspection areas using isopropyl alcohol (30, Table 1, Chapter 91-00-00). Fungus contamination will appear as black sludge (Ref. Figure 203). NOTE: For ease of sealant removal it is permissible to use sealant emulsifier, PolyGone 300-AG. Do not allow PolyGone 300-AG to come into contact with non-metal parts. Any areas where primer has been removed or dissolved during sealant removal must be cleaned and primed. (5) Inspect sealant for condition. Remove any that appears deteriorated. NOTE: All open fuel lines and disconnected fuel system components must be capped and plugged to prevent contamination of the fuel system. (6) Perform the ELECTRIC FUEL BOOST PUMP REMOVAL procedure (Ref. 28-21-00) on both collector tanks. (7) Disconnect the cross transfer fuel lines (4) from the forward end of both collector tanks and leave loose (Ref. Figure 201). (8) Remove the fuel strainers (2) from the forward end of both collector tanks. (9) Remove the fuel probes (3) from the wall of both collector tanks leaving the wires attached and set aside in the tank.

28-12-00

Page 201 Nov 1/12

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (10) Remove the transfer jet pump fuel pickup strainers (Ref. Figure 202). (11) Completely dry out all fuel from the following areas: (a) The entire collector tank (Ref. Figure 203). (b) An area along the main wing spar from WS 124.58 to WS 135.86 that extends forward to the fuel pickup strainer access panel (Ref. Figure 204). (c) An area along the main wing spar from WS 124.58 to WS 135.86 that extends aft to the main tank sump access panel. (12) Allow certified technicians to perform the following procedure and have them check the structures noted in the FUEL SYSTEM TANK INSPECTION (Ref. 28-11-00), and also those noted in the Model 1900 Airliner Series Corrosion Control Manual, FUEL TANK INTERNAL INSPECTION (Ref. Chapter 28-10-01). CAUTION: Ensure no personnel are in the area during the X-ray procedure.

B. Personnel & Safety •

All x-ray producing equipment can only be operated by personnel certified to level 2 or 3.

•

All x-ray producing equipment will be operated in accordance with all state and NCR regulations.

•

Any x-ray assistants must have required safety training.

•

Radiation boundaries must be marked with ropes and signs to keep out unauthorized personnel.

C. Test Equipment •

X-ray tube must have an effective focal spot size of 1.5 mm or less, and be able to obtain a range of 50 to 80 KV.

•

X-ray tubes must be of the portable type.

NOTE: For guidance beyond the scope of this procedure refer to ASTM E 1742.

D. Film (1) Double load film cassettes with fine grain/coarse grain film for complete coverage. (2) Place a letter “B” on the back side of the film to detect backscatter. (3) If back scatter is present, place a 0.005 inch minimum lead screen on the back side to protect film from backscatter.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

E. Film/X-ray Tube Placement and Exposure NOTE: No penetrameters are required for detection of corrosion. (1) Place film on top side of lower wing skin. Refer to Figure 203 for film placement. NOTE: Area to be x-rayed consist of the entire collector tank, the main tank sump area along the main wing spar from WS 124.58 to WS 135.86 that extends forward to the fuel pickup strainer access panel, and the main tank sump area along the main wing spar from WS 124.58 to WS 135.86 that extends aft to the main tank sump access panel., both right hand and left hand sides of the aircraft (Ref. Figure 204). (2) Place film I.D. and lead location markers on the bottom side of the lower wing skin to assure coverage. (3) Place x-ray tube below the lower wing skin, facing up at a 32”-48” tube-to-skin distance, x-ray source must be perpendicular with lower wing skin +/- 10 degrees. (4) Expose film using 60- 75 KV and proper milli-amper-seconds to produce a radiograph of 1.8 to 4.0 H&D density in the area of interest.

F. Indication/Evaluation (1) Any dark pit or irregular shaped indications must be interpreted as corrosion unless positively identified as something else such as adhesive. (2) Any areas not corrosion related such as grinding marks, or drill marks will be evaluated by engineering or NDI level 3. (3) Percentage of wall thickness remaining in an area of corrosion may be estimated by x-ray of a standard of similar material type and areas machined to different thicknesses is used for comparison. (4) If indication is thought to be a film artifact, radiograph must be retaken to verify (5) If corrosion is found, please contact Hawker Beechcraft Technical Support at 316.676.2000 before taking further action. If no corrosion is found proceed to next Step.

G. Completion (1) Repair any sealant that was removed in Step (5). Refer to the Model 1900 Airliner Series Structural Repair Manual, Chapter 51-30-05 for preparation and sealing instructions. (2) Remove all caps and plugs from all opened fuel lines and fuel system components prior to installation. Check for and eliminate all FOD. (3) Install the transfer jet pump fuel pickup strainers (Ref. Figure 202). (4) Install the fuel strainers (2) in the forward end of both collector tanks (Ref. Figure 201). (5) Install the fuel probes (3) to the wall of both collector tanks. (6) Connect the cross transfer fuel lines (4) to the forward end of both collector tanks.

28-12-00

Page 203 Nov 1/12

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (7) Perform the ELECTRIC FUEL BOOST PUMP INSTALLATION procedure (Ref. 28-21-00) on both collector tanks. NOTE: Ensure all fuel lines that were disconnected in previous Steps are properly connected before installing access panels. (8) Check the three inspection areas for FOD before installing access panels (Ref. Figure 203). (9) Perform FUEL TANK FILLING procedure (Ref. Chapter 12-10-00) and check for leaks. NOTE: To prevent contamination from occurring or returning:

Page 204 Nov 1/12

•

Drain fuel sumps daily.

•

Periodically sterilize the fuel system every 90 days if evidence of microbiological contamination is detected (Ref. Chapter 12-10-00).

•

Clean and inspect collector tanks and main tank sump areas for evidence of microbiological attack or sealant deterioration as required by Table 201 (Ref. Chapter 5-10-00).

28-12-00

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 201 Collector Tank Fuel System Components

28-12-00

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 202 Equipment Bay Fuel System Components

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28-12-00

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

1. COLLECTOR TANK 2. FORWARD MAIN TANK SUMP AREA 3. AFT MAIN TANK SUMP AREA

MAIN TANK SUMP ACCESS PANEL (BOTTOM SURFACE)

2

3 1

MAIN TANK SUMP ACCESS PANEL (BOTTOM SURFACE)

COLLECTOR TANK ACCESS PANEL (TOP SURFACE)

UE28B 082444AA.AI

Figure 203 Fuel System Wing Tank Inspection Area

28-12-00

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 204 Wing Fuel Tank Access Panels

Page 208 Nov 1/12

28-12-00

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

28-20-00 00

FUEL FUEL DISTRIBUTION (UA-1 AND AFTER; UB-1 AND AFTER) DESCRIPTION AND OPERATION 1. GENERAL The fuel system consists of two separate systems connected by a cross-transfer system. Each wing system has one filler cap, located in each outboard wing leading edge. Fuel flows by gravity through fuel cell interconnects, from the outboard wing fuel cells to the center section fuel cell and then to the collector tank. A cross-transfer line connects the two collector tanks. A switch-controlled cross-transfer valve is externally connected into the line at the forward outboard corner of the LH center section fuel cell. When the valve is in its normally closed position, each engine draws fuel from its respective fuel tank system. A manually operated cross-transfer control switch is mounted on the upper fuel control panel, just above the fuel quantity gages. When the cross-transfer control switch is actuated, power is drawn from a circuit breaker on the lower fuel control panel to the solenoid of the cross-transfer valve. The cross-transfer valve then opens to allow an electrically-driven fuel pump to transfer fuel to either the left wing or right wing fuel system. The electrically-driven pump located in the bottom of each collector tank is provided as a backup pump should the primary jet pump fail; it is also used for all cross-transfer operations. In the event of an inoperative electric pump, cross-transfer can only be accomplished from the side of the operative pump. The engine-driven fuel pump is mounted on the accessory case in conjunction with the fuel control unit. The engine-driven boost pump is mounted on a drive pad on the aft accessory section of the engine. In the event of failure of the engine-driven boost pump, the respective red FUEL PRESSURE light in the annunciator panel will illuminate. The light illuminates when pressure decreases below 1 psig. The light will be extinguished by switching on the backup boost pump on that side, thus increasing pressure to above 1 psig. CAUTION: Operation with the fuel pressure light on is limited to 10 hours between overhaul or replacement of the engine-driven fuel pump.

28-20-00

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28-20-02 200200

FUEL FUEL FILTERS AND SCREENS (UA-1 AND AFTER; UB-1 AND AFTER) MAINTENANCE PRACTICES 1. PROCEDURES A. Servicing Each engine is equipped with a fuel filter located in the landing gear wheel well, aft of the main spar. The filter elements may be removed and cleaned as outlined in the following procedures.

B. Cleaning Fuel Filters (1) Close the manual shutoff valve. (2) Remove the drain hose from the filter and disconnect the drain line. (3) Cut the lockwire securing the filter case assembly retaining nut. (4) Remove the retaining nut and packing seal. (5) Remove the filter case assembly. (6) Remove the filter element. (7) If an Air-Maze filter is installed, perform AIR-MAZE FUEL FILTER CLEANING procedure in Chapter 28-00-00. Otherwise, proceed to Step (8). (8) Clean the filter element, filter case assembly and mounting hardware with solvent (2, Table 1, Chapter 91-00-00) or its equivalent. NOTE: The maximum permissible pressure drop across the fuel filter after cleaning is 0.6 psig at 2400 pounds per hour fuel flow. (9) Install the filter element. (10) Install the filter case assembly. (11) Install the packing seal and retaining nut. (12) Safety the retaining nut and lockwire. (13) Install the drain hose and connect the drain lines. (14) Open the manual shutoff valve.

C. Engine-Driven Fuel Pump Screens and Filters Refer to the Pratt and Whitney Engine Maintenance Manual for the PT6A-65B engine for proper servicing procedures for the engine-driven fuel pump screens and filters. NOTE: All fuel filters, screens and the fuel tank sump should be cleaned any time the submerged fuel boost pump is removed.

28-20-02

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FUEL FUEL PUMPS (UA-1 AND AFTER; UB-1 AND AFTER) MAINTENANCE PRACTICES

28-20-03 200200

1. ELECTRIC FUEL BOOST PUMP A. Removal (1) Make sure that all electrical power to the airplane is disconnected. (2) Drain the wing fuel cells. Perform DRAINING THE FUEL SYSTEM procedure in Chapter 12-10-00. (3) Remove the large access cover on the bottom of the center section. Remove the access plate from the bottom of the collector tank (Ref. Figure 201). (4) Disconnect the electrical lead from the pump. (5) Disconnect the outlet line from the pump. (6) Remove the bolts and washers securing the pump. (7) Remove the pump from the airplane.

B. Installation (1) Working through the access opening on the bottom of the center section, position the new gasket and pump in the collector tank (Ref. Figure 201). (2) Install the washers and bolts securing the pump. Safety the bolts. (3) Install the outlet line to the pump. (4) Install the electrical lead to the pump. (5) Install the access plate on the bottom of the fuel collector tank and install the attaching washers and bolts. Torque the bolts 20 to 30 inch-pounds and safety wire. (6) Install the access cover on the bottom of the center section.

2. JET TRANSFER PUMP A. Removal (1) Make sure all electrical power to the airplane is disconnected. (2) Drain the wing fuel cells. Perform DRAINING THE FUEL SYSTEM procedure in Chapter 12-10-00. (3) Remove the large access cover on the bottom of the center section. Remove the access plate from the bottom of the collector tank (Ref. Figure 201). (4) The system is equipped with two jet transfer pumps. One is located at the forward end and one at the aft end of the collector tank. (5) Disconnect the fuel lines from the pump.

28-20-03

Page 201 Nov 1/09

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (6) Remove the safety wire, washers and bolts securing the pump and remove the pump from the airplane.

B. Installation (1) Working through the access opening on the bottom of the center section, position the jet pump in the collector tank (Ref. Figure 201). (2) Install the washers and bolts securing the pump and safety wire the bolts. (3) Connect the fuel inlet and outlet lines to the pump. (4) Install the access plate on the bottom of the fuel collector tank and install the attaching washers and bolts. Torque the bolts 20 to 30 inch-pounds and safety wire. (5) Install the access cover on the bottom of the center section.

3. PRIMARY FUEL JET PUMP A. Removal (1) Make sure that all electrical power to the airplane is disconnected. (2) Drain the wing fuel cells. Perform DRAINING THE FUEL SYSTEM procedure in Chapter 12-10-00. (3) Remove the large access cover on the bottom of the center section. Remove the access plate from the bottom of the collector tank (Ref. Figure 201). (4) Disconnect the inlet line from the pump. (5) Remove the safety wire, bolts, and washers securing the pump. (6) Remove the pump from the airplane.

B. Installation (1) Working through the access opening on the bottom of the center section, position the new gasket and pump in the collector tank (Ref. Figure 201). (2) Install the washers and bolts securing the pump. Safety the bolts. (3) Install the inlet line to the pump. (4) Install the access plate on the bottom of the fuel collector tank. (5) Install the access cover on the bottom of the center section.

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28-20-03

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 201 Fuel Supply Collector Tank (UA-1 and After; UB-1 and After)

28-20-03

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FUEL FUEL VALVES (UA-1 AND AFTER; UB-1 AND AFTER) MAINTENANCE PRACTICES

28-20-05 200200

1. CROSS-TRANSFER VALVE The line-mounted cross-transfer valve allows cross-transfer of fuel from either system when circumstances warrant. The external cross-transfer valve is located at the forward outboard corner of the center section fuel cell (Ref. Figure 201).

A. Removal (1) Make sure all electrical power to the airplane is off. (2) Drain the wing fuel cells. Perform DRAINING THE FUEL SYSTEM procedure in Chapter 12-10-00. Drain to a level below the cross-transfer valve. (3) Disconnect the electrical lead from the valve. (4) Disconnect the inlet and outlet lines from the valve. (5) Remove the safety wire and mounting bolts and remove the valve from the mounting bracket.

B. Installation (1) Place the valve in its mounting bracket, then install the mounting bolts and safety wire. (2) Connect the inlet and outlet lines to the valve. (3) Connect the electrical lead to the valve.

2. FIREWALL SHUTOFF VALVES Electrically operated, gate-type shutoff valves are mounted behind the firewall on the outboard side of each nacelle. Relief valves are incorporated in the valves to relieve thermal expansion downstream of the valve (Ref. Figure 202).

A. Removal (1) Make sure all electrical power to the airplane is off. (2) Drain the fuel system. Perform DRAINING THE FUEL SYSTEM procedure in Chapter 12-10-00. Drain to a level below the shutoff valve. (3) Remove the access door on the outboard side of the nacelle, just forward of the wing leading edge. (4) Disconnect the electrical lead from the valve. (5) Disconnect the inlet and outlet lines from the valve. (6) Remove the safety wire and four mounting bolts and remove the valve from the mounting bracket.

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B. Installation (1) Position the valve in its mounting bracket, then install the mounting bolts and safety wire. (2) Connect the inlet and outlet lines. (3) Connect the electrical leads. (4) Replace the access cover.

C. Functional Check (1) Start the engine as outlined in the Model 1900C Airplane Flight Manual. (2) Upon engine stabilization, perform the following check: CAUTION: Observe all fire precautions and safety practices when working on the fuel system. Refer to Chapter 28-00-00. The fire extinguisher system is armed when the FIRE PULL “T” handle is pulled. Actuating the extinguisher push light/switch will release the extinguishing agent. (a) Close left hand firewall shut-off valve by activating (pulling) the FIRE PULL “T” handle. NOTE: If engine fuel flow does not decrease to zero (0) within 10 seconds of fuel flow shut-off valve activation, replace the motive flow check valve and repeat this procedure. (b) 2. Observe that appropriate FUEL PRESS LO annunciator illuminates and engine fuel flow deceases to zero (0) within 10 seconds following activation of fuel firewall shut-off valve. (c) Repeat the above procedure for other engine.

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Figure 201 Cross-Transfer and Electric Pump Schematic (UA-1 and After; UB-1 and After)

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Figure 202 Firewall Shutoff Valves Schematic (UA-1 and After; UB-1 and After)

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FUEL FUEL DISTRIBUTION (UC-1 AND AFTER) DESCRIPTION AND OPERATION

28-21-00 00

1. GENERAL The fuel system consists of two independent sections connected by a cross-transfer line. Fuel is moved from the outboard areas of the main tank to the collector tank by gravity. The fuel from the auxiliary tank is pumped to the collector tank with a manually activated electric pump (Ref. Figures 1, 2 and 3). A cross-transfer line connects the two collector tanks. A switch-controlled cross-transfer valve is externally connected into the line at the forward outboard corner of the LH center section fuel cell. When the valve is in its normally closed position, each engine draws fuel from its respective fuel tank system. A manually operated cross-transfer control switch is mounted on the upper fuel control panel, just above the fuel quantity gages. When the cross-transfer control switch is actuated, power is drawn from a circuit breaker on the lower fuel control panel to the solenoid of the cross-transfer valve. The cross-transfer valve then opens to allow an electrically-driven fuel pump to transfer fuel to either the left wing or right wing fuel system. The electrically driven fuel boost pump, located in the bottom of the collector tank, serves a dual purpose. It transfers fuel from the side on which it is mounted to the collector tank in the opposite wing and serves as a backup for the engine-driven fuel boost pump. Should one of the electric fuel boost pumps malfunction, cross-transfer of fuel can only flow from the side with the functioning fuel pump. CAUTION: Airplane operation with the fuel pressure warning light on is limited to 10 hours before the engine-driven fuel pump must be overhauled or replaced. The engine-driven boost pump is mounted on a pad at the aft accessory section of the engine. Should this pump malfunction, the red L FUEL PRESS or R FUEL PRESS annunciator will illuminate when the pressure drops below 10 psi. The annunciator will remain illuminated until the electric boost pump is switched on and fuel pressure rises above 10 psi.

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Figure 1 Collector Tank Fuel System Components (UC-1 and After) Page 2 Nov 1/09

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Figure 2 Equipment Bay Fuel System Components (UC-1 and After)

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Figure 3 Fuel System Access Covers (UC-1 and After)

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FUEL FUEL DISTRIBUTION (UC-1 AND AFTER) MAINTENANCE PRACTICES

200200

1. ACCESS PANEL A. Removal (1) Drain the fuel system (Ref. Chapter 12-10-00). (2) Insert a 10-32 screw into the threaded hole in the center of the exterior surface of the panel to be removed (Ref. Figure 201). (3) Remove the countersunk screws from the perimeter of the access panel to be removed. (4) Holding the panel with the 10-32 screw inserted in Step (2), push in on the panel to release the seal and rotate the panel as necessary to remove it from the opening in the lower wing skin. (5) Remove the packing seal from the perimeter of the access panel and discard the seal.

B. Installation (1) Insert a 10-32 screw into the threaded hole in the center of the exterior surface of the panel to be installed (Ref. Figure 202). (2) Install a new packing into the groove in the perimeter of the access panel. NOTE: The 10-32 holding screw should be to the outboard end of the panels with the offset center holes. (3) Holding the panel by the 10-32 screw inserted in Step (1), insert the panel into the access opening in the lower wing skin and rotate it into position in the opening. (4) Apply corrosion inhibitive compound (140, Table 1, Chapter 91-00-00) to the countersunk screws and the internal threads around the perimeter of the access panel. NOTE: It is important to tighten access panels in a definite pattern to ensure uniform compression of seals and prevent possible fuel seepage. The example in Figure 202 shows the sequence to be followed when tightening retaining screws in these panels. In the event a panel of different size or a different number of fasteners is encountered, the same basic pattern should be followed and altered to accommodate the differences. When securing the panel, initially snug the screws in the defined sequence, then tighten the screws in the same sequence. (5) Install the countersunk screws around the perimeter of the access panel and tighten the panel using the tightening sequence shown (Ref. Figure 202). Torque the screws 20 to 25 inch-pounds.

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Figure 201 Fuel System Access Covers (UC-1 and After)

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Figure 202 Access Panel Tightening (UC-1 and After)

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FUEL FUEL FITTINGS (UC-1 AND AFTER) MAINTENANCE PRACTICES

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1. PROCEDURES A. Fuel Fittings Any time the clamp-type fittings connecting the fuel lines are loosened, new packings must be installed in the assembly before the fitting is clamped (Ref. Figure 201). When installing these fittings, ensure that the safety spring is seated in the clamp lever. Lubricate all AN type fittings and B-nut seats with an approved antiseize thread compound prior to tightening. Thoroughly inspect all areas of maintenance for leaks and seepage.

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Figure 201 Wiggins Fittings Installation (UC-1 and After)

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FUEL FUEL FILTERS AND SCREENS (UC-1 AND AFTER) MAINTENANCE PRACTICES

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1. PROCEDURES A. Servicing Each engine is equipped with a fuel filter, located in the equipment bay adjacent to the outboard side of the nacelle and forward of the main spar. The filter elements may be removed and replaced or cleaned as outlined in the following Steps (Ref. Figure 201).

2. FUEL FILTER ELEMENT A. Removal and Cleaning (Aircraft Porous Media) (1) Remove the lower panel of the leading edge cover. (2) Cut the safety wire securing the fuel filter shutoff valve control knob and turn the knob 90° to the OFF position (Ref. Figure 2, 28-21-00, Description and Operation section). (3) Cut the safety wire securing the filter bowl nut to the main body of the filter (Ref. Figure 201). (4) Twist the filter bowl counterclockwise until the bowl is free of the filter body. NOTE: These filters are equipped with a bypass indicator button. If this button is extended 3/16 inch, it indicates that the filter element has been bypassed and that there may be some form of fuel contamination or a filter malfunction. This button should be pressed back into position after the filter has been serviced. (5) Remove the filter element from the filter body nipple. Replace or clean the filter element, bowl and interior surfaces of the filter body with cleaning solvent (2, Table 1, Chapter 91-00-00).

B. Installation (Aircraft Porous Media) NOTE: The maximum permissible pressure drop across a clean or new filter is 0.6 psid at 2400 pph fuel flow. (1) Install a new packing in the filter body (Ref. Figure 201). (2) Install the filter element on the filter body nipple. NOTE: Check the filter orifice to assure that the packing is in place. (3) Install the filter element housing. (4) Safety wire the element housing to the filter body. Turn the shutoff valve to the open position and safety wire the knob in place (Ref. Chapter 20-07-00). (5) Check the filter assembly for leaks and seepage. (6) Install the leading edge cover panel.

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C. Removal and Cleaning (Air Maze) (1) Remove the lower panel of the leading edge cover. (2) Cut the safety wire securing the filter fuel shutoff valve control knob and turn the knob 90° to the OFF position (Ref. Figure 2, 28-21-00, Description and Operation section). (3) Cut the safety wire securing the filter bowl nut in place and loosen the filter bowl nut (Ref. Figure 202). (4) Remove the bowl and nut assembly from the filter body. (5) Remove the nut holding the filter element in place and slide the element down and off the mounting stud. (6) Clean the interior surfaces of the filter body and element bowl with solvent (51 or 2, Table 1, Chapter 91-00-00). NOTE: These filters are equipped with a red bypass indicator button. If this button is extended 0.19 inch, it indicates that the filter element has been bypassed and that there may be some form of fuel contamination or a filter malfunction. This button should be pressed back (reset flush) into position after the filter has been serviced. (7) Perform the AIR-MAZE FUEL FILTER CLEANING procedure in this section.

D. Installation (Air Maze) NOTE: The maximum permissible pressure drop across a clean or new filter is 0.6 psid at 2400 pph fuel flow. Torque the nut until it is flush with the filter element and the filter element is snug against the filter body, then tighten the nut 1/4 turn. (1) Position the filter element on the mounting stud and install the retaining nut (Ref. Figure 202). (2) Install a new packing in the filter body groove where the filter body interfaces with the filter bowl. (3) Install the filter bowl and torque the nut to a value of 80 to 120 inch-pounds. (4) Safety wire the bowl retaining nut to the filter body. (5) Turn the filter fuel shutoff valve to ON and safety wire the knob in position (Ref. Chapter 28-21-00). (6) Check the filter assembly for leaks and seepage. (7) Install the leading edge cover panel.

E. Engine-Driven Fuel Pump Screens and Filters Refer to the Pratt and Whitney PT6A-65B Engine Maintenance Manual for engine-driven fuel pump screen and filter servicing intervals and procedures.

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Figure 201 Fuel Filter Assembly (Aircraft Porous Media) (UC-1 and After)

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A

A

RESET POSITION (FLUSH) ACTIVATED POSITION (POPPED OUT 0.19 INCH)

FILTER BYPASS INDICATOR BUTTON VIEW LOOKING DOWN ON TOP OF FILTER

FILTER BODY

FILTER ELEMENT MOUNTING STUD

PACKING FILTER PACK ELEMENT (CONTAINING SIX FILTER DISCS) FILTER BOWL ELEMENT RETAINING NUT

BOWL NUT RETAINER CLIP

SEAL

BOWL RETAINING NUT DRAIN PORT

3.50 DIA. DETAIL

A-A

Figure 202 Fuel Filter Assembly (Air Maze) (UC-1 and After)

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3. AIR-MAZE FUEL FILTER A. Cleaning - Primary Method The primary method of cleaning the Air-Maze fuel filters used on this airplane is by use of an ultrasonic cleaner. The following data provides information on the methods to be used for this purpose.

B. Initial Cleaning - Filter Pack Element (1) Carefully remove the filter pack element from the housing and remove all packings. (2) Do not disassemble the filter pack element for this initial cleaning. (3) Plug open end(s) of filter(s) packs with rubber plugs and immerse in a recommended cleaner (51 or 2, Table 1, Chapter 91-00-00). (4) Allow items to soak for 30 to 60 minutes to remove sludge and baked on contaminates. (5) Agitate parts for several minutes then remove the parts from cleaner. (6) Allow parts to drain and then rinse the parts in new cleaner. (7) Use filtered low pressure air (not exceeding 30 psi) to remove the excess cleaner and to facilitate drying. (8) Preform the FINAL CLEANING - FILTER PACK ELEMENT procedure in this section. NOTE: The ultrasonic cleaning tank used in this procedure must have a minimum of 3 watts of power per square inch of tank bottom area.

C. Final Cleaning - Filter Pack Element (1) Fill tank with a recommended ultrasonic cleaner (51 or 2, Table 1, Chapter 91-00-00). (2) Unplug the filters for this cleaning. (3) Disassemble and remove the filter discs from the center tube of the filter pack element. Slide the discs onto a rod rack which is supported horizontally in the tank. Provide a 1/4 inch space between the filter discs. (4) Metal valve parts, etc., may be cleaned simultaneously in a separate basket suspended in the solution. (5) Clean the filters for 20 minutes. (6) Remove the filter discs from the tank and drain or use filtered low pressure air (not exceeding 30 psi) to remove the excess cleaner and to facilitate drying. (7) Examine filter discs for cleanliness and screen damage. If the screen is damaged, replace the filter disc. Examine parts for damage and replace if required. Inspect the filter discs with a 60X scope to ensure debris removal. (8) Repeat Steps (3) thru (7) if item is not clean. (9) Reassemble the filter pack element.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (10) Replace packings and gaskets (if any) and reassemble valve parts and filter pack element to housing.

D. Cleaning - Secondary Method The secondary method of cleaning the Air-Maze fuel filters used on this airplane is by use of an chemical cleaners. The filter pack elements are cleaned twice when chemicals are used. The following data provides information on the methods to be used for this purpose.

E. Initial Cleaning - Secondary Method - Filter Pack Element (1) Carefully remove the filter pack element from the housing and remove all packings. (2) Do not disassemble the filter pack element for this initial cleaning. (3) Plug open end(s) of filter(s) packs with rubber plugs and immerse in a recommended cleaner (51 or 2, Table 1, Chapter 91-00-00). (4) Allow items to soak for 30 to 60 minutes to remove sludge and baked on contaminates. (5) Remove the filter pack element and parts from cleaner. (6) Allow items to air dry or use filtered low pressure air (not exceeding 30 psi) to remove the excess cleaner and to facilitate drying. (7) Preform the FINAL CLEANING - SECONDARY METHOD - FILTER PACK ELEMENT procedure in this section.

F. Final Cleaning - Secondary Method - Filter Pack Element (1) Immerse the filter pack element in a recommended cleaner (51 or 2, Table 1, Chapter 91-00-00). (2) Agitate solution for twenty minutes. (3) Use a stiff bristled non-metallic brush to remove any remaining dirt or deposits. (4) Unplug the filter and then drain or use filtered low pressure air (not exceeding 30 psi) to remove the excess cleaner and to facilitate drying. (5) Examine filter discs for cleanliness and screen damage. If the screen is damaged, replace the filter disc. (6) Repeat Steps (1) thru (5) if item is not clean. (7) Replace packings and gaskets (if any) and reassemble valve parts and filter pack element to housing.

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FUEL FUEL PUMPS (UC-1 AND AFTER) MAINTENANCE PRACTICES

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1. ELECTRIC FUEL BOOST PUMP A. Removal (1) Disconnect all electrical power to the airplane. (2) Drain the fuel from the wing (Ref. Chapter 12-10-00). (3) Remove the lower outboard nacelle panel that covers a portion of the boost pump access cover. (4) Remove the round access cover from the bottom of the outboard wing (between BL 114.250 and WS 124.588 and stringers K and L) (Ref. Figure 3, 28-21-00, Description and Operation section). (5) Remove the access cover from the upper outboard wing (between BL 114.250 and WS 124.588 and stringers D and E). (6) Disconnect the electrical lead from the electric fuel boost pump (Ref. Figure 1, 28-21-00, Description and Operation section). (7) Disconnect the outlet line from the pump. (8) Working through the lower access port, cut the safety wire and remove the 12 bolts and washers attaching the pump housing to the mounting structure. (9) Remove the pump through the lower access port.

B. Installation NOTE: Ensure that the bond strap is attached to the position from which it was removed. (1) Position the pump in the mounting structure and install the attaching bolts and washers (Ref. Figure 1, 28-21-00, Description and Operation section). (2) Torque the bolts to a value of 45 to 50 inch-pounds. Safety wire the bolt heads (Ref. Chapter 20-07-00). (3) Connect the outlet line to the pump. (4) Connect the electrical lead to the pump. NOTE: Install new packings on the access covers anytime they are removed. (5) Install the access covers. (6) Install the nacelle access panel.

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2. TRANSFER JET PUMP A. Removal (1) Disconnect all electrical power from the airplane. (2) Drain the fuel from the wing (Ref. Chapter 12-10-00). (3) Remove the oval access cover from the bottom of the outboard wing (between WS 124.588 and WS 135.860 and stringers K and L) (Ref. Figure 3, 28-21-00, Description and Operation section). (4) Remove the access cover from the upper outboard wing (between BL 114.250 and WS 124.588 and stringers D and E). (5) Disconnect the fuel lines from the pump (Ref. Figure 1, 28-21-00, Description and Operation section). (6) Remove the bolts and washers securing the adapter and pump to the wing rib. (7) Remove the pump through the upper access port.

B. Installation NOTE: Install a new packing on the adapter prior to installing the pump. (1) Place the pump and adapter in position on the wing rib (Ref. Figure 1, 28-21-00, Description and Operation section). (2) Install the bolts and lock washers attaching the adapter and pump to the wing rib. (3) Attach the plumbing to the pump. NOTE: Install new packings on the access covers anytime they are removed. (4) Install the upper and lower access cover.

3. MAIN TANK JET PUMP A. Removal (1) Disconnect all electrical power on the airplane. (2) Drain the fuel from the wing (Ref. Chapter 12-10-00). (3) Remove the access cover from the upper outboard wing (between BL 114.250 and WS 124.578 and stringers D and E) (Ref. Figure 3, 28-21-00, Description and Operation section). (4) Remove the fuel lines connected to the pump (Ref. Figure 1, 28-21-00, Description and Operation section). (5) Remove the nut and packing holding the pump to the main spar web. (6) Remove the pump through the upper access port.

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B. Installation (1) Install a new packing on the fitting at the inlet (FWD) end of the pump (Ref. Figure 1, 28-21-00, Description and Operation section). (2) Install the pump inlet fitting through the main spar web. (3) Connect the fuel lines to the pump and install the attaching nut. (4) Install the access port cover with a new packing in place.

4. AUXILIARY FUEL TRANSFER PUMP A. Removal (1) Disconnect all electrical power from the airplane. (2) Drain the fuel from the auxiliary tank (Ref. Chapter 12-10-00). (3) Remove the center access cover from the bottom of the center section, aft of the main spar, (between FS 312 and 319 and WS 29.06 and WS 41.85) (Ref. Figure 3, 28-21-00, Description and Operation section). (4) Remove the forward access cover from the bottom of the center section, aft of the main spar, (between FS 304.62 and FS 312 and WS 30.06 and WS 41.85). (5) Working through the forward access port, disconnect the fuel lines from the pump (Ref. Figure 201). (6) Disconnect the electrical wiring. (7) Remove the bolts that attach the pump to the structure. (8) Pull the pump out of the mounting structure.

B. Installation (1) Install a new gasket on the pump and position the pump in the mounting structure (Ref. Figure 201). NOTE: Assure that the ground cable is attached to the area from which it was disconnected. (2) Install the attaching bolts and washers. Torque the bolts to a value of 45 to 50 inch-pounds. Safety wire the bolt heads (Ref. Chapter 20-07-00). (3) Connect the electrical wiring. (4) Connect the fuel lines to the pump. NOTE: Install a new packing on the forward access cover. (5) Install the access covers.

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Figure 201 Auxiliary Fuel Tank (UC-1 and After)

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FUEL FUEL MANIFOLDS (UC-1 AND AFTER) MAINTENANCE PRACTICES

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1. PROCEDURES A. Removal (1) Disconnect all electrical power in the airplane. (2) Drain the fuel from the wing (Ref. Chapter 12-10-00). (3) Remove the access cover from the upper outboard wing (between BL 114.250 and WS 124.588 and stringers D and E) (Ref. Figure 3, 28-21-00, Description and Operation section). (4) Remove the oval access cover from the bottom of the outboard wing (between WS 124.588 and WS 135.860 and stringers K and L). (5) Working through the upper access port, disconnect the fuel lines at the manifold (Ref. Figure 1, 28-21-00, Description and Operation section). (6) Remove the four bolts and washers that attach the manifold to the wing rib. These bolts are reached through the lower access port. (7) Remove the manifold through the upper access port.

B. Installation (1) Position the manifold in place and install the four attaching bolts and washers (Ref. Figure 1, 28-21-00, Description and Operation section). (2) Connect the fuel tubes to the manifold. NOTE: Install new packings on the access covers anytime they are removed. (3) Install the access cover.

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FUEL FUEL VALVES (UC-1 AND AFTER) MAINTENANCE PRACTICES

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1. FILTER FUEL SHUTOFF VALVE A. Removal (1) Disconnect all electrical power from the airplane. (2) Drain the fuel from the wing (Ref. Chapter 12-10-00). (3) Remove the access cover from the upper outboard wing (between BL 114.250 and WS 124.588 and stringers D and E) (Ref. Figure 3, 28-21-00, Description and Operation section). (4) Remove the lower panel of the leading edge cover (on the leading edge adjacent to the outboard side of the nacelle). (5) Disconnect the fuel lines from the forward and aft ends of the shutoff valve (Ref. Figure 1, 28-21-00, Description and Operation section). (6) Remove the nut and sealer from the aft end of the valve and pull the valve forward and out of the structure.

B. Installation (1) Apply sealer (94, Table 1, 91-00-00) to the valve and place the valve in position and install the attaching nut (Ref. Figure 1, 28-21-00, Description and Operation section). (2) Attach the fuel lines to the valve. (3) Connect the fuel lines to the forward and aft ends of the shutoff valve 3. (4) Install the lower panel of the leading edge cover. NOTE: Install a new packing on the upper access cover. (5) Install the upper access cover.

2. FIREWALL FUEL SHUTOFF VALVE A. Removal (1) Remove all electrical power from the airplane. (2) Remove the lower panel of the leading edge cover (on the leading edge adjacent to the outboard side of the nacelle). (3) Cut the safety wire on the filter fuel shutoff valve and move it 90° to the off position (Ref. Figure 2, 28-21-00, Description and Operation section). (4) Disconnect the electrical plug from the firewall shutoff valve. (5) Disconnect the fuel line from the aft side of the fuel filter.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (6) Disconnect the fuel line from the firewall shutoff valve. (7) Cut the safety wire and remove the two bolts securing the fuel filter to its mounting bracket. (8) Remove the filter and shutoff valve. (9) Loosen the jam nut between the filter and the shutoff valve and screw the shutoff valve out of the filter assembly.

B. Installation (1) Install a jam nut on the end of the valve that has the double set of threads (Ref. Figure 2, 28-21-00, Description and Operation section). (2) Move the jam nut to the thread set nearest the valve and install a new packing in the area between the thread sets. (3) Screw the shutoff valve into the filter. The first set of threads on the shutoff valve must be screwed into the filter far enough that the jam nut can be tightened against the filter without pressing the packing against the threads. (4) Install the filter and shutoff valve in the filter bracket with the two attaching bolts and washers. NOTE: The solenoid portion of the shutoff valve must be up. (5) Safety wire the bolts in place (Ref. Chapter 20-07-00). (6) Connect the fuel lines to the filter and shutoff valve. (7) Connect the electrical plug to the shutoff valve. (8) Move the filter fuel shutoff valve to the ON position and safety wire the knob in place. (9) Install the lower panel of the leading edge cover.

C. Functional Check (1) Start the engine as outlined in the Model 1900/1900C Airplane Flight Manual. (2) Upon engine stabilization, perform the following check: CAUTION: Observe all fire precautions and safety practices when working on the fuel system (Ref. Chapter 28-00-00). The fire extinguisher system is armed when the FIRE PULL “T” handle is pulled. Actuating the extinguisher push light/switch will release the extinguishing agent. (a) Close left hand firewall shutoff valve by activating (pulling) the FIRE PULL “T” handle. NOTE: If engine fuel flow does not decrease to zero (0) within 10 seconds of fuel flow shutoff valve activation, replace the motive flow check valve and repeat this procedure.

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(b) Observe that appropriate FUEL PRESS LO annunciator illuminates and engine fuel flow deceases to zero (0) within 10 seconds following activation of fuel firewall shutoff valve. (c) Repeat the above procedure for other engine.

3. CROSS-TRANSFER VALVE A. Removal (1) Disconnect all electrical power from the airplane. (2) Drain the fuel from the airplane (Ref. Chapter 12-10-00). (3) Working through the LH wheel well, disconnect the fuel lines from the valve mounted on the forward side of the main spar. (4) Cut the safety wire from the attaching bolt heads and remove the bolts. (5) Remove the cross-transfer valve from the airplane.

B. Installation (1) Position the valve on its mounting bracket and install the attaching bolts. (2) Safety wire the bolt heads (Ref. Chapter 20-07-00). (3) Connect the fuel lines to the valve. (4) Connect the electrical plug.

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FUEL LH AND RH FUEL LINES (UC-1 AND AFTER) MAINTENANCE PRACTICES

28-21-06 200200

1. PROCEDURES A. Inspection (1) Ensure the landing gear is down, locked and pinned. (2) Open the 2-amp landing gear control circuit breaker. (3) Remove all electrical power from the airplane and disconnect the airplane battery (Ref. Chapter 24). NOTE: If Split Fender Kit 114-9801 is installed only the upper half of the inner fenders require removal (Ref. Chapter 54-10-01). (4) Perform NACELLE INNER FENDER Removal procedure (Ref. Chapter 54-10-01) in the left and right hand wheel wells to gain access to the areas located immediately behind the fenders. (5) Remove the split convolex tubing from the fuel transfer tubes (Ref. SB 32-3839 which supersedes SB 32-3616 R2) before accomplishing fuel transfer tube inspection. WARNING: Wear rubber gloves, goggles and protective wet weather clothing when using MIL-C-85570 cleaning compounds. If cleanser is splashed in eyes, rinse thoroughly with fresh water for 15 minutes and seek medical aid. Remove clothing saturated with cleaning solution immediately and flush exposed skin areas with fresh water. (6) Clean the 118-920000-29 LH and 118-920000-35 RH fuel lines using MIL-C-85570 Type II cleanser (138, Table 1, Chapter 91-00-00). Mix one part cleanser to four parts water and apply with soft brush or cloth for one minute. Rinse with fresh water and dry. (7) Inspect the LH/RH fuel line for chafing. Pay particular attention to the areas located behind the openings for the main landing gear uplock switch wire harnesses. (8) If chafing has occurred, measure the depth of the chafing to determine if additional action is required (Ref. Table 201). (9) If there is no visible chafing to the fuel lines, measure the distance between the fuel lines and the main landing gear uplock switch wire harness. If the distance measured is less than 1/2 inch, complete Steps 1. and 2.. (a) Using tie wraps (139, Table 1, Chapter 91-00-00), remove excess slack from the main landing gear uplock wire harnesses by securing the harnesses to existing wire harnesses which route thru the top of the main landing gear wheel well side panels. (b) Ensure a minimum of 1/2 inch clearance between the main landing gear uplock switch wire harnesses and all tubing located in the main landing gear wheel wells. (10) If the distance measured between the main landing gear uplock switch wire harness is greater than 1/2 inch, install the inner fenders assuring adequate clearance between all wiring and tubing.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL NOTE: If Split Fender Kit 114-9801 is installed and only the upper halves of the inner fenders were removed, install inner fender upper halves (Ref. Chapter 54-10-01). Main gear up position switch adjustment is not required if the following conditions are met: •

Split Fender Kit 114-9801-0001 is installed.

•

Only the upper half of the inner fenders were removed.

•

Main gear up position switch was not disturbed.

(11) Perform NACELLE INNER FENDER Installation procedure (Ref. Chapter 54-10-01) in the left and right hand wheel wells. Table 201 Action Required for Fuel Line Chafing CHAFING DEPTH (INCH)

ACTION REQUIRED

0.001 to 0.007

Perform the CORROSION PREVENTION FOR CHAFED FUEL LINES procedures in this section. Ensure the main landing gear wire harness is secure and there is a minimum clearance of 1/2 inch distance between the fuel line and the wire harness.

0.008 to 0.015

Replace the fuel line before the next 100 hour interval.

0.016 or Greater

Replace the fuel line before the next flight.

B. Corrosion Prevention for Chafed Fuel Lines (1) Mask off areas adjacent to the chafed area using MIL-T-23397 masking tape (136, Table 1, Chapter 91-00-00). (2) Remove all corrosion using 240 grit, P-C-451 Type 1 aluminum-oxide cloth (134, Table 1, Chapter 91-00-00). If the tube assembly is painted, the corroded area and the immediate surrounding surface area must be sanded to bare metal. (3) Polish the affected area using 320 grit, P-C-451 Type 1 aluminum-oxide cloth (135, Table 1, Chapter 91-00-00). If the surrounding surface area is painted, feather the edges of the paint. (4) Measure the depth of the chafing after all corrosion has been removed (Ref. Table 201). If the depth of the affected area is 0.016 inches or greater, discontinue this corrosion removal procedure and replace the fuel line before the next flight. WARNING: Wear rubber gloves, goggles and protective wet weather clothing when using MIL-C-85570 cleaning compounds. If cleanser is splashed in eyes, rinse thoroughly with fresh water for 15 minutes and seek medical aid. Remove clothing saturated with cleaning solution immediately and flush exposed skin areas with fresh water. (5) Wipe the affected area clean using a clean cloth dampened with MIL-C-85570 Type II cleanser (138, Table 1, Chapter 91-00-00). Mix one part cleanser to nine parts water. (6) Rinse with water. Ensure all traces of oil and finger prints have been removed from the affected area. If traces of oil contaminants still exist, repeat Steps (5) and (6).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL WARNING: MIL-C-81706 is moderately toxic to skin, eyes, and the respiratory tract. Eye and skin protection are required. General ventilation is adequate. (7) Surface treat the affected area with MIL-C-81706, Class 1A, Form III (137, Table 1, Chapter 91-00-00) for two to four minutes using a sponge stick applicator. (8) Lightly rinse the affected area with a spray bottle and clean water immediately after applying the surface treatment. (9) Allow to dry for a minimum of 1 hour. (10) Apply one coat of MIL-P-23377 epoxy primer (5, Table 1, Chapter 91-00-00). Wait thirty minutes and apply a second coat of the epoxy primer. (11) If the fuel transfer tube assembly is painted, apply two coats of Matterhorn White No. 6160 urethane paint (133, Table 1, Chapter 91-00-00).

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28-40-00 00

FUEL FUEL QUANTITY INDICATING (UA-1 AND AFTER; UB-1 AND AFTER) DESCRIPTION AND OPERATION 1. GENERAL A. Capacitance Fuel Quantity Indicating System The fuel quantity indicating system is a capacitance type that is compensated for specific gravity and reads in pounds on a linear scale. Electronic circuits in the system process the signals from the fuel quantity (capacitance) probes in the various fuel cells for an accurate readout by the fuel quantity indicators (Ref. Figure 1). Each side of the airplane has an independent gaging system, consisting of a fuel quantity (capacitance) probe in the collector tank, one in the aft inboard fuel cell, two in the integral (wet wing) fuel cell, two in the inboard leading edge fuel cell, and two in the center section fuel cell. Fuel density and dielectric constant vary with respect to temperature, fuel type and fuel batch. The capacitance gaging system is designed to sense and compensate for these variables. The fuel quantity probe is simply a variable capacitor comprised of two concentric tubes. The inner tube is profiled by changing the diameter as a function of height so that the capacitance between the inner and outer tube is proportional to the tank volume. The tubes serve as fixed electrodes and the fuel of the tank in the space between the tubes acts as the dielectric of the fuel quantity probe. The capacitance of the fuel quantity probe varies with respect to the change in the dielectric that results from the ratio of fuel to air in the fuel cell. As the fuel level between the inner and outer tubes rises, air with a dielectric constant of one is replaced by fuel with a dielectric constant of approximately two, thus increasing the capacitance of the fuel quantity probe. This variation in the volume of fuel contained in the fuel cell produces a capacitance variation that is a linear function of that volume. This function is converted to a linear current that energizes the fuel quantity indicator. Capacitance fuel probes are designed to produce a capacitance variation that is linear in relation to variation in weight, even though the weight is nonlinear with respect to the fuel level. By varying the capacitance per inch, the nonlinear level signal is changed to a linear function. In addition to its capacitance sensing tubes, each fuel quantity probe contains a small circuit network that produces an output current whose average value is directly proportional to fuel level, while automatically compensating for fuel temperature-density variations. Regulated +28 vdc is impressed across the low Z side (outer tube) of the fuel quantity probe. The ensuing signal is further processed by a DC amplifier, inside the indicator, which contains a potentiometer for adjusting the “Full” and “Empty” settings. The DC amplifier controls response time and drives a D'Arsonval meter (Gull Airborne) or a servo-type meter (Regen Data Systems). Proper operation of capacitance fuel quantity indicating systems depends on good wire connections. The Fuel Quantity Test Set (14, Table 7, Chapter 91-00-00) and the Probe Selector Unit (15, Table 7, Chapter 91-00-00), are used to check the operation and calibration of the fuel quantity system.

B. Fuel Quantity Capacitance Probes Each capacitance probe consists of two concentric tubes. The tubes serve as fixed electrodes and the fuel of the tank in the space between the tubes acts as the dielectric of the probe. The capacitance of the probe varies with respect to the change in the dielectric that results from the ratio of fuel to air in the tank. The tubular capacitor elements of each capacitance probe are profiled to match the contour of the tank in which it is installed to provide linear capacitance versus volume in the tank at varying flight attitudes.

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C. Low Fuel Quantity Warning System Sensors in the center section tanks provide for low fuel quantity annunciation by sensing the level of fuel in the tank and transmitting energizing current for the annunciator lights. Sensors in the collector tanks provide the signals for low fuel feed annunciation in the same manner. The system is functionally tested by pressing the test switch on the warning annunciator panel mounted on the glareshield. The low fuel quantity sensor and the low fuel feed sensor are connected electrically in parallel at the 28-vdc-in lead, test signal lead and ground lead. The 28-vdc-out leads are connected individually to their respective annunciator circuits. Five-ampere circuit breakers provide operating current for the RH and LH systems.

Figure 1 Low Fuel Quantity Sensors Page 2 Nov 1/09

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D. Low Fuel Quantity Sensors The low fuel quantity sensors take advantage of a principal which allows light, produced within the sensor by a reference light source, to pass through the prism of the tip of the sensor when the prism is submerged in fuel. When the prism is not submerged, the light is reflected off of the bevel of the prism back into the probe. The reflecting light is then sensed by a light sensitive transistor which conducts electrically in response to the light and triggers the current to allow the annunciator to illuminate. An integral self-test circuit provides for functionally testing the sensor and its associated circuitry by energizing another light source (test light source) in the probe which is positioned to shine into the transistor even though the prism of the sensor is submerged in fuel (Ref. Figure 1). When the warning annunciator test switch is pressed, the annunciators will not illuminate immediately, due to a time delay circuit incorporated into the electronics of the sensor itself. The time delay has been designed to allow up to a 4-second delay before the annunciators will be illuminated. After illuminating, the annunciators will remain on for up to 15 seconds. These time delays prevent continuous cycling of the annunciators when the fuel is sloshing around the inside of the tanks. As a result of design, the low fuel sensors are self-diagnosing. Circuit failures within the sensor will always be in the “on” mode; therefore, a false annunciation will result from circuit failure. A failure of the reference light source will cause the light sensitive transistor to be biased, allowing it to conduct and resulting in a false annunciation. Failure of the test light source in the sensor will result in the annunciator lights not being illuminated during test.

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100100

FUEL FUEL QUANTITY INDICATING (UA-1 AND AFTER; UB-1 AND AFTER) TROUBLESHOOTING 1. PROCEDURES Troubleshooting of the fuel quantity indicating system requires the use of special tools available from Hawker Beechcraft Corporation Parts and Service Operations. The Fuel Quantity Test Set (14, Table 7, Chapter 91-00-00) simulates the signals necessary to check and calibrate the fuel quantity indicating system. The Probe Selector Unit (15, Table 7, Chapter 91-00-00) is used with the test set to check probes and probe harnesses in groups or individually. One 9-vdc battery inside the test set provides the power required to read insulation (conductance) and capacitance. The LCD shows capacitance in picofarads (pF) and conductance in nano siemens (nS). Aircraft battery power or an auxiliary power source is required to simulate capacitance for indicator checking and calibration. Troubleshooting checks are carried out in five phases: checking the system harness insulation, probe harness insulation, system capacitance, probe capacitance and checking the indicator. If the indicator does not check or is replaced, it must be calibrated.

A. System Insulation Check NOTE: Connect a regulated 28-vdc external power supply to the airplane when troubleshooting or calibrating. This test procedure may be performed with full, partially full or empty tanks. Display readings are presented in nano Siemens (nS), equivalent megohms are derived by dividing 1000 by the LCD display reading: 1000/50 nS = 20 Megohms. These checks are identical for left and right fuel systems. (1) Place the Fuel Quantity Test Set (14, Table 7, Chapter 91-00-00) in a convenient location for the procedures that follow. (2) Ensure that the test set ON/OFF switch is in the OFF position. CAUTION: Power to the airplane must be off when lowering or raising the fuel control panel to preclude the possibility of shorting any wiring connections to ground. (3) Ensure that the battery switch is OFF. (4) Open the fuel quantity system circuit breakers. (5) Gain access to the fuel quantity indicator. (6) Disconnect the airplane's harness from the fuel quantity indicator. (7) Connect the test set to the airplane's harness. NOTE: Do not connect the test set to the fuel quantity indicator. (8) Rotate the TEST FUNCTION selector to IND AMP. (9) Place the INSULATION/SYSTEM switch to IN nS. (10) Rotate the INS TEST POINTS selector to LO-Z GND.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (11) Place the shorting switch to the NO position. (12) Place the ON/OFF switch to ON and allow sufficient time for the LCD to stabilize. (13) The LCD reading should be less than 50 nS. (14) Rotate the INS TEST POINTS selector through the remaining positions. Allow sufficient time for the LCD to stabilize in each position. The LCD reading in each position should be less than 50 nS, except in the SIG/RTN position which will show an overrange indication (1---). NOTE: An overrange reading will display in the LO-Z/RTN position if a 114-389001-17 probe is installed in the collector tank. (15) Rotate the INS TEST POINTS selector to RTN/SIG. (16) Close the fuel system circuit breakers that were opened in Step (5). (17) Turn the battery master switch to ON. The maximum time delay from power OFF to power ON is 10 seconds. (18) Rotate the INS TEST POINTS selector through the remaining positions. Allow sufficient time for the LCD to stabilize in each position. The LCD reading in each position should be less than 50 nS, with the same exceptions as previously noted. Due to continuity through the diodes inside the probe, when the test set detects a short from RTN to GND, that short will also be seen as a short from SIG to GND. (19) When any of the above checks are out of tolerance, isolate the airplane's wire harness from the probe wire harnesses. The main tank wire harness connectors are behind access panel 4 (Ref. Chapter 6-50-00 WING ACCESS PANELS). (20) Repeat Steps (8) through (18). Acceptable readings are 5 nS or less. If the insulation readings are acceptable, the fault lies in the main tank wire harness to the probes. If not, the fault is in the airplane's wire harness between the indicator in the cockpit out to the wing. (21) Perform the PROBE INSULATION TEST procedure in this section when any faults have been isolated to the probe wire harnesses. NOTE: The results of the preceding checks must be within the stated limits. If any results are slightly outside of these limits, repeat the entire procedure to ensure that failure of the checks was not due to human error. (22) Turn the test set OFF and disconnect the test set. (23) Connect the airplane's wire harnesses to the original configuration.

B. Probe Insulation Test The probe insulation check may be performed with the fuel tanks full, partially full or empty. (1) Ensure that the battery master switch is OFF. (2) Open all fuel system circuit breakers. (3) Locate the Fuel Quantity Test Set (14, Table 7, Chapter 91-00-00) conveniently for working in the area of the access panels for the main tank wire harness connectors.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (4) Disconnect the main tank wire harness from the airplane's wire harness. The main tank wire harness connectors are behind access panel 4 (Ref. Chapter 6-50-00 WING ACCESS PANEL). (5) Ensure that the test box ON/OFF switch is in the OFF position. (6) Connect the Probe Selector Unit (15, Table 7, Chapter 91-00-00) to the bulkhead connector of the main tank wire harness. (7) Connect the test set to the probe selector unit with the ACFT connector. NOTE: Do not connect the test set IND connector to the airplane's wire harness. (8) Rotate the TEST FUNCTION switch to IND AMP. (9) Place the INSULATION/SYSTEM switch to IN nS. (10) Rotate the INS TEST POINTS selector to LO-Z GND. (11) Rotate the probe selector unit selector switch to ALL. (12) Place the test set ON/OFF switch in ON and allow sufficient time for the LCD to stabilize. (13) The LCD reading should be less than 50 nS. (14) Rotate the INS TEST POINTS selector to its remaining positions, allowing sufficient time for the LCD to stabilize in each position: the SIG/RTN position should show an overrange indication (1---); all others should indicate less than 50 nS; however, if a 114-389001-17 probe is installed in the collector tank, the LO-Z/RTN position will show an overrange indication. (15) Using the switch on the selector unit, select each individual probe and repeat Steps (8) through (14). The reading for each probe should be less than 50 nS, and the total readings for all the probes must not exceed 50 nS except in the SIG/RTN position, which should show an overrange indication. When the collector tank probe is selected, an overrange indication will appear if a 114-389001-17 probe is installed. (16) Should any probe be out of tolerance, that probe or its associated wiring to the bulkhead connector must be replaced. (17) Once the fault has been isolated and corrected, the applicable portion of the preceding checks should be repeated to ensure fault correction. NOTE: If any component has been changed, perform the SYSTEM AND PROBE CAPACITANCE CHECK and INDICATOR CALIBRATION procedure in this section. (18) Place the test set ON/OFF switch in the OFF position, and restore the airplane to its original condition.

C. System and Probe Capacitance Check (1) Ensure that the battery master switch is OFF. (2) Open all fuel system circuit breakers.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (3) Locate the Fuel Quantity Test Set (14, Table 7, Chapter 91-00-00) and Probe Select Unit (15, Table 7, Chapter 91-00-00) conveniently for working in the area of the access panels for the main tank wire harness connectors. (4) Disconnect the main tank wire harness from the airplane's wire harness. The main tank wire harness connectors are behind access panel 4 (Ref. Chapter 6-50-00 WING ACCESS PANELS). (5) Ensure that the test box ON/OFF switch is in the OFF position. (6) Connect the probe selector unit to the bulkhead connector of the main tank wire harness. (7) Connect the test set to the probe selector unit with the ACFT connector. NOTE: Do not connect the test set IND connector to the airplane's wire harness. (8) Rotate the TEST FUNCTION selector to PROBES. (9) Place the INSULATION/SYSTEM switch to SYSTEM. (10) Place the MAIN/TOT-AUX/NAC to MAIN/TOT. (11) Place the 200 pF/1000 pF switch to 200 pF if the tanks are empty or to 1000 pF if the tanks are full. (12) Rotate the probe selector unit selector switch to ALL. (13) Place the test set ON/OFF switch in the ON position. (14) Depress and hold the PRESS TO READ CAP (pF) push button: the reading should be within the ranges shown in Tables 101 and 102. Record these empty and full capacitance values in the airframe log for future reference. (15) Place the 200 pF/1000 pF switch to 200 pF. (16) Rotate the selector switch to each probe position and measure the capacitance of the individual probes: the capacitance readings should be within the values listed in Table 103; otherwise, replace the out-of-tolerance probe or its associated wiring, then repeat the appropriate checks to confirm that the fault has been corrected. NOTE: If any component or components have been replaced, perform SYSTEM AND PROBE CAPACITANCE CHECK procedure in this section. (17) Place the test set ON/OFF switch in the OFF position and restore the airplane to original configuration.

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Table 101 EMPTY SYSTEM CAPACITANCE VALUES - MAIN TANK Nom.

Min.

Max.

157.3

141.6

173.0

Table 102 FULL SYSTEM CAPACITANCE VALUES - MAIN TANK Nom.

Min.

Max.

259.2

233.3

285.1

Table 103 INDIVIDUAL PROBE CAPACITANCE VALUES Empty Tanks

Full Tanks

Tank Location

Probe P/N

Min.

Max.

Min.

Max.

Integral Outboard

114-389001-29/-31

17.91

21.89

26.4

32.3

Integral Inboard

114-389001-25/-27

20.03

24.48

31.2

38.1

37.94

46.37

57.6

70.4

Integral Combined Leading Edge Outboard

114-389001-1/-3

21.01

25.67

32.1

39.2

Leading Edge Inboard

114-389001-9/-11

24.75

30.25

40.3

49.2

Box Section

114-389001-13/-15

15.89

19.43

24.4

29.9

Center Section Outboard

114-389001-5/-7

19.82

24.22

39.1

47.8

Center Section Inboard

114-389001-21/-23

14.91

18.23

29.1

35.6

Collector Tank

114-389001-17/-19

7.24

8.84

10.5

12.9

Table 104 CAPACITANCE SIMULATION VALUES - MAIN TANK 0 lbs

300 lbs

600 lbs

900 lbs

1200 lbs

189.0

210.0

231.0

252.0

273.0

D. Setting Capacitance Simulation When performing the indicator linearity check or calibrating the indicator, the test set uses the airplane's 28 volts and it's circuitry to simulate a capacitance for the indicator to read. This value is shown on the indicator face as a fuel quantity. During a bench test the procedures for simulating capacitance is the same; the only difference being an auxiliary 28 vdc power supply is used in lieu of the airplane's power. Use these procedures and Table 104 as required to set capacitance values for checking and calibrating the indicator.

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NOTE: Connect the Fuel Quantity Test Set (14, Table 7, Chapter 91-00-00) ACFT connector to the airplane’s wire harness and the IND connector to the indicator. (1) Set the TEST FUNCTION selector switch to CAP SIM CAL (NORM SYS). (2) Set the INSULATION/SYSTEM switch to SYSTEM. (3) Set the MAIN/TOT-AUX/NAC switch to MAIN/TOT. (4) Set the shorting switch to NO. (5) Set the pF switch to 200 for values from 0 - 200 pF or in the 1000 pF position for values over 200. (6) Turn the test set ON. (7) Set the CAP SIM (pF) 100's and 10's thumb wheels to the desired capacitance value to be simulated. (8) Press and hold the PRESS TO READ CAP (pF) button to read the capacitance in the LCD. (9) Use the trimmer control to fine adjust the capacitance setting. (10) Set the TEST FUNCTION switch to IND AMP. (11) Turn on the battery switch and apply auxiliary power (28 vdc) to airplane. (12) The fuel quantity indicator will read the capacitance simulated by the test set and display a fuel quantity in pounds.

E. Indicator Linearity Check The following linearity check may be performed on the fuel quantity indicator with the fuel tanks full, partially full or empty. (1) Ensure that the battery master switch is OFF. (2) Locate the Fuel Quantity Test Set (14, Table 7, Chapter 91-00-00) convenient to the fuel control panel on the pilot's sidewall. (3) Open the appropriate fuel quantity system circuit breakers. (4) Gain access to the fuel quantity indicator. (5) Connect the test set ACFT connector to the indicator wire harness and the IND connector to the indicator. (6) Set the test set to the values shown in Table 104. Refer to SETTING CAPACITANCE SIMULATION procedure in this section. (7) Set the indicator needle deflection, using the EMP 1/E1 or E adjustment, for each required quantity shown in Table 104. (8) Set the indicator needle deflection by using the EMP 1/E1 or E adjustment to the required quantity.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (9) If the indicator will not adjust to the required quantity reading for the capacitance simulated, it must be replaced. (10) Turn the battery switch OFF. (11) Turn the test set OFF. (12) Disconnect the test set. (13) Connect the airplane's wire harness to the indicator and return the airplane to the original configuration.

F. Indicator Calibration The fuel quantity indicator may be calibrated with the fuel tanks either completely empty (undrainable fuel only) or completely full; however, the preferred procedure is to calibrate the indicator with the tanks completely empty. While calibrating the indicator with empty tanks is more reliable than the alternate procedure, the accuracy of the alternate procedure (full tanks) can be appreciably enhanced if the empty-tanks value is known. It is recommended that when the empty capacitance value is determined, it should be recorded in the airframe log for future use in calibrating. The preferred procedure (undrainable fuel only) makes use of the actual empty-tank value in adjusting the zero indication. The alternate procedure (full tanks) applies a nominal empty-tank value to the system by the test set for the zero indication. Both procedures will be outlined in the following text; however, it must be remembered that the alternate procedure is a stop-gap procedure only and the system should be recalibrated using the preferred procedure at the earliest convenience of the operator.

G. Empty Tanks Calibration (Preferred Procedure) (1) Defuel the airplane and drain sumps. (2) Locate the Fuel Quantity Test Set (14, Table 7, Chapter 91-00-00) conveniently for working in the cockpit. (3) Ensure that the ON/OFF switch on the test set is OFF. (4) Ensure that the battery switch is off and the fuel quantity circuit breakers are open. (5) Gain access to the fuel quantity indicator. (6) Disconnect the wiring connector from the indicator. (7) Connect the test set ACFT connector to the indicator wire harness and connect the test set IND connector to the indicator. (8) Rotate the TEST FUNCTION selector to CAP SIM CAL (NORM SYS). (9) Place the INSULATION/SYSTEM switch to SYSTEM. (10) Place the MAIN/TOT-AUX/NAC switch to MAIN/TOT. (11) Place the shorting switch to the NO position.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (12) Place the ON/OFF switch to ON. (13) Close the appropriate fuel quantity circuit breakers. (14) Turn the battery switch ON. NOTE: There are two different indicators which may be installed in the Model 1900 Series Airplanes: one type of indicator has a power-off dot below the “0” graduation. (15) The fuel quantity indicator needle should rest between the “0” indication and the power-off dot on indicators which have the power-off dot, or the needle should rest approximately one needle width below the “0” indication on indicators without the power-off dot. Should the indication be inconsistent with the preceding guidelines, adjust the EMP 1/E1 or E/EMP adjustment on the back of the indicator. (16) Use the procedures in SETTING CAPACITANCE SIMULATION to set a capacitance value on the test set as follows: ADD CAP = 104.0 pF. (17) Rotate the TEST FUNCTION selector to ADD CAP. (18) The value set on the test set is added to the current empty tank reading which will simulate a full tank. Adjust the F/FULL adjustment on the indicator to read the maximum fuel quantities shown in Table 104. NOTE: Check empty and full adjustments made to the indicator to ensure accurate calibration. (19) Turn the battery switch OFF. (20) Turn the ON/OFF switch on the test set OFF and disconnect the test set from the airplane's system. (21) Restore the airplane to original configuration.

H. Full Tanks Calibration (Alternate Procedure Only) With tanks FULL use Steps (1) through (20), then (24) to end. When tanks are PARTIALLY FULL use Steps (1) through (19), then Steps (21) to end. (1) Locate the Fuel Quantity Test Set (14, Table 7, Chapter 91-00-00) conveniently for working in the cockpit. (2) Ensure that the ON/OFF switch on the test set is OFF. (3) Ensure that the battery switch is OFF and the fuel quantity circuit breakers are open. (4) Gain access to the fuel quantity indicator. (5) Disconnect the wiring connector from the indicator. (6) Connect the test set ACFT connector to the indicator wire harness and connect the test set IND connector to the indicator. (7) Rotate the TEST FUNCTION selector to CAP SIM CAL (NORM SYS). (8) Place the INSULATION/SYSTEM switch to SYSTEM.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (9) Place the MAIN/TOT-AUX/NAC switch to MAIN/TOT. (10) Place the shorting switch to the NO position. (11) Place the ON/OFF switch to ON. (12) Close the appropriate fuel quantity circuit breakers. (13) Turn the battery switch ON. (14) Set the test set to a nominal empty capacitance value for the main tank as shown in Table 101. Check the airframe log for an empty tank value and, if available, use that value in place of the values given in Table 101. (15) Rotate the TEST FUNCTION switch to IND AMP. (16) Turn the battery switch ON. (17) The indicator needle should indicate between the “0” position and the power-off dot for indicators with the power-off dot or approximately one needle width below the “0” indication on indicators without the power-off dot. Should the indicator reading be inconsistent with the preceding guidelines, adjust the EMP 1/E1 or E/EMP adjustment on the back of the indicator. (18) Set the fuel select switch to MAIN. (19) Rotate the TEST FUNCTION switch to CAP SIM CAL (NORMAL SYS). (20) The indicator should read the maximum value for a full tank (1447 lbs). Make adjustments with the F/FULL adjustment as required. (21) Set the test set to a nominal full capacitance value for the main tank as shown in Table 102. Check the airframe log for an empty tank value and, if available, use that value in place of the values given in Table 101. (22) Rotate the TEST FUNCTION switch to IND AMP. (23) The indicator should read 1400 lbs. NOTE: Check the empty and full adjustments made to the indicator to ensure accurate calibration. (24) Place test set ON/OFF switch in the OFF position. (25) Turn the battery master switch OFF. (26) Open the appropriate fuel quantity circuit breakers. (27) Disconnect the test set from the airplane's system. (28) Restore the airplane to the original configuration.

I. Fuel Quantity Probe Bench Check Should a fuel quantity probe become suspect as a result of the preceding checks, the following bench check procedure must be performed on the probe prior to returning the probe to Hawker Beechcraft Corporation.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (1) Ensure that the ON/OFF switch on the test set is in the OFF position. (2) Rotate the TEST FUNCTION selector to PROBES. (3) Place the INSULATION/SYSTEM switch to SYSTEM. (4) Place the 200 (pF)/1000 (pF) switch to 200 (pF). (5) Connect the probe adapter P/N 101-00814 to the PROBES socket on the test set. (6) Connect the color-coded leads from the adapter to the matching color and threads of the probe terminals as follows: L-YELLOW (10-32), S-RED (8-32), R-GREEN (6-32). Connect the ground clip to the flange of the probe and place the shorting switch in the YES position. NOTE: The grounding clip is not used when testing an integral tank probe that does not have a flange. The shorting switch is still placed in the YES position. (7) Place the ON/OFF switch on the test set in the ON position. (8) Depress the PRESS TO READ CAP (pF) push button to read the probe capacitance on the LCD allowing sufficient time for the LCD to stabilize. (9) The probe capacitance value must be within the tolerance as specified by part number in Table 103. (10) If the probe capacitance is not within tolerance, the probe should be replaced. (11) Rotate the TEST FUNCTION selector to IND AMP. (12) Place the INSULATION/SYSTEM switch in the IN nS position. (13) Place the shorting switch to the NO position. (14) Rotate the INS TEST POINT selector to LO-Z GND. (15) Allow sufficient time for the LCD to stabilize: the LCD reading must be less than 50 nS. (16) Rotate the INS TEST POINT selector to each of the remaining position, allowing sufficient time for the LCD to stabilize in each position: all readings must be less than 50 nS, except in the SIG/RTN position which will display and overrange indication (1---). NOTE: An overrange indication should also be present in the LO-Z/RTN position if a 114-389001-17 probe is being tested. (17) Should any of the preceding indications be inconsistent with the expected results, the probe should be replaced. (18) Place the ON/OFF switch on the test set in the OFF position and disconnect the test set adapter from the probe.

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FUEL FUEL QUANTITY INDICATING (UA-1 AND AFTER; UB-1 AND AFTER) MAINTENANCE PRACTICES 1. FUEL PROBE - INBOARD LEADING EDGE AND INBOARD AFT FUEL CELLS A. Removal (1) Remove the access cover over the fuel quantity probe (Ref. Figure 201). CAUTION: Handle the fuel quantity probe carefully, for damage to the surface of the probe tubes will destroy the accuracy of the probe as a measuring device. (2) Unsafety and remove the five mounting screws and remove the fuel quantity probe from the fuel cell. (3) Disconnect the electrical leads at the three terminals on the side of the fuel quantity probe.

B. Installation CAUTION: Handle the fuel quantity probe carefully, for damage to the surface of the probe tubes will destroy the accuracy of the probe as a measuring device. (1) Attach the lead wires to the fuel capacitance probe (Ref. Figure 201). (2) Insert the fuel quantity probe into its opening in the fuel cell. (3) Secure the fuel quantity probe in place with the attaching bolts. Torque the bolts to 25 ± 5 inch-pounds, then safety wire them. NOTE: Install the free end of the bonding jumper under one of the probe mounting bolts. (4) Install the cover over the fuel quantity probe and secure it in place with the attaching bolts. (5) Perform EMPTY TANKS CALIBRATION (PREFERRED PROCEDURE) or FULL TANKS CALIBRATION (ALTERNATE PROCEDURE ONLY) in 28-40-00, 101.

C. Installation - Collector Tank CAUTION: Handle the fuel quantity probe carefully, for damage to the surface of the probe tube will destroy the accuracy of the probe as a measuring device. (1) Position the new gasket and fuel quantity probe in the opening in the collector tank (Ref. Figures 201 and 202). (2) Connect the electrical lead to the fuel quantity probe. (3) Install the attaching washers and bolts and safety wire. (4) Install the large access cover on the bottom of the center section. (5) Perform EMPTY TANKS CALIBRATION (PREFERRED PROCEDURE) or FULL TANKS CALIBRATION (ALTERNATE PROCEDURE ONLY) in 28-40-00, 101.

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2. FUEL PROBE - INTEGRAL (WET WING) FUEL CELL A. Removal (1) Drain the main fuel system. (2) Remove the inboard access panel for the integral fuel cell from the underside of the wing (Ref. 28-00-00 FUEL SYSTEM SCHEMATIC) (Ref. Chapter 6-50-00 WING ACCESS PANELS). (3) Disconnect the interconnect at the inboard end of the fuel tank. (4) Remove the access panel adjacent to the fuel quantity probe from the underside of the wing. (5) Working through the access opening, disconnect the electrical wiring from the terminals on the side of the fuel quantity probe. Disengage the leads from the strain relief clip on the probe. CAUTION: Take precautions to avoid damaging the shielding around the wiring. Damage to the shielding may change system capacitance. (6) Remove the retaining clamp from the bottom and unsafety. Remove the retaining clip from the top of the fuel capacitance probe, then remove the probe from the fuel cell. CAUTION: Handle the fuel quantity probe carefully, for damage to the surface of the probe tubes will destroy the accuracy of the probe as a measuring device.

B. Installation (1) Insert the top of the fuel quantity probe into the retaining clip and safety wire the clip. Install the clamp that anchors the bottom of the fuel quantity probe in place (Ref. Figure 201) (Ref. 28-00-00 FUEL SYSTEM SCHEMATIC). CAUTION: Handle the fuel quantity probe carefully, for damage to the surface of the probe tubes will destroy the accuracy of the probe as a measuring device. (2) Connect the electrical leads to the terminals in accordance with the wiring diagram. Make sure that the cables are secured under the strain relief clip of the probe. CAUTION: Take precautions to avoid damaging the shielding around the wiring. Damage to the shielding may change system capacitance. (3) Connect the interconnect at the inboard end of the fuel tank. (4) Install the access covers and gaskets on the underside of the wing (Ref. Chapter 6-50-00 WING ACCESS PANELS). (5) Perform EMPTY TANKS CALIBRATION (PREFERRED PROCEDURE) or FULL TANKS CALIBRATION (ALTERNATE PROCEDURE ONLY) in 28-40-00, 101.

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3. FUEL PROBE - CENTER SECTION A. Removal CAUTION: Handle the fuel quantity probe carefully, for damage to the surface of the probe tubes will destroy the accuracy of the probe as a measuring device. (1) Remove the small round cover adjacent to the fuselage to gain access to the inboard fuel quantity probe of the center section fuel cell (Ref. Figure 201). (2) Unsafety and remove the five mounting screws securing the jumper wire and capacitance probe in place, then remove the probe from the fuel cell. (3) Remove the large oblong cover aft of the main spar on top of the center section to gain access to the outboard center section fuel quantity probe. (4) Disconnect the three electrical leads of the fuel quantity probe. (5) Unsafety and remove the five mounting screws securing the jumper wire and fuel quantity probe in place, then remove the probe from the fuel cell.

B. Installation (1) Insert the outboard fuel quantity probe into the oblong opening and secure in place with the attaching bolts. Install the free end of the bonding jumper under one of the probe mounting bolts. Torque the bolts to 25 ± 5 inch-pounds, then safety the bolts (Ref. Figure 201). CAUTION: Handle the fuel quantity probe carefully, for damage to the surface of the probe tubes will destroy the accuracy of the probe as a measuring device. (2) Plug the connector of the electrical lead of the fuel quantity probe into the airplane wiring. (3) Install the large oblong access cover on top of the center section. (4) Insert the inboard fuel quantity probe into the small round access opening adjacent to the fuselage. CAUTION: Handle the fuel quantity probe carefully, for damage to the surface of the probe tubes will destroy the accuracy of the probe as a measuring device. (5) Secure in place with the attaching bolts. Install the free end of the bonding jumper under one of the probe mounting bolts. Torque the bolts to 25 ± 5 inch-pounds and safety wire. (6) Plug the connector of the electrical lead of the fuel quantity probe into the airplane wiring. (7) Install the small round access cover, adjacent to the fuselage on top of the center section. (8) Perform EMPTY TANKS CALIBRATION (PREFERRED PROCEDURE) or FULL TANKS CALIBRATION (ALTERNATE PROCEDURE ONLY) in 28-40-00, 101.

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4. FUEL PROBE - COLLECTOR TANK A. Removal CAUTION: Handle the fuel quantity probe carefully, for damage to the surface of the probe tube will destroy the accuracy of the probe as a measuring device. (1) Remove the large access cover on the bottom of the center section (Ref. Figures 201 and 202). (2) Disconnect the electrical connection to the fuel quantity probe. (3) Unsafety and remove the five washers and bolts attaching the probe. (4) Remove the probe from the collector tank.

B. Installation CAUTION: Handle the fuel quantity probe carefully, for damage to the surface of the probe tube will destroy the accuracy of the probe as a measuring device. (1) Position the new gasket and fuel quantity probe in the opening in the collector tank (Ref. Figures 201 and 202). (2) Connect the electrical lead to the fuel quantity probe. (3) Install the attaching washers and bolts and safety wire. (4) Install the large access cover on the bottom of the center section. (5) Perform EMPTY TANKS CALIBRATION (PREFERRED PROCEDURE) or FULL TANKS CALIBRATION (ALTERNATE PROCEDURE ONLY) in 28-40-00, 101.

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Figure 201 Fuel Quantity Probes and Fuel Level Sensors (UA-1 and After; UB-1 and After)

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Figure 202 Fuel Quantity Probe and Fuel Level Sensor (Collector Tank) (UA-1 and After; UB-1 and After)

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FUEL FUEL LEVEL SENSORS MAINTENANCE PRACTICES

28-40-01 200200

1. PROCEDURES A. Removal (1) Drain all fuel from the airplane (Ref. Chapter 12-10-00). (2) Remove ground power from the airplane (Ref. Chapter 24-40-00). (3) Perform BATTERY DISCONNECTION procedure (Ref. Chapter 24-31-00). NOTE: (UA-1 and After; UB-1 and After) The main tank sensor (Detail C), which operates the L FUEL QTY or R FUEL QTY annunciator, may be accessed from the top of the wing. The collector tank sensor (Detail D), which operates the L FUEL FEED or R FUEL FEED annunciator, may be reached from the under side of the wing (Ref. Figures 201 and 202). (UC-1 and After) The main tank sensor (Forward Detail B), which operates the L FUEL QTY or R FUEL QTY annunciator, may be accessed by removing the lower panel (21) of the leading edge cover (Ref. Chapter 06-00-00, Figure 11). The collector tank sensor (Aft Detail B), which operates the L FUEL FEED or R FUEL FEED annunciator, may be reached through the wheel well about even with wing access panel (2) (Ref. Figure 203). (4) Disconnect the electrical plug from the sensor. CAUTION: HANDLE THE SENSOR CAREFULLY; IF THE PRISM IS DAMAGED, THE SENSOR WILL NOT OPERATE. (5) Remove the bolts and washers from the sensor and remove the sensor from the tank. If the tank is to be open for an extended period, cover the sensor port to prevent any contaminants from entering the tank.

B. Installation CAUTION: HANDLE THE SENSOR CAREFULLY; IF THE PRISM IS DAMAGED, THE SENSOR WILL NOT OPERATE. NOTE: Install an AN960-10L washer under the head of each attaching bolt and a NAS1598-3R washer (lubricated with petroleum jelly) under each AN960-10L washer. (1) Position the sensor and new gasket in the fuel tank and install the attaching bolts and washers (Ref. Figure 201 and 202 for UA-1 and After; UB-1 and After) (Ref. Figure 203 for UC-1 and After). (2) Torque the bolts to a value of 20 to 30 inch-pounds. (3) Connect the electrical harness to the sensor. (4) Refuel the airplane and carefully check the installation for leaks and seepage. (5) Perform BATTERY CONNECTION procedure (Ref. Chapter 24-31-00). (6) Install all removed access panels.

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C. Functional Test (1) Perform the FUEL LEVEL SENSOR REMOVAL procedure in this section. (2) Connect the electrical connector to the low fuel quantity sensor. (3) Perform BATTERY CONNECTION procedure (Ref. Chapter 24-31-00). (4) Connect external electrical power to the airplane (Ref. Chapter 24-40-00). (5) Select the BATT switch to the ON position. (6) Select the EXT PWR switch to the EXT PWR position. (7) Position one person in the flight compartment and one at the low fuel quantity sensor. (8) Observe the affected annunciator is illuminated (i.e. L FUEL QTY, R FUEL QTY, L FUEL FEED or R FUEL FEED). NOTE: The container should be a solid container that will block light with a lid that has a suitable hole that will allow just the sensor to fit. (9) Submerge the sensor into a container with at least two inches of fuel. (10) The annunciator should extinguish within a few seconds. (11) With the sensor submerged, press and hold the PRESS TO TEST switch on the annunciator. The annunciator should illuminate within a few seconds. (12) Release the PRESS TO TEST switch. The annunciator should extinguish within a few seconds. (13) While observing the annunciator remove the sensor from the container. The annunciator should illuminate within a few seconds. (14) If any of the above requirements are not met, troubleshoot the system. (15) If all the above checks are satisfactory perform the FUEL LEVEL SENSOR INSTALLATION procedure in this section. (16) Select the EXT PWR switch to the off position. (17) Select the BATT switch to the OFF position. (18) Disconnect external electrical power from the airplane (Ref. Chapter 24-40-00).

2. FUEL QUANTITY INDICATOR A. Removal and Installation Refer to the Chapter 39-10-00 FUEL CONTROL PANEL SECTION AND COMPONENT REMOVAL for the removal and installation of the indicator. Refer to the INDICATOR CALIBRATION in the Troubleshooting section for Fuel Quantity Indicator Calibration checks (Ref. 28-40-00).

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Figure 201 Fuel Quantity Probes and Fuel Level Sensors (UA-1 and After; UB-1 and After)

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Figure 202 Fuel Quantity Probe and Fuel Level Sensor (Collector Tank) (UA-1 and After; UB-1 and After)

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

FUEL WALL STRUCTURE

A 15-423 GASKET

LOW LEVEL SENSOR UNIT

NAS1598-3R WASHER

MS 20073-03-05 BOLT NOTE INSTALL AN960-10L WASHER UNDER BOLT HEAD. APPLY PETROLEUM JELLY TO THE NAS 1598-3R WASHER. TORQUE THE BOLTS TO A TORQUE VALUE OF 20 TO 30 INCH-POUNDS

AN960-10L WASHER

DETAIL

B LEADING EDGE COVER

PANEL #2

B B

DETAIL

A

VIEW LOOKING DOWN AT LEFT WING

UC28B 084424AA.AI

Figure 203 Low Level Sensor Installation (UC-1 and After) (UA-1 and After; UB-1 and After)

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FUEL FUEL QUANTITY INDICATING (UC-1 AND AFTER) DESCRIPTION AND OPERATION

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1. GENERAL A. Capacitance Fuel Quantity Indicating System The fuel indicating system is a capacitance type that is compensated for specific gravity and reads in pounds on a linear scale. Electronic circuits in the system process the signals from the fuel quantity (capacitance) probes in the fuel tanks for an accurate reporting of fuel quantity (Ref. Figure 1). Each side of the airplane has an independent gaging system with five probes strategically located in various areas of the main tank, one in the collector tank and two in the auxiliary tank. Fuel density and dielectric constant vary with respect to temperature, fuel type and fuel batch. The capacitance gaging system is designed to sense and compensate for these variables. The fuel quantity probe is simply a variable capacitor comprised of two concentric tubes. The inner tube is profiled by changing the tube diameter so that the capacitance between the inner and outer tube is proportional to the tank volume. The tubes serve as fixed electrodes and the fuel between the tubes acts as the dielectric for the probe. The capacitance of the fuel quantity probe varies with respect to changes in the dielectric that results from the ratio of fuel to air in the fuel cell. As the fuel level between the inner and outer tubes rises, air with a dielectric constant of one is replaced by fuel with a dielectric constant of approximately two, thus increasing the capacitance of the fuel quantity probe. This variation in the volume of fuel contained in the fuel cell produces a capacitance variation that is a linear function of that volume. This function is converted to a linear current that energizes the fuel quantity indicator. Capacitance fuel probes are designed to produce a capacitance variation that is linear in relation to variation in weight, even though the weight is nonlinear with respect to the fuel level. By varying the capacitance per inch, the nonlinear level signal is changed to a linear function. In addition to its capacitance sensing tubes, each fuel quantity probe contains a small circuit network that produces an output current whose average value is directly proportional to fuel level, while automatically compensating for fuel temperature density variations. The ensuing signal is further processed by a DC amplifier inside the indicator, which contains a potentiometer for adjusting the “FULL” and “EMPTY” settings. The DC amplifier controls response time and drives a D'Arsonval type meter on Gull Airborne indicators or a servo-type meter on Ragen Data Systems indicators. Correct operation of the capacitance fuel quantity indicating system depends on proper system connections. These connections should be thoroughly inspected before performing other maintenance procedures. For detailed system checks and calibration, refer to FUEL QUANTITY INDICATING TROUBLESHOOTING in this Chapter.

B. Low Fuel Quantity Warning System Sensors in the inboard wall of the main tank provide low fuel quantity notification when the fuel quantity drops below the sensing element prism. When the prism is not submerged, the probe triggers current to the appropriate annunciator light. Sensors in the collector tank area of the main tank provide the signal for the low fuel feed annunciator light in the same manner. The system may be functionally tested by pressing the test switch mounted adjacent to the annunciator panel (Ref. Figure 2). The low fuel quantity sensors and the low fuel feed sensors are connected electrically in parallel at the 28 vdc in, test signal, and ground leads. The 28 vdc out leads are connected individually to their respective annunciator circuits. Five-ampere circuit breakers provide operating current for the RH and LH systems.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL The low fuel quantity sensors operate on a principal that allows a light produced within the sensor by a reference light source to pass through a prism on the tip of the sensor when the probe is submerged in fuel. When the prism is not submerged, the light is reflected back, off the prism bevel and back into the probe. The reflecting light is then sensed by a light sensitive transistor which triggers the current for the appropriate annunciator light. An integral self-test circuit provides for functionally testing the sensor and its associated circuitry by energizing a test light source in the probe. The test light is positioned to shine into the light-sensitive transistor even when the probe is submersed in fuel. When the warning annunciator test switch is pressed, illumination of the annunciators is not immediate because of a time delay built into the circuitry. The time delay is designed to allow up to a 4 second delay before the annunciator light is activated. The time delay feature prevents false cycling of the annunciator light when the fuel is moving around in the wing. Circuit failures within the sensor will always be in the ON mode; therefore, a false annunciator will result if the circuit fails. A failure of the reference light source will cause the light sensitive transistor to be biased, allowing it to conduct and resulting in a false annunciation.

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Figure 1 Fuel Quantity Indicating Schematic

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Figure 2 Low Fuel Quantity Warning Block Diagram

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FUEL FUEL QUANTITY INDICATING (UC-1 AND AFTER) TROUBLESHOOTING

100100

1. PROCEDURES When the low fuel quantity indicating system fails to test, a complete check of the system wiring is indicated. Once the technician is satisfied as to the integrity of the system wiring, the fault may be assumed to be in the sensor, and the inoperative sensor may be replaced according to the procedure under FUEL QUANTITY INDICATING - MAINTENANCE PRACTICES. NOTE: No adjustments or repairs can be made to the sensor; therefore, an improperly operating sensor must be replaced.

A. System Insulation Check NOTE: Connect a regulated 28 vdc external power supply to the airplane when troubleshooting or calibrating. This test procedure may be performed with full, partially full or empty tanks. Display readings are presented in nano Siemens (nS), equivalent megohms are derived by dividing 1000 by the LCD display reading: 1000/ 50 nS = 20 Megohms. These checks are identical for left and right fuel systems. (1) Locate the Fuel Quantity Test Set (14, Table 7, Chapter 91-00-00) conveniently for the procedures that follow. (2) Ensure that the test set ON/OFF switch is in the OFF position. CAUTION: Power to the airplane must be off when lowering or raising the fuel control panel to preclude the possibility of shorting any wiring connections to ground. (3) Ensure that the battery switch is OFF. (4) Open the fuel quantity system circuit breakers. (5) Set the fuel select switch to MAIN. (6) Gain access to the fuel quantity indicator. (7) Disconnect the airplane's harness from the fuel quantity indicator. (8) Connect the test set to the airplane’s harness. NOTE: Do not connect the test set to the fuel quantity indicator. (9) Rotate the TEST FUNCTION selector to IND AMP. (10) Place the INSULATION/SYSTEM switch to IN nS. (11) Rotate the INS TEST POINTS selector to LO-Z GND. (12) Place the shorting switch to the NO position. (13) Place the ON/OFF switch to ON and allow sufficient time for the LCD to stabilize. (14) The LCD reading should be less than 50 nS.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (15) Rotate the INS TEST POINTS selector through the remaining positions. Allow sufficient time for the LCD to stabilize in each position. The LCD reading in each position should be less than 50 nS, except in the SIG/RTN position which will show an overrange indication (1---). (16) Rotate the INS TEST POINTS selector to RTN/SIG. (17) Close the fuel system circuit breakers that were opened in Step (4). (18) Turn the battery master switch to ON. The maximum time delay from power OFF to power ON is 10 seconds. (19) Rotate the INS TEST POINTS selector through the remaining positions. Allow sufficient time for the LCD to stabilize in each position. The LCD reading in each position should be less than 50 nS, with the same exceptions as previously noted. Due to continuity through the diodes inside the probe, when the test set detects a short from RTN to GND, that short will also be seen as a short from SIG to GND. (20) Set the fuel select switch to AUX. Repeat Steps (11) through (19). (21) When any of the above checks are out of tolerance, isolate the airplane's wire harness from the probe wire harnesses. The main tank wire harness connectors are behind access panel 21 and the auxiliary tank wire harness connectors are behind panel 25. Refer to Chapter 6-50-00 WING ACCESS PANELS illustration in the Dimensions and Areas section. (22) Repeat Steps (9) through (19). Acceptable readings are 5 nS or less. If the insulation readings are acceptable, the fault lies in the main tank wire harness to the probes. If not, the fault is in the airplane's wire harness between the indicator in the cockpit out to the wing. (23) Perform PROBE INSULATION TEST when any faults have been isolated to the probe wire harnesses. NOTE: The results of the preceding checks must be within the stated limits. If any results are slightly outside of these limits, repeat the entire procedure to ensure that failure of the checks was not due to human error. (24) Turn the test set OFF and disconnect the test set. (25) Connect the airplane's wire harnesses to the original configuration.

B. Probe Insulation Test The probe insulation check may be performed with the fuel tanks full, partially full or empty. (1) Ensure that the battery master switch is OFF. (2) Open all fuel system circuit breakers. (3) Set the fuel select switch to MAIN. (4) Locate the Fuel Quantity Test Set (14, Table 7, Chapter 91-00-00) conveniently for working in the area of the access panels for the main tank wire harness connectors.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (5) Disconnect the probe wire harnesses from the airplane's wire harness. The main tank wire harness connectors are behind access panel 21 and the auxiliary wire harness connectors are behind access panel 25. Refer to Chapter 6-50-00 WING ACCESS PANELS illustration in the Dimensions and Areas section. (6) Ensure that the test box ON/OFF switch is in the OFF position. (7) Connect the Probe Selector Unit (15, Table 7, Chapter 91-00-00) (Ref. Figure 101) to the bulkhead connector of the main tank wire harness and to the bulkhead connector of the auxiliary tank wire harness. (8) Connect the test set to the probe selector unit with the ACFT connector. NOTE: Do not connect the test set IND connector to the airplane's wire harness. (9) Rotate the TEST FUNCTION switch to IND AMP. (10) Place the INSULATION/SYSTEM switch to IN nS. (11) Rotate the INS TEST POINTS selector to LO-Z GND. (12) Set the probe selector ALL/EACH toggle switch to ALL. (13) Set the probe selector MAIN/AUX toggle switch to MAIN. (14) Place the test set ON/OFF switch in ON and allow sufficient time for the LCD to stabilize. (15) The LCD reading should be less than 50 nS. (16) Rotate the INS TEST POINTS selector to its remaining positions, allowing sufficient time for the LCD to stabilize in each position: the SIG/RTN position should show an overrange indication (1---); all others should indicate less than 50 nS. (17) Set the probe selector ALL/EACH toggle switch to EACH. (18) Using the switch on the selector unit, select each individual probe and repeat Steps (10) through (15): the reading for each probe should be less than 50 nS, and the total readings for all the probes must not exceed 50 nS except in the SIG/RTN position, which should show an overrange indication. (19) Set the fuel select switch to AUX. (20) Set the probe selector MAIN/AUX toggle switch to AUX. (21) Set the probe selector BOTH/EACH toggle switch to BOTH. (22) Repeat Steps (11) and (14) through (16) to read the insulation values of the auxiliary tank probe harness. (23) Set the probe selector BOTH/EACH toggle switch to EACH and repeat Step (18). (24) Should any probe be out of tolerance, that probe or its associated wiring to the bulkhead connector must be replaced.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (25) Once the fault has been isolated and corrected, the applicable portion of the preceding checks should be repeated to ensure fault correction. NOTE: If any component has been changed, perform the SYSTEM CAPACITANCE CHECKS and INDICATOR CALIBRATION procedures in this section. (26) Place the test set ON/OFF switch in the OFF position, and restore the airplane to its original condition.

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Figure 101 Probe Selector Unit

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C. System Capacitance Check The following checks must be performed with the airplane's fuel tanks either completely empty or completely full. (1) Locate the Fuel Quantity Test Set (14, Table 7, Chapter 91-00-00) convenient to the indicator for these checks. (2) Ensure that the ON/OFF switch on the test set is in the OFF position. (3) Ensure that the battery master switch is off. (4) Open the fuel quantity system circuit breakers. (5) Place the fuel select switch to MAIN. (6) Gain access to fuel quantity indicator. (7) Disconnect the fuel quantity indicator harness from the rear of the indicator. (8) Connect the test set with the ACFT connector to the indicator harness. NOTE: Do not connect the IND connector to the fuel quantity indicator. (9) Rotate the TEST FUNCTION selector on the test set to PROBES. (10) Place the INSULATION/SYSTEM switch to SYSTEM. (11) Place the MAIN/TOT-AUX/NAC switch to MAIN/TOT. (12) Place the shorting switch to the YES position. (13) Place 200 pF/1000 pF switch to the 200 pF position if tanks are empty or to the 1000 pF position if the tanks are full. (14) Place the test set ON/OFF switch to ON. (15) Press and hold the PRESS TO READ CAP (pF) push button: the reading should be within the range shown in Table 101 or Table 102. Record these empty and full capacitance values in the airframe log for future reference. (16) Place the fuel select switch to AUX. (17) Press and hold the PRESS TO READ CAP (pF) push button: the reading should be within the range shown in Table 101 or Table 102. Record these empty and full capacitance values in the airframe log for future reference. (18) If the system capacitance is out of tolerance, perform the PROBE CAPACITANCE CHECK. (19) Place the test set ON/OFF switch in the OFF position and disconnect the test set from the indicator harness.

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D. Probe Capacitance Check (1) Ensure that the battery master switch is OFF. (2) Open all fuel system circuit breakers. (3) Set the fuel select switch to MAIN. (4) Locate the Fuel Quantity Test Set (14, Table 7, Chapter 91-00-00) and Probe Selector Unit (15, Table 7, Chapter 91-00-00) conveniently for working in the area of the access panels for the main tank wire harness connectors and the auxiliary tank wire harness connectors. (5) Disconnect the main tank wire harness from the airplane's wire harness. The main tank wire harness connectors are behind access panel 21, and the auxiliary tank wire harness connectors are behind panel 25. Refer to Chapter 6-50-00 WING ACCESS PANELS illustration in the Dimensions and Areas section. (6) Ensure that the test set ON/OFF switch is in the OFF position. (7) Connect the probe selector unit to the bulkhead connector of the main tank wire harness and to the bulkhead connector of the auxiliary tank wire harness. (8) Connect the test set to the probe selector unit with the ACFT connector. NOTE: Do not connect the test set IND connector to the airplane's wire harness. (9) Rotate TEST FUNCTION selector to PROBES. (10) Place INSULATION/SYSTEM switch to SYSTEM. (11) Place MAIN/TOT-AUX/NAC to MAIN/TOT. (12) Place 200 pF/1000 pF switch to 200 pF if the tanks are empty or to 1000 pF if the tanks are full. (13) Set probe selector ALL/EACH switch to ALL and set the probe selector MAIN/AUX switch to MAIN. (14) Place test set ON/OFF switch in the ON position. (15) Press and hold PRESS TO READ pF push button: the reading should be within the ranges shown in Tables 101 and 102. Record these empty and full capacitance values in the airframe log for future reference. (16) Release PRESS TO READ pF push button. (17) Place 200 pF/1000 pF switch to 200 pF. (18) Set probe selector ALL/EACH switch to EACH. (19) Rotate selector switch to each probe position and measure the capacitance of the individual probes: the capacitance readings should be within the values listed in Table 103; otherwise, replace the out of tolerance probe or its associated wiring, then repeat the appropriate checks to confirm that the fault has been corrected. (20) Set fuel select switch to AUX.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (21) Set probe selector BOTH/EACH switch to BOTH and set the probe selector MAIN/AUX switch to AUX. (22) Repeat Steps (9) through (12) and (14) through (17). (23) Set probe selector BOTH/EACH switch to EACH and repeat Step (19). NOTE: If any component or components have been replaced, perform the SYSTEM CAPACITANCE CHECK and INDICATOR CALIBRATION procedures in this section.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Table 101 EMPTY SYSTEM CAPACITANCE VALUES Main Tank

Auxiliary Tank

Nom

Min

Max

Nom

Min

Max

159.3

156.3

162.2

64.1

63.1

65.1

Table 102 FULL SYSTEM CAPACITANCE VALUES Main Tank

Auxiliary Tank

Nom

Min

Max

Nom

Min

Max

295.6

292.6

298.6

115.2

114.2

116.2

Table 103 INDIVIDUAL PROBE CAPACITANCE VALUES Empty Tanks

Full Tanks

Probe Location

Probe P/N

Min

Max

Min

Max

WS 41.85 (Aux)

118-389004-41

36.6

37.6

66.0

67.0

WS 79.65 (Aux)

118-389004-43

26.5

27.5

48.2

49.2

WS 114.25

118-389004-45

6.7

7.7

13.8

14.8

WS 135.86

118-389004-13

48.2

49.3

90.5

91.5

WS 171.29

118-389004-47

23.6

24.6

42.2

43.2

WS 194.85

118-389004-21

23.0

24.0

43.0

44.0

WS 223.49

118-389004-25

28.6

29.6

52.6

53.6

WS 276.01

118-389004-49

26.4

27.4

50.5

51.5

Table 104 CAPACITANCE SIMULATION VALUES - MAIN TANK 0 lbs

300 lbs

600 lbs

900 lbs

1200 lbs

1600 lbs

159.4

184.8

210.3

235.7

261.1

295.0

Table 105 CAPACITANCE SIMULATION VALUES - AUXILIARY TANK EMPTY

----

----

---

----

FULL

64.1

----

----

---

----

115.2

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

E. Setting Capacitance Simulation When performing the indicator linearity check or calibrating the indicator, the Fuel Quantity Test Set (14, Table 7, Chapter 91-00-00) uses the airplane's 28 volts and it's circuitry to simulate a capacitance for the indicator to read. This value is shown on the indicator face as a fuel quantity. During a bench test the procedures for simulating capacitance is the same, the only difference being that an auxiliary 28 vdc power supply is used in lieu of the airplane's power. Use these procedures and Tables 104 and 105 as required to set capacitance values for checking and calibrating the indicator. NOTE: Connect the test set ACFT connector to the airplane’s wire harness and the IND connector to the indicator. (1) Set the TEST FUNCTION selector switch to CAP SIM CAL (NORM SYS). (2) Set the INSULATION/SYSTEM switch to SYSTEM. (3) Set the MAIN/TOT-AUX/NAC switch to MAIN/TOT. (4) Set the shorting switch to NO. (5) Set the pF switch to 200 for values from 0 to 200 pF or in the 1000 pF position for values over 200. (6) Turn the test set ON. (7) Set the CAP SIM (pF) 100's and 10's thumb wheels to the desired capacitance value to be simulated. (8) Press and hold the PRESS TO READ CAP (pF) button to read the capacitance in the LCD. (9) Use the trimmer control to fine adjust the capacitance setting. (10) Set the TEST FUNCTION switch to IND AMP. (11) Turn on the battery switch and apply auxiliary power (28 vdc) to airplane. (12) The fuel quantity indicator will read the capacitance simulated by the test set and display a fuel quantity in pounds.

F. Indicator Linearity Check The following linearity check may be performed on the fuel quantity indicator with the fuel tanks full, partially full or empty. (1) Ensure that the battery master switch is OFF. (2) Locate the Fuel Quantity Test Set (14, Table 7, Chapter 91-00-00) convenient to the fuel control panel on the pilot's sidewall. (3) Open the appropriate fuel quantity system circuit breakers. (4) Gain access to the fuel quantity indicator. (5) Connect the test set ACFT connector to the indicator wire harness and the IND connector to the indicator.

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28-41-00

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (6) Set the fuel select switch to MAIN. (7) Use the SETTING CAPACITANCE SIMULATION procedures in this section, to set the test set to the values shown in Table 104. (8) Set the indicator needle deflection, using the EMP 1/E1 or E adjustment, for each required quantity shown in Table 104. (9) Set the fuel select switch to AUX and check the empty and full readings of the indicator in accordance with the values in Table 105. (10) Set the indicator needle deflection by using the EMP 1/E1 or E adjustment to the required quantity. (11) If the indicator will not adjust to the required quantity reading for the capacitance simulated, it must be replaced. (12) Turn the battery switch OFF. (13) Turn the test set OFF. (14) Disconnect the test set. (15) Connect the airplane's wire harness to the indicator and return the airplane to the original configuration.

G. Indicator Calibration The fuel quantity indicator may be calibrated with the fuel tanks either completely empty (undrainable fuel only) or completely full; however, the preferred procedure is to calibrate the indicator with the tanks completely empty. While calibrating the indicator with empty tanks is more reliable than the alternate procedure, the accuracy of the alternate procedure (full tanks) can be appreciably enhanced if the empty-tanks value is known. It is recommended that when the empty capacitance value is determined, it should be recorded in the airframe log for future use in calibrating. The preferred procedure (undrainable fuel only) makes use of the actual empty-tank value in adjusting the zero indication. The alternate procedure (full tanks) applies a nominal empty-tank value to the system by the test set for the zero indication. Both procedures will be outlined in the following text; however, it must be remembered that the alternate procedure is a stop-gap procedure only and the system should be recalibrated using the preferred procedure at the earliest convenience of the operator.

H. Empty Tanks Calibration (Preferred Procedure) (1) Defuel the airplane and drain sumps. (2) Locate the Fuel Quantity Test Set (14, Table 7, Chapter 91-00-00) conveniently for working in the cockpit. (3) Ensure that the ON/OFF switch on the test set is OFF. (4) Ensure that the battery switch is OFF and the fuel quantity circuit breakers are open. (5) Gain access to the fuel quantity indicator.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (6) Disconnect the wiring connector from the indicator. (7) Connect the test set ACFT connector to the indicator wire harness and connect the test set IND connector to the indicator. (8) Rotate the TEST FUNCTION selector to CAP SIM CAL (NORM SYS). (9) Place the INSULATION/SYSTEM switch to SYSTEM. (10) Place the MAIN/TOT-AUX/NAC switch to MAIN/TOT. (11) Set the fuel select switch to MAIN. (12) Place the shorting switch to the NO position. (13) Place the ON/OFF switch to ON. (14) Close the appropriate fuel quantity circuit breakers. (15) Turn the battery switch ON. NOTE: There are two different indicators which may be installed in the 1900 Series Airplanes: one type of indicator has a power-off dot below the “0” graduation. (16) The fuel quantity indicator needle should rest between the “0” indication and the power-off dot on indicators which have the power-off dot, or the needle should rest approximately one needle width below the “0” indication on indicators without the power-off dot. Should the indication be inconsistent with the preceding guidelines, adjust the EMP 1/E1 or E/EMP adjustment on the back of the indicator. (17) Set the fuel select switch to AUX. The indicator should read “0”. Make adjustments as required using the on E2/EMP 2 adjustment on the back of the indicator. (18) Set the fuel select switch back to MAIN. (19) Use the procedures in SETTING CAPACITANCE SIMULATION to set a capacitance value on the test set as follows: ADD CAP = 135.9 pF. (20) Rotate the TEST FUNCTION selector to ADD CAP. (21) The value set on the test set is added to the current empty tank reading which will simulate a full tank. Adjust the F/FULL adjustment on the indicator to read the maximum fuel quantities shown in Table 104. NOTE: Check empty and full adjustments made to the indicator to ensure accurate calibration. (22) Turn the battery switch OFF. (23) Turn the ON/OFF switch on the test set OFF and disconnect the test set from the airplane's system. (24) Restore the airplane to the original configuration.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

I. Full Tanks Calibration (Alternate Procedure Only) NOTE: With tanks FULL use Steps (1) through (23), then (26) to end. When tanks are PARTIALLY FULL use Steps (1) through (22), then Steps (24) to end. (1) Locate the Fuel Quantity Test Set (14, Table 7, Chapter 91-00-00) conveniently for working in the cockpit. (2) Ensure that the ON/OFF switch on the test set is OFF. (3) Ensure that the battery switch is OFF and the fuel quantity circuit breakers are open. (4) Gain access to the fuel quantity indicator. (5) Disconnect the wiring connector from the indicator. (6) Connect the test set ACFT connector to the indicator wire harness and connect the test set IND connector to the indicator. (7) Rotate the TEST FUNCTION selector to CAP SIM CAL (NORM SYS). (8) Place the INSULATION/SYSTEM switch to SYSTEM. (9) Place the MAIN/TOT-AUX/NAC switch to MAIN/TOT. (10) Set the fuel select switch to MAIN. (11) Place the shorting switch to the NO position. (12) Place the ON/OFF switch to ON. (13) Close the appropriate fuel quantity circuit breakers. (14) Turn the battery switch ON. (15) Set the test set to a nominal empty capacitance value for the main tank as shown in Table 101. Check the airframe log for an empty tank value and, if available, use that value in place of the values given in Table 101. (16) Rotate the TEST FUNCTION switch to IND AMP. (17) Turn the battery switch ON. (18) The indicator needle should indicate between the “0” position and the power-off dot for indicators with the power-off dot or approximately one needle width below the “0” indication on indicators without the power-off dot. Should the indicator reading be inconsistent with the preceding guide lines, adjust the EMP 1/E1 or E/EMP adjustment on the back of the indicator. (19) Set the fuel select switch to AUX and set the test set to a nominal empty capacitance value for the auxiliary tank (Ref. Table 101). Check the airframe log for an empty tank value and, if available, use that value in place of the values given in Table 101. The indicator should read “0”. Make adjustments as required in accordance with the previous Step. (20) Set the fuel select switch to MAIN.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (21) Rotate the TEST FUNCTION switch to CAP SIM CAL (NORMAL SYS). (22) The indicator should read the maximum value for a full tank (1637 pounds). Make adjustments with the F/FULL adjustment as required. (23) Set the test set to a nominal full capacitance value for the main tank as shown in Table 102. Check the airframe log for an empty tank value and, if available, use that value in place of the values given in Table 101. (24) Rotate the TEST FUNCTION switch to IND AMP. (25) The indicator should read 1600 lbs. NOTE: Check the empty and full adjustments made to the indicator to ensure accurate calibration. (26) Place test set ON/OFF switch in the OFF position. (27) Turn the battery master switch OFF. (28) Open the appropriate fuel quantity circuit breakers. (29) Disconnect the test set from the airplane's system. (30) Restore the airplane to the original configuration.

J. Fuel Quantity Probe Bench Check Should a fuel quantity probe become suspect as a result of the preceding checks, the following bench check procedure must be performed on the probe prior to returning the probe to Hawker Beechcraft Corporation. (1) Ensure that the ON/OFF switch on the Fuel Quantity Test Set (14, Table 7, Chapter 91-00-00) is in the OFF position. (2) Rotate the TEST FUNCTION selector to PROBES. (3) Place the INSULATION/SYSTEM switch to SYSTEM. (4) Place the 200 (pF)/1000 (pF) switch to 200 (pF). (5) Connect the Probe Adapter (16, Table 7, Chapter 91-00-00) to the PROBES socket on the test set. (6) Connect the color-coded leads from the adapter to the matching color and threads of the probe terminals as follows: L-YELLOW (10-32), S-RED (8-32), R-GREEN (6-32). Connect the ground clip to the flange of the probe and place the shorting switch in the YES position. NOTE: The grounding clip is not used when testing an integral tank probe that does not have a flange. The shorting switch is still placed in the YES position. (7) Place the ON/OFF switch on the test set in the ON position. (8) Press the PRESS TO READ CAP (pF) push button to read the probe capacitance on the LCD allowing sufficient time for the LCD to stabilize.

Page 114 Nov 1/09

28-41-00

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (9) The probe capacitance value must be within the tolerance as specified, by part number in Table 103. (10) If the probe capacitance is not within tolerance, the probe should be replaced. (11) Rotate the TEST FUNCTION selector to IND AMP. (12) Place the INSULATION/SYSTEM switch in the IN nS position. (13) Place the shorting switch to the NO position. (14) Rotate the INS TEST POINT selector to LO-Z GND. (15) Allow sufficient time for the LCD to stabilize: the LCD reading must be less than 50 nS. (16) Rotate the INS TEST POINT selector to each of the remaining position allowing sufficient time for the LCD to stabilize in each position: all readings must be less than 50 nS except in the SIG/RTN position which will display and overrange indication (1---). (17) Should any of the preceding indications be inconsistent with the expected results, the probe should be replaced. (18) Place the ON/OFF switch on the test set in the OFF position and disconnect the test set adapter from the probe.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

FUEL FUEL QUANTITY INDICATING (UC-1 AND AFTER) MAINTENANCE PRACTICES

28-41-00 200200

1. FUEL PROBE A. Removal (1) Drain the fuel from the airplane. CAUTION: DO NOT remove or attempt to remove any of the lower wing access panels until it is certain that all fuel has been drained from the wings. (2) Disconnect all electrical power from the airplane. (3) Remove the appropriate access cover that allows access to the probe that is to be removed (Ref. Figure 201). (4) Unsnap and loosen the clamps securing the probe to the mounting brackets. (5) Disconnect the electrical harness attached to the probe terminal block. (6) Remove the probe from the tank.

B. Installation CAUTION: Handle the fuel quantity probe carefully. Damage to the surface of the probe may destroy the accuracy. (1) Connect the wiring to the fuel probe terminal block. (2) Position the probe in the mounting bracket clamps. (3) Assure that the clamps are in position around the probe collars and lock the clamp levers. (4) Install the access cover with a new packing. (5) Perform EMPTY TANKS CALIBRATION (PREFERRED PROCEDURE) or FULL TANKS CALIBRATION (ALTERNATE PROCEDURE ONLY) in 28-41-00, 101.

28-41-00

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 201 Fuel Quantity Probe Installation

Page 202 Nov 1/13

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

28-41-01 200200

FUEL FUEL LEVEL AND LOW FUEL QUANTITY SENSORS (UC-1 AND AFTER) MAINTENANCE PRACTICES 1. LOW FUEL QUANTITY SENSOR A. Removal (1) Drain all fuel from the airplane (Ref. Chapter 12-10-00). (2) Remove ground power from the airplane (Ref. Chapter 24-40-00). (3) Perform BATTERY DISCONNECTION procedure (Ref. Chapter 24-31-00). NOTE: The main tank sensor (Forward Detail B), which operates the L FUEL QTY or R FUEL QTY annunciator, may be accessed by removing the lower panel (21) of the leading edge cover (Ref. Chapter 6-50-00, Figure 11). The collector tank sensor (Aft Detail B), which operates the L FUEL FEED or R FUEL FEED annunciator, may be reached through the wheel well about even with wing access panel (2) (Ref. Figure 201). (4) Disconnect the electrical plug from the sensor. CAUTION: Handle the sensor carefully; if the prism is damaged, the sensor will not operate. (5) Remove the bolts and washers from the sensor and remove the sensor from the tank. If the tank is to be open for an extended period, cover the sensor port to prevent any contaminants from entering the tank.

B. Installation CAUTION: Handle the sensor carefully; if the prism is damaged, the sensor will not operate. NOTE: Install an AN960-10L washer under the head of each attaching bolt and a NAS1598-3R washer (lubricated with petroleum jelly) under each AN960-10L washer. (1) Position the sensor and new gasket in the fuel tank and install the attaching bolts and washers (Ref. Figure 201). (2) Torque the bolts to a value of 20 to 30 inch-pounds. (3) Connect the electrical harness to the sensor. (4) Refuel the airplane and carefully check the installation for leaks and seepage. (5) Perform BATTERY CONNECTION procedure (Ref. Chapter 24-31-00). (6) Install all removed access panels.

2. FUEL LEVEL SENSOR A. Functional Test (1) Perform the FUEL LEVEL SENSOR REMOVAL procedure in this section.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (2) Connect the electrical connector to the low fuel quantity sensor. (3) Perform BATTERY CONNECTION procedure (Ref. Chapter 24-31-00). (4) Connect external electrical power to the airplane (Ref. Chapter 24-40-00). (5) Select the BATT switch to the ON position. (6) Select the EXT PWR switch to the EXT PWR position. (7) Position one person in the flight compartment and one at the low fuel quantity sensor. (8) Observe the affected annunciator is illuminated (i.e. L FUEL QTY, R FUEL QTY, L FUEL FEED or R FUEL FEED). NOTE: The container should be a solid container that will block light with a lid that has a suitable hole that will allow just the sensor to fit. (9) Submerge the sensor into a container with at least two inches of fuel. (10) The annunciator should extinguish within a few seconds. (11) With the sensor submerged, press and hold the PRESS TO TEST switch on the annunciator. The annunciator should illuminate within a few seconds. (12) Release the PRESS TO TEST switch. The annunciator should extinguish within a few seconds. (13) While observing the annunciator remove the sensor from the container. The annunciator should illuminate within a few seconds. (14) If any of the above requirements are not met, troubleshoot the system. (15) If all the above checks are satisfactory perform the FUEL LEVEL SENSOR INSTALLATION procedure in this section. (16) Select the EXT PWR switch to the off position. (17) Select the BATT switch to the OFF position. (18) Disconnect external electrical power from the airplane (Ref. Chapter 24-40-00).

3. FUEL QUANTITY INDICATOR A. Removal and Installation Refer to the Chapter 39-10-00 FUEL CONTROL PANEL SECTION AND COMPONENT REMOVAL for the removal and installation of the indicator. Refer to the INDICATOR CALIBRATION in the Troubleshooting section for Fuel Quantity Indicator Calibration checks.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

FUEL WALL STRUCTURE

A 15-423 GASKET

LOW LEVEL SENSOR UNIT

NAS1598-3R WASHER

MS 20073-03-05 BOLT NOTE INSTALL AN960-10L WASHER UNDER BOLT HEAD. APPLY PETROLEUM JELLY TO THE NAS 1598-3R WASHER. TORQUE THE BOLTS TO A TORQUE VALUE OF 20 TO 30 INCH-POUNDS

AN960-10L WASHER

DETAIL

B

LEADING EDGE COVER PANEL #2

B B

DETAIL

A

UC28B 084424AA.AI

VIEW LOOKING DOWN AT LEFT WING

Figure 201 Low Fuel Quantity Sensor Installation

28-41-01

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Model 1900/1900C Airliner (UA-1 and After) (UB-1 and After) (UC-1 and After)

Maintenance Manual

Volume 2 Chapter 30 thru Chapter 91

Copyright © 2017 Beechcraft Corporation. All rights reserved. Hawker and Beechcraft are trademarks of Beechcraft Corporation. P/N 114-590021-7 Issued: November 12, 1982

P/N 114-590021-7C12 Revised: January 1, 2017

Published by Beechcraft Corporation P.O. Box 85 Wichita, Kansas 67201-0085 USA

The export of these commodities, technology or software are subject to the US Export Administration Regulations. Diversion contrary to US law is prohibited. For guidance on export control requirements, contact the Commerce Department’s Bureau of Export Administration at 202-482-4811 or visit the US Department of Commerce website.

REV4, 4/20/07

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

CHAPTER 30 - ICE AND RAIN PROTECTION TABLE OF CONTENTS SUBJECT

PAGE

GENERAL INFORMATION 30-00-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1

BRAKE DEICE SYSTEM 30-01-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1

AIRFOIL 30-10-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Surface Deicer System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Troubleshooting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 101 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 101 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Surface Deicer Boot . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Cleaning . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Surface Deicer Fast Boots Removal and Installation (Pressure Sensitive Adhesion) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202 Shortened 1300L Dry Time For Standard Pneumatic Deicers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 204 Test Control Strips . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 206 Adhesion Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 206 Boot Acceptability . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 206 Surface Deicer Boot Age Master and Icex Application . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 206 Surface Deicer Boot Repairs . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 207 Resurfacing Deicer Boots . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 208 Scuff Damage . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 208 Tube Area Damage . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 209 Loose Surface Ply In Non-Inflatable Area . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 209 Loose Surface Ply In Tube Area . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 210 Fabric Back Ply Damage During Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 210 Surface Deicer Operational Checks . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 210 With Engines Operating . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 210 Without Engines Operating . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 210 Bleed Air Pressure Regulator Relief Valve Adjustment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 211 Vacuum Regulator Valve Adjustment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 211

30-CONTENTS

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

CHAPTER 30 - ICE AND RAIN PROTECTION TABLE OF CONTENTS (CONTINUED) SUBJECT

PAGE

AIR INTAKES 30-20-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Engine Air Inlet Lip Anti-Ice . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Inertial Separation System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Air Intake Anti-ice Lip (UA-1 and After, UB-1 thru UB-60 Without Kit No. 114-9014-1 S Installed) . . . . . . . . . . . . . . . . . . . . . . 201 Air Intake Anti-ice Lip (UB-61 and After, UC-1 and After and Airplanes With Kit No. 114-9014-1 S Installed) . . . . . . . . . . . . . . 201 Engine Air Inlet Anti-Ice Lip . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Crack Limits . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Inertial Anti-icing System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202 Rigging the Engine Inertial Anti-Icing Systems . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202

WINDOWS AND WINDSHIELDS 30-40-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Windshield Anti-Ice . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Windshield Wiper . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Windshield Heating Elements . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Resistance Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Windshield Wiper Motor . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202 Windshield Wiper Converter . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 203 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 203 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 203 Windshield Wiper . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 204 Adjustment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 204

PROPELLER 30-60-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Propeller Electric Deicer System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Troubleshooting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 101 Propeller Deicer Boot . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 101 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Propeller Deice Ammeter . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Calibration . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Propeller Deicer Boot . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

ICE AND RAIN PROTECTION GENERAL INFORMATION DESCRIPTION AND OPERATION

30-00-00 00

1. GENERAL The airplane is equipped with a variety of ice and rain protection systems that can be utilized during operation under inclement weather conditions. Electrical heating elements embedded in the windshield provide adequate protection against the formation of ice while air from the cabin heating systems prevents fogging to ensure visibility during operation under icing conditions. Heavy duty windshield wipers for both the pilot and copilot provide further visibility during rainy flight conditions. Pneumatic deicer boots on the wings, the stabilons, and the horizontal stabilizers prevent the formation of ice during flight. Regulated bleed air pressure and vacuum are cycled to the pneumatic boots for the inflation-deflation cycle. The selector switch that controls the system permits automatic single cycle operation or manual operation. Ice protection for the engine is provided by an inertial separation system utilizing a moveable vane and bypass door actuated through an electric actuator for each engine. When icing conditions are encountered, the moveable vane and ice bypass door are lowered into the inlet airstream to induce an abrupt turn in the airflow before entering the engine plenum. The heavy ice-laden air is then discharged overboard through a bypass duct in the lower cowling at the aft end of the air duct while directing the lighter ice-free air into the engine plenum. The lip around each engine air inlet is heated by hot exhaust gases to prevent the formation of ice. The system is functioning anytime the engine is running and does not require any crew action. The propellers are protected against icing by electrothermal boots on each blade that automatically cycle upon actuation. Heat from the boots reduces the grip of the ice, which is then removed by the centrifugal effect of the propeller rotation and the blast of the air stream. Engine bleed air is routed by a hose through a solenoid operated shutoff valve to a distributor manifold that directs the hot air to the brakes for deicing. A heating element in each pitot mast and alternate static button prevents them from becoming clogged with ice. The heating elements are connected into the airplane electrical system. Two circuit breaker switches on the left subpanel control the heating elements.

30-00-00

Page 1 Nov 1/09

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

ICE AND RAIN PROTECTION BRAKE DEICE SYSTEM DESCRIPTION AND OPERATION

30-01-00 00

1. GENERAL The heated air for brake deicing is supplied by bleeding air from the compressor of each engine. This engine bleed air is routed by a line on the left side of each nacelle to a solenoid operated shutoff valve in the center of the firewall. From the shutoff valve, bleed air is routed through a hose secured to the aft side of the landing gear strut and down to a distributor manifold attached to the two bottom bolts in the piston and axle assembly. The bleed air is directed to the brake for each wheel through orifices around the circumference of each ring of the distributor manifold. The Brake Deice System is controlled by an ON-OFF toggle switch mounted on the subpanel to the right of the pilot's control column. When this switch is placed in the ON position, power from the airplane electrical system is supplied through a 5-ampere circuit breaker in the copilot's sidepanel to a control module located under the center aisle floorboard at FS 213.25, in line with the first cabin window. This module supplies current to open the solenoid shutoff valves in each wheel well, allowing the hot bleed air to enter the distributor manifold for diffusion through the orifices to deice the brakes. A switch on the bottom of each solenoid simultaneously provides a signal to illuminate the green L BK DEICE ON and R BK DEICE ON annunciator lights. The annunciator panel lights are located forward of the pedestal in the center of the subpanel. If the airplane takes off without the control switch for the brake deice system having been switched OFF, a circuit is completed through the uplock switch to a timing circuit in the control module when the main landing gears reach the up and locked position. This timing circuit will cycle the deice system off after 10 minutes of operation, thereby closing the solenoid valve in the wheel well to shut off the flow of bleed air to the brakes so that adjacent components in the wheel well will incur no damage through overheating. Should the bleed air pressure activated tubes melt, LH and RH overheat pressure switches provide a signal to illuminate the amber L BK DI OVHT and R BK DI OVHT lights in the annunciator panel. The system cannot be activated until the landing gear is cycled from the up and locked position. Refer to Chapter 32-41-00 for Removal and Installation procedures of the Brake Deicer.

30-01-00

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

ICE AND RAIN PROTECTION AIRFOIL DESCRIPTION AND OPERATION

30-10-00 00

1. GENERAL A. Surface Deicer System The deicer and vacuum system is operated with pressure obtained by tapping off bleed air from the engine compressors (Ref. Figure 1). This air is routed through a regulator valve that is set to maintain the pressure required to inflate the deicer boots on the leading edge of each wing, the stabilons, and the horizontal stabilizers. To assure operation of the system should one engine fail, a check valve is incorporated in the bleed air line from each engine to prevent the escape of air pressure into the chamber of the inoperative compressor. The bleed air from the engine is also routed through an ejector that employs the venturi effect to produce vacuum for deflation of the deicer boots and operation of the instruments. The inflation and deflation phases of operation are controlled by means of a distributor valve. The deicer system is actuated by a three-way toggle switch on the LH subpanel. This switch is spring-loaded to return to the OFF position from either the MANUAL or SINGLE cycle position. When the switch is pushed to the SINGLE cycle position, one complete cycle of deicer operation automatically follows as the distributor valves open to inflate the deicer boots. The wing deicer boots inflate during the first six-second inflation period. The center section, stabilons and horizontal stabilizer deicer boots inflate during the second six-second inflation period. After the inflation period, a timer relay switches the distributor valves to OFF, or VACUUM position for deflation of the deicer boots. When the switch is pushed to the MANUAL position, all the boots will inflate and will continue to hold in the inflated positions as long as the switch is held in position. Upon release of the switch, the distributor valve returns to the OFF position and the deicer boots remain deflated until the switch is actuated again. The deicer system on this airplane is equipped with a combination vacuum and pressure line from a central distributor valve to provide pressure for inflation and vacuum for deflation of the boots. The vacuum line for the instruments is routed through a suction relief valve that is designed to admit into the system the amount of air required to reduce venturi vacuum sufficiently for proper operation of the instruments. CAUTION: Do not operate the deicer boots below -40°F. Exceeding this limit can result in permanent damage to the deicers.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

A

DETAIL

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UE30B 993956AA.AI

Figure 1 Surface Deicer

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

ICE AND RAIN PROTECTION AIRFOIL TROUBLESHOOTING

100100

1. PROCEDURES Table 101 Surface Deicer System PROBLEM 1. Lack of adequate vacuum to deflate boots completely.

2. Lack of adequate air pressure to inflate boots completely.

PROBABLE CAUSE

CORRECTIVE ACTION

a. Clogged lines.

a. Clean out lines.

b. Leaking lines.

b. Repair or replace lines as required.

c. Pressure regulator valve fails to operate properly or needs adjustment.

c. Adjust per Chapter 36-00-00 or replace as required.

d. Vacuum regulator valve fails to operate properly or needs adjustment.

d. Adjust per Chapter 37-00-00 or replace as required.

e. Distributor valve fails to operate properly.

e. Replace valve.

a. Clogged lines.

a. Clean out lines.

b. Leaking lines.

b. Repair or replace lines as required.

c. Pressure regulator valve needs adjustment.

c. Adjust per Chapter 36-00-00 or replace as required.

d. Distributor valve is not operating properly.

d. Replace valve.

3. Boots fail to inflate with switch in a. Tripped or malfunctioning circuit either SINGLE or MANUAL position. breaker.

a. Reset circuit breaker and check circuit as required.

4. Boots fail to inflate with switch in SINGLE position.

a. Surface deice control PCB assembly fails to operate properly.

a. Replace PCB assembly.

5. Boots inflate with switch in SINGLE position, but deflate instantaneously when the spring-loaded switch returns to OFF position.

a. Surface deice control PCB assembly fails to operate properly.

a. Replace PCB assembly.

6. Boots fail to inflate over 12 to 15 seconds with switch in SINGLE position.

a. Surface deice PCB assembly fails to operate properly.

a. Replace PCB assembly.

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Page 101 Nov 1/09

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Table 101 Surface Deicer System (Continued) PROBLEM

PROBABLE CAUSE

7. Boots inflation time either more or less than 5 to 8 seconds with switch in SINGLE position.

a. Surface deice PCB assembly fails to operate properly.

a. Replace PCB assembly.

8. Gage pressure below normal.

a. Gage not functioning properly.

b. Repair or replace as required.

b. Leakage in lines.

b. Repair or replace as required.

c. Pressure regulator valve needs readjustment or is not functioning properly.

c. Adjust per Chapter 36-00-00 or replace as required.

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30-10-00

CORRECTIVE ACTION

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

ICE AND RAIN PROTECTION AIRFOIL MAINTENANCE PRACTICES

200200

1. PROCEDURES Since the deicer boots and related components operate on clean air bled directly from the engine, little is required in the way of servicing the system. The only possible source of contamination is through the air entering the vacuum regulator valve in the vacuum system. The filter for the vacuum system is installed outside the pressurized area in the nose compartment. This filter should be periodically replaced according to the time limits stated in Chapter 5. Other than this, the only servicing required is for the vacuum regulator valve, which prevents the vacuum system from exceeding operating limits by allowing the proper amount of air to enter the vacuum lines. The vacuum regulator valve is mounted on the pressure bulkhead in the nose compartment. Since the vacuum regulator valve bleeds outside air into the vacuum system, it is essential that the regulator valve filter be kept clean. Frequency of cleaning the filter will vary with the conditions under which the airplane is operated; however, should it appear that the valve needs adjustment, especially to lower the vacuum, the filter should be cleaned and the setting rechecked before readjusting the valve. The valve can be removed for cleaning by disconnecting the lines from the valve and removing the retaining nuts. The valve should be cleaned in cleaning solvent (2, Table 1, Chapter 91-00-00) and dried with compressed air.

2. SURFACE DEICER BOOT A. Cleaning The surface of the deicer boots should be checked for indications of engine oil after servicing and at the end of each flight. Any oil spots that are found should be removed with a mild detergent soap and water solution. Care should be exercised during cleaning to avoid scrubbing the surface of the boots, as this will tend to remove the special Icex application. It should also be remembered during servicing of the airplane that the deicer boots are made of soft flexible stock, which may be damaged if refueling hoses are dragged over the surface of the boots, or if ladders and platforms are rested against them.

B. Surface Deicer Fast Boots Removal and Installation (Pressure Sensitive Adhesion) For airplanes equipped with Kit 129-4029-1, or any combination of Fast Boots, (Ref. Chapter 30 of the Model 1900 Airliner Series Component Maintenance Manual and B. F. Goodrich Pneumatic De-Icer Installation and Removal Instructions; 30-10-70, latest revision).

C. Removal CAUTION: Check the part identification of the deicer boot before using methyl propyl ketone on the deicer boot. Should the boot be labeled ESTANE, solvent (18, Table 1, Chapter 91-00-00) may be used on the boot. DO NOT ALLOW METHYL PROPYL KETONE TO COME IN CONTACT WITH ESTANE BOOTS. (1) Starting at one corner of the upper trailing edge of the boot, soften the adhesion line with solvent (14, Table 1, Chapter 91-00-00) on neoprene boots or solvent (18, Table 1, Chapter 91-00-00) for Estane boots. (2) Apply a minimum amount of solvent to the seam line while carefully applying tension to peel back the boot.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (3) Using a brush or trigger type oil can to apply solvent, separate the boot from the leading edge for a chordwise distance of four inches across the upper trailing edge. NOTE: The area between the boot and leading edge which has been separated will act as a reservoir for the solvent; consequently, the boot may be pulled forward with uniform tension. (4) From the centerline of the leading edge, continue to apply solvent while tension is being applied to undercut the adhesive until the boot is removed. (5) Thoroughly clean all adhesive from the leading edge with clean, grease-free cloths dampened with solvent (14, Table 1, Chapter 91-00-00) for neoprene boots or solvent (18, Table 1, Chapter 91-00-00) for Estane boots.

D. Installation CAUTION: Check the part identification of the deicer boot before using methyl propyl ketone on the deicer boot. Should the boot be labeled ESTANE, toluol may be used on the boot. DO NOT ALLOW METHYL PROPYL KETONE TO COME IN CONTACT WITH ESTANE BOOTS. (1) With 3/4-inch masking tape, mask off the area to be covered by the boot, allowing 1/2 to 3/4-inch margin beyond the actual area. Mask off accurately to avoid unnecessary clean up of excess adhesive. (2) Clean the area to be covered by the boot with a clean, grease-free cloth dampened with solvent (14, Table 1, Chapter 91-00-00) for neoprene boots or solvent (18, Table 1, Chapter 91-00-00) for Estane boots. Change cloths frequently, and never dip a used cloth in a clean supply of solvent (14, Table 1, Chapter 91-00-00) or solvent (18, Table 1, Chapter 91-00-00). (3) Reclean the metal surface with a clean cloth dampened with methyl propyl ketone for neoprene boots or toluol for Estane boots, then quickly wipe the surface dry with a clean, grease-free dry cloth. CAUTION: Use plastic, not metal, containers for Turco Metal-Glo No. 3 (15, Table 1, Chapter 91-00-00). It is also recommended that persons handling the cleaner wear rubber gloves, even though the cleaner is harmless to skin if immediately washed off with soap and water. (4) Vigorously scrub the metal surface with a clean, grease-free cloth dampened with Turco Metal-Glo No. 3 (15, Table 1, Chapter 91-00-00). After one minute contact, wipe the surface dry with a clean, grease-free dry cloth. Cover the cleaned surface with craft paper and wait an hour before rechecking the surface for cleanliness with a clean, dry, grease-free white cloth. If the clean cloth picks up any dirt, repeat the preceding Steps. (5) Carefully clean the rough back surface of the boot at least twice with a clean, grease-free cloth moistened with methyl propyl ketone for neoprene boots or toluol for Estane boots. Change cloths frequently to avoid recontaminating the cleaned portions of the boot. If necessary, continue cleaning the boot until completely clean. NOTE: If the finish on the back of the boot is smooth, mechanically roughen the bonding surface before cleaning. (6) Prepare test strips (Ref. the TEST CONTROL STRIPS procedure in this section).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (7) Open the container and thoroughly stir the adhesive (16, Table 1, Chapter 91-00-00) with a clean stick. Do not attempt to use adhesive that has jelled too much to drip from the stick. NOTE: Do not apply adhesive at temperature below 40°F or when the relative humidity is above 80%. If this requirement cannot be met, allow an extended drying time. Do not apply adhesive when the air is contaminated with dust or other elements. The following table will serve as a guide for adhesive application: Temperature In °F

Drying Time In Minutes

Above 80°

30

60° to 80°

45

40° to 60°

60

(8) Either of the two following methods may be used to apply the cement: NOTE: Brushing too long in one area will cause the cement to roll or ball up. (a) Brush one even coat of adhesive (16, Table 1, Chapter 91-00-00) on both the deicer boot back and metal surfaces. After permitting the adhesive to dry for at least 30 minutes, apply a second even coat to both surfaces. Allow the second coat to dry for at least 30 minutes, preferably an hour, before installing the deicer boot. Observe the above note concerning ball up for both applications. NOTE: For spraying consistency, dilute the adhesive with a solvent blend of 2 parts solvent (17, Table 1, Chapter 91-00-00) to 1 part solvent (14, Table 1, Chapter 91-00-00) until a No. 4 Ford cup will empty in 20 seconds. (b) Using a spray gun (Binks No. 7 with No. 63 PB head, De Vilbiss JGA 75X, or their equivalent), spray one even coat of adhesive (16, Table 1, Chapter 91-00-00) on both the deicer boot back and metal surfaces. After permitting the adhesive to dry for at least 30 minutes, spray a second even cross coat on both surfaces. Allow the second coat to dry for at least 30 minutes, preferably an hour, before installing the deicer boot. (9) Using No. 50 holes drilled into the leading edge at the extremities of boot locations as a guide, snap a chalk line along the leading edge of the airfoil. Trace the chalk line with a ball point pen and remove the chalk line with a dry cloth. The line should be intensified with a ball point pen after the adhesive is applied. (10) Securely attach hoses to the deicer boot connections without leaving fingerprints on the adhesive side of the deicer boot. Have a sufficient number of personnel present to handle the boot effectively. (11) Using a clean, lint-free cloth heavily moistened (but not dripping) with solvent (18, Table 1, Chapter 91-00-00), reactivate the adhesive on the airfoil leading edge and boot in spanwise strips about 6-inches wide and 24-inches long. Match the boot reference line to the leading edge chalk line, and roll the boot surface spanwise along the leading edge with a 2-inch rubber roller to remove all air from between the adjoining surfaces. Rubbing the adhesive surfaces excessively will remove some of the adhesive. If the boot is off course, carefully peel back the boot, using a minimum amount of methyl propyl ketone for neoprene boots or toluol for Estane boots, and reposition it properly.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL CAUTION: Do not allow excessive amounts of solvent to contact the boot; the solvent is to undercut the adhesive. If the adhesive is removed from either the boot or leading edge, it must be replaced and allowed to dry for 30 minutes minimum before continuing. (12) Continue to activate the surfaces and roll on each top half and lower half of the boot in sequence. Exerting pressure, use the 2-inch rubber roller on the entire surface of the boot parallel with the inflatable tubes and use the narrow steel stitcher roller on the trailing edge. Take care to avoid entrapping air throughout the rolling operation. If an air blister does occur, carefully remove the boot, using a brush dipped in methyl propyl ketone for neoprene boots or toluol for Estane boots, so that the air can escape, then press the surface down until the adhesive surface seals. (13) Using a sharp knife, trim the boot as required for proper fit. A 0.75-inch edge distance must be maintained between the inflatable tube and the edge of the trim. After the trimming process has been completed, re-roll the entire surface of the deicer boot. (14) Cut out the wing deice boot around the stall warning transducer and any inspection plates which fall beneath the boot installation to provide 0.02 to 0.06-inch of clearance around the perimeter of the transducer or inspection plates. (15) Remove all masking tape and clean excessive adhesive off with a clean white cloth dampened with methyl propyl ketone for neoprene boots or toluol for Estane boots, being careful not to allow the solvent to run under the edges of the boot. (16) Perform an adhesion test of the test strips for boot adhesion acceptability. (Ref. the TEST CONTROL STRIPS procedure in this section). (17) Install masking tape approximately 1/4-inch around the boot for uniform straight lines, then cover all exposed cement and fair around all cut edges and trailing edges of the boot with sealer (19, Table 1, Chapter 91-00-00). Remove the tape before the sealing compound cures. NOTE: The boots must not be operated for a 48 hour set time after installation. If the deicers are flexed before the 48-hour set time, the deicer boots must be checked for signs of lifting as soon as possible and be reinstalled or replaced if visible signs of lifting are present. The set time may be shortened (Ref. the SHORTENED 1300L DRY TIME FOR STANDARD PNEUMATIC DEICERS procedure in this section).

E. Shortened 1300L Dry Time For Standard Pneumatic Deicers NOTE: The following instructions are included to supply relief for the recommended 48 hours dry time for 1300L adhesive on pneumatic deicers. This applies to the deicers with 1 1/4-inch or narrower tubes operating at 20 psi or less pressure and is not applicable to heated deicers/ anti-icers with larger than 1 1/4-inch tubes, or to pneumatic deicers operating at pressures higher than 20 psi. CAUTION: Attempts to speed up drying time by applying heat externally through the deicers are not recommended since it is important for the bondline to be at the indicated temperature. The deicers will insulate the bondline from the heat being applied through it, and excess heat could cause deterioration of the deicers. NOTE: While the Chart allows a reduction in set time after installation, it is recommended that the actual set time before inflation of the deicers be extended as much as possible.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL The chart gives the reduced set times after installation of the applicable pneumatic deicers (Ref. Figure 202). The aircraft may be operated after this set time as long as the deicer system is not used until after the 48-hour period. If the deicers are flexed before a 48-hour dry time, the deicer boots must be checked for signs of lifting as soon as possible and be reinstalled or replaced if visible signs of lifting are present. (1) First determine the room temperature and humidity. (2) Locate the point on the Chart where the temperature (scale at the top of the Chart) and humidity (scale on the side of the Chart) intercept. NOTE: If the intercept point falls the Installation Not Recommended area, do not install the de-icers. If intercept point falls in the 24 Hour Dry Time area, allow installation to dry for 24 hours before inflating the de-icers. CAUTION: Allow at least the set time indicated by the Chart before inflating the deicers. If the deicers are flexed before a 48-hour dry time, the deicer boots must be checked for signs of lifting as soon as possible and be reinstalled or replaced if visible signs of lifting are present. (3) If intercept point falls in the middle area, read the recommended dry time by moving straight down to the set time scale (at the bottom of the Chart).

Figure 201 1300L Set Time for Standard Pneumatic Deicers

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3. TEST CONTROL STRIPS Deicer boot adhesion shall be checked through the use of a test strip applied at the time of boot installation. A one inch by eight inches test strip shall be prepared from a full thickness piece of boot material (scrap or previously rejected boot material). One adhesion test strip shall be prepared for each batch or mixture of adhesive used. The strip shall be applied in a convenient area adjacent to the boot. The metal and rubber surface area shall be prepared in the same manner as prescribed for the boot installation. NOTE: The test strip may be run on a separate aluminum panel of the same alloy as the leading edge skin. The same surface preparations and bonding requirements as previously stated shall apply. When applying the test strip, leave approximately an inch of one end free of adhesive for the purpose of attaching a clamp for the adhesion test. Allow the test strip to cure for the same time prescribed for the boot installation.

A. Adhesion Test Using a spring scale attached to a clamp on the test strip unbonded end, measure the force required to remove the strip at a rate of four-inches per minute. The pull shall be applied 180° to the surface (strip doubled back on itself). A minimum requirement of seven pounds tension shall be required to remove the test strip. In the event of failure, the acceptance of the boot installation shall be determined by the BOOT ACCEPTABILITY requirements in the following paragraph.

B. Boot Acceptability In the event of test strip failure, acceptability shall be based on the following: (1) Carefully loosen and lift one corner of the newly installed deicer boot sufficiently to attach a clamp. (2) Attach a spring scale to the clamp and pull with a force 180° to the surface in such a direction that the boot tends to be removed on a diagonal. (3) If a force of seven pounds-per-inch of width can be exerted under these conditions, the installation shall be considered as satisfactory. (4) Attach the boot corner (Ref. the SURFACE DEICER BOOT INSTALLATION procedure in this section). (5) Failure of the boot to meet the above requirements shall result in complete removal and reinstallation of the boot.

C. Surface Deicer Boot Age Master and Icex Application Rubber protect agent (29, Table 1, Chapter 91-00-00) is a liquid that protects rubber products from the ozone and extends the life of these parts. Age Master No. 1 and Icex compound are both products of the B.F. Goodrich Company. Compound (20, Table 1, Chapter 91-00-00) is a silicone base material specifically compounded to lower the strength of ice adhesion on the surfaces of deicer boots. Icex will not harm the boots and offers added ozone protection. When properly applied and renewed at periodic intervals, Icex provides a smooth, polished film that evens out microscopic irregularities on the rubber surfaces, so that ice formations have less chance to cling and are removed faster and cleaner when the boots are operated. Rubber protect agent (29, Table 1, Chapter 91-00-00) and compound (20, Table 1, Chapter 91-00-00) are applied as follows:

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (1) Wash the boot with mild soap and water. Thoroughly rinse the surface with clean water and allow to dry. (2) Solvent (30, Table 1, Chapter 91-00-00) may be used to remove substances which cannot be removed using soap and water. However, the boot must be washed again with mild soap and water, rinsed with clean water and allowed to dry. (3) Apply masking tape adjacent to the boots to be treated to prevent staining. CAUTION: To avoid the loss of vital components needed for the protective agent to penetrate the rubber and for fire safety precautions, do not spray Age Master No. 1. (4) Using a brush or three-inch trim roller or a two-inch by four-inch swab of lint-free cloth (wet but not dripping), apply a heavy coat of Age Master No. 1. Plastic or rubber gloves should be used with the swab method to prevent staining the skin. Apply at a rate of 0.4 to 0.5 fluid ounces per square foot (130 to 160 ml. per square meter). Cover the surface completely and evenly in uniform strokes. (5) Allow the first coat of Age Master No. 1 to dry from five to ten minutes. NOTE: A minimum two coats of Age Master No. 1 are required for a complete treatment. (6) Apply a second coat of Age Master No. 1 as outlined in Step (4) Apply a total of 0.75 fluid ounces per square foot (240 ml. per square meter) for effective protection. Allow the second coat to dry 20 to 30 minutes before handling. NOTE: After applying the second coat of Age Master No. 1 to the boot, let it dry for a minimum of 24 hours before applying Icex compound. Without further cleaning, Icex compound is then applied directly over the Age Master No. 1. (7) Allow the two coats of Age Master No. 1 to dry 24 hours. (8) Thin the compound (20, Table 1, Chapter 91-00-00) with Trichloroethane (21, Table 1, Chapter 91-00-00) to obtain a thin coat for application. (9) Use a clean cloth saturated with the thinned Icex compound for application. Hand buff with a clean cloth until a smooth glossy surface is obtained. NOTE: Both the Age Master No. 1 and Icex compound should be reapplied every 150 flight hours.

4. SURFACE DEICER BOOT REPAIRS NOTE: Patch sizes can be trimmed to accommodate small areas, as long as the patch extends a minimum of 1/2-inch beyond the damaged area on all sides. To maintain the maximum functional efficiency of the surface deicer boots, the following limits are allowed: NOTE: For pin hole repair, use Kit (172, Table 1, Chapter 91-00-00). Refer to instructions provided with the kit for repair procedure. • Pin holes 1/16 in. or smaller - no more than 20 pinhole repairs should be made per 12-inch square or • Small patches 1 1/4-inch x 2 1/2-inch - 3 patches per 12-inch square or • Medium patches 2 1/2-inch x 5-inch - 2 patches per 12-inch square or • Large patches 5-inch x 10-inch - 1 patch per 12-inch square or • 2 small patches and 1 medium patch per 12-inch square.

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A. Goodrich Pneumatic De-Icer Patch Replacement The following guidelines are recommended in assessing the need for immediate replacement of repair patches on Goodrich pneumatic de-icers (Ref. Figure 202). (1) If the lifted (debonded) area of the patch results in loss of air at the site of the patch, the patch should be replaced prior to the next flight. (2) If the lifted (debonded) area of the patch extends more than 0.19 inch inward from the edge of the patch, the patch should be replaced prior to the next flight. (3) If the lifted (debonded) area of the patch extends less than 0.19 inch inward from the edge of the patch and the patch is not leaking air, the patch should be replaced at the next available maintenance opportunity.

ALLOWABLE DE-BOND AREA

NON-ALLOWABLE DEBOND AREA

0.19 INCH

UE30B 092939AA.AI

Figure 202 Goodrich Pneumatic De-Icer Patch Replacement

B. Resurfacing Deicer Boots Resurface the seems if the surfacing material has abraded off or if the surfacing has developed cracks. The following procedures should be followed when resurfacing deicer boots: (1) Roughen the entire surface of the boot with fine sandpaper. (2) Clean the deicer boot thoroughly with solvent (18, Table 1, Chapter 91-00-00) or uncontaminated unleaded aviation gasoline.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (3) Clean the surface again with a clean lint-free cloth moistened with toluol or uncontaminated unleaded aviation gasoline. (4) Apply masking tape beyond the upper and lower trailing edges, leaving a 1/4-inch gap of bare metal. NOTE: If cement (22, Table 1, Chapter 91-00-00) has aged 3 months or more, it may be necessary to dilute it with solvent (18, Table 1, Chapter 91-00-00) to obtain the proper brushing consistency. Mix thoroughly approximately 5 parts cement to 1 part toluol. (5) Brush one coat of cement (22, Table 1, Chapter 91-00-00) on the boot and allow it to dry at least one hour. Then apply a second coat and allow it to dry at least four hours before operating the deicers. The airplane may be flown as soon as the cement is dry.

C. Scuff Damage This type of damage is most commonly encountered and generally will require no repair; however, close visual inspection when the boots are cycled may expose deep scuffs and cuts. On those occasions, if the entire ply surface (0.010-inch) has been removed, or if the boot is cut, it is necessary to patch the damage. B.F. Goodrich Kit 74-451-C (for neoprene boots) or Kit 74-451-H (for Estane boots) contain cold patches suitable for repairing the damaged area. B.F. Goodrich Kit 74-451-P is an acceptable alternate for Kit 74-451-H (for Estane boots). Cold patch repairs can be made as follows: (1) Clean the area around the damage with a clean, grease-free cloth dampened with solvent (18, Table 1, Chapter 91-00-00) or unleaded gasoline. (2) Using steel wool, buff the area around the damage so that it is moderately, but completely roughened. NOTE: A locally manufactured buffing shield, fabricated from any thin sheet material, will assure a neater repair. (3) Wipe the buffed area with a clean grease-free cloth dampened with solvent (18, Table 1, Chapter 91-00-00) to remove loose particles. (4) Select a repair patch which will extend at least 1/2-inch beyond the damaged area in all directions. (5) Apply one even coat of No. 4 cement, provided in the repair kit, to the patch and the damaged area. Allow the cement to dry until tacky. (6) Apply the patch to the boot with an edge, or the center, adhering first. Work the patch down carefully to avoid air pockets. Thoroughly roll the patch down with a stitcher roller. NOTE: The patches are manufactured to stretch in one direction only. Be sure to cut and apply the selected patch so that stretch is in the widthwise (chordwise) direction of the inflatable tubes. The length of the patch shall run along the length of the tubes. CAUTION: Check the part identification of the deicer boot before using methyl propyl ketone on the deicer boot. Should the boot be labeled ESTANE, toluol may be used on the boot. DO NOT ALLOW METHYL PROPYL KETONE TO COME IN CONTACT WITH ESTANE BOOTS. (7) Allow the patch to set for 10 to 15 minutes, then clean the patch and surrounding area with a clean cloth dampened with solvent (14, Table 1, Chapter 91-00-00) for neoprene boots or solvent (18, Table 1, Chapter 91-00-00) for Estane boots.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (8) Satisfactory adhesion of the patch will be reached in about four hours; however, the boot may be inflated to check the repair after a minimum of 30 minutes. (9) Apply a smooth light coat of sealer (19, Table 1, Chapter 91-00-00) to seal and feather the edges of the patch to the boot. (10) After the cement and sealing compound has dried and cured, apply a thin coating of compound (20, Table 1, Chapter 91-00-00) to the patch and surrounding area.

D. Tube Area Damage Repair cuts, tears or ruptures to the tube area as follows: (1) Clean the area around the damage with a clean, grease-free cloth dampened with toluol or unleaded gasoline. (2) Buff the area around the damage with a buffing stick, provided in the repair kit. (3) Wipe the buffed area with a clean, grease-free cloth dampened with toluol to remove loose particles. (4) Select a patch of ample size from the repair kit to cover the damaged area and extend at least 5/ 8-inch beyond the ends and edges of the cut or tear. NOTE: If none of the patches are the size required, cut a patch of the desired size from one of the larger patches. If this is done, bevel the edges by cutting with shears at an angle. (5) Complete the installation of the patch (Ref. SCUFF DAMAGE, Steps (4) thru (10) in this section).

E. Loose Surface Ply In Non-Inflatable Area Repair the damaged area as follows: (1) Peel and trim the loose surface ply to a point where adhesion of the surface ply is good. (2) Scrub (roughen) the area in which the surface ply is removed with steel wool. The scrubbing motion must be parallel to the cut edge of the ply to prevent the ply from loosening. (3) Scrub with steel wool and toluol directly over all edges, but parallel to the edges of the ply, to taper edges of the ply. (4) Cut a piece of surface ply material, provided in the repair kit, large enough to cover the damaged area and extend at least one inch beyond the damaged area in all directions. (5) Mask off an area 1 1/2-inches larger in length and width than the size of the damaged area. (6) Apply one even coat of No. 4 cement, provided in the repair kit, to the damaged area and the surface ply material. Allow the cement to dry until tacky. (7) Roll the surface ply to the deicer boot with a two-inch rubber roller, applying enough tension to prevent wrinkling. If air blisters appear, remove with a hypodermic needle and re-roll the area. (8) Remove the masking tape and complete the installation (Ref. SCUFF DAMAGE, Steps (7) thru (10) in this section).

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F. Loose Surface Ply In Tube Area Loose surface ply in the tube area is usually an indication of the boot starting to flex fail. This type of failure is easily detected, as a blister under the surface ply when the boot is pressurized. If detected while a small blister (1/4 to 3/8-inch in diameter) and immediately patched, service life of the boot will be appreciably extended. Repair of this type damage is the same as SCUFF DAMAGE.

G. Fabric Back Ply Damage During Removal If the adhesive from the leading edge skin has adhered to back surface of the boot, remove the adhesive with steel wool and methyl propyl ketone (for neoprene boots) or toluol (for Estane boots). In spots where the coating has pulled off the fabric, leaving bare fabric exposed, apply at least two additional coats of adhesive (16, Table 1, Chapter 91-00-00). Allow each coat to dry thoroughly.

5. SURFACE DEICER OPERATIONAL CHECKS A. With Engines Operating (1) With both engines operating at 70 to 80% N1 momentarily place the deicer switch in the SINGLE position; the deicer boots should inflate for a period of 5 to 8 seconds, then deflate to the vacuum hold-down condition. During inflation, the pneumatic pressure gage should register approximately 18 psi. (2) Hold the deicer control switch in MANUAL position for a few seconds. The deicer boots should inflate and remain inflated while the switch is retained in this position. Check for correct system pressure (Ref. BLEED AIR PRESSURE REGULATOR RELIEF VALVE ADJUSTMENT in this section). (3) Release the deicer control switch, permitting it to return to the OFF position; the deicer boots should deflate to the vacuum hold-down position. (4) Repeat Steps (1) thru (3) with each engine operating individually at 70 to 80% N1.

B. Without Engines Operating (1) Remove floor panels and carpeting forward of spar (Ref. Chapter 6-50-00). (2) Connect shop air to the inlet side of the pneumatic pressure regulator (Ref. Chapter 36-00-00 PRESSURE REGULATOR AND PNUEMATIC SYSTEM DIAGRAM) illustrations in the Maintenance Practices section. (3) With external power applied to the airplane - turn battery switch ON. (4) Momentarily place the deicer switch in the SINGLE position; the deicer boots should inflate for a period of 5 to 8 seconds, then deflate to the vacuum hold-down condition. During inflation, the pneumatic pressure gage should register approximately 18 psi. (5) Hold the deicer control switch in MANUAL position for a few seconds. The deicer boots should inflate and remain inflated while the switch is retained in this position. Check for correct system pressure (Ref. PRESSURE REGULATOR RELIEF VALVE ADJUSTMENT in this section). (6) Release the deicer control switch, permitting it to return to the OFF position; the deicer boots should deflate to the vacuum hold-down position.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (7) Turn the battery switch OFF and remove external power. (8) Disconnect shop air and restore pneumatic system. (9) Install floor panels and carpet.

6. BLEED AIR PRESSURE REGULATOR RELIEF VALVE ADJUSTMENT Refer to Chapter 36-00-00, REGULATOR/RELIEF VALVE ADJUSTMENT.

7. VACUUM REGULATOR VALVE ADJUSTMENT Refer to Chapter 37-00-00, VACUUM REGULATOR ADJUSTMENT.

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ICE AND RAIN PROTECTION AIR INTAKES DESCRIPTION AND OPERATION

30-20-00 00

1. GENERAL A. Engine Air Inlet Lip Anti-Ice The lip around each engine air inlet is heated by hot exhaust gases to prevent the formation of ice during inclement weather (Ref. Figure 201, Maintenance Practices section). A scoop in the engine LH exhaust stack deflects the hot exhaust gases and is ducted into the hollow lip tube that encircles the engine air inlet. The gases are exhausted out of the opposite exhaust stack. No shutoff or temperature indicator is necessary for this system.

B. Inertial Separation System An inertial separation system is built into each engine air inlet to prevent moisture particles from entering the engine inlet plenum during freezing conditions (Ref. Figure 202, Maintenance Practices section). When icing conditions are encountered, a moveable inertial vane and ice bypass door are lowered into the inlet airstream to induce an abrupt turn in the airflow before entering the engine plenum. The heavy, ice laden air is then discharged overboard through a bypass duct in the lower cowling at the aft end of the air duct. The inertial vane and bypass door are extended or retracted simultaneously through a linkage system connected to an electric actuator. The electric actuator for each engine consists of a primary and backup motor with a single actuator rod assembly. The system is normally driven by the primary motor. In the event the primary motor malfunctions, the backup motor will provide the power necessary to operate the system through the standby switch for the actuators. CAUTION: Should the actuator primary motor malfunction, the cause of the malfunction must be determined and corrected before the next flight. The primary actuator motors are energized through switches placarded ICE VANE - LEFT - RIGHT EXT - RET. The actuator switches are located on the lower left subpanel, to the left of the pilot's control column. The standby system may be activated through a second set of switches, located just below the primary ICE VANE switches. The standby system switches are placarded ACTUATORS - STANDBY MAIN. In normal operation, the ice vane is retracted out of the airstream and the bypass door is fully extended. When the ice vane is fully extended, a sense switch on each internal vane linkage will cause the L ENG ANTI-ICE and R ENG ANTI-ICE (green) lights to illuminate in the annunciator panel. When the control switches on the subpanel are activated, a second sense switch, also located on each inertial vane linkage, will energize a 30 to 40 second time-delay circuit. If full extension of the ice vanes is not attained in this time, L ENG ICE FAIL and R ENG ICE FAIL (amber) lights will illuminate in the annunciator signaling a fault in the prImary motor. These lights receive power through the annunciator power pcb. Full extension must be accomplished with the backup actuator motor. The ice vanes and bypass doors should be fully extended or fully retracted; there are no intermediate positions. When the ice vane is retracted, the annunciator lights will be off.

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ICE AND RAIN PROTECTION AIR INTAKES MAINTENANCE PRACTICES

200200

1. AIR INTAKE ANTI-ICE LIP (UA-1 AND AFTER, UB-1 THRU UB-60 WITHOUT KIT NO. 114-9014-1 S INSTALLED) Due to slight movement of the engine during operation, it is important that the bolts in the exhaust seal plate be left slightly loose. Only about three threads should show below the nut after installation. The slight movement allowed between the exhaust seal plate and the air intake anti-ice lip will alleviate stress on the exhaust duct and the anti-ice assemblies (Ref. Chapter 78-00-00 ENGINE EXHAUST SYSTEM illustration in the Maintenance Practices section).

2. AIR INTAKE ANTI-ICE LIP (UB-61 AND AFTER, UC-1 AND AFTER AND AIRPLANES WITH KIT NO. 114-9014-1 S INSTALLED) Flex pipes allow movement and alleviate stress on the exhaust duct and anti-ice assemblies caused by engine movement (Ref. Figure 201).

3. ENGINE AIR INLET ANTI-ICE LIP A. Removal (1) Remove the upper forward, upper aft and upper mid engine cowlings in that order (Ref. COWLING REMOVAL, Chapter 71-10-00). (2) Remove the lower forward engine cowling and perform the necessary service or replacement on the anti-ice lip.

B. Installation (1) Install the forward lower engine cowling. NOTE: For UA-1 and After, UB-1 thru UB-60 and Airplanes without Kit No. 114-9014-1 S. Due to the slight movement of the engine during operation, it is important that the bolts in the exhaust seal plate be left slightly loose. Only about three threads should show below the nut after installation. The slight movement allowed between the exhaust seal plate and the tube on the engine air inlet anti-ice lip will alleviate stress between the exhaust stack duct and the anti-ice lip. (2) Install the upper mid, upper aft and upper forward engine cowlings in that order (Ref. COWLING INSTALLATION, Chapter 71-10-00).

C. Crack Limits If the crack does not exceed the following criteria it may be disregarded and not repaired. (1) No exhaust leaks through the crack. (2) Crack length limit of 1.25-inch maximum.

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(3) Must be located on the outside of the anti-ice lip assembly between the lip assembly aft edge and the lip assembly leading edge. If any of the above criteria is exceeded, the anti-ice lip must be repaired (Ref. Model 1900 Airliner Series Structural Repair Manual, ENGINE ANTI-ICE LIP CRACK REPAIR, Chapter 54-90-08.

4. INERTIAL ANTI-ICING SYSTEM To help control wear and keep the inertial anti-icing system of the engine operating freely, the actuating linkage should be periodically lubricated with grease (23, Table 1, Chapter 91-00-00). NOTE: Do not lubricate with oil. Beyond lubrication, virtually no servicing of the inertial anti-icing system is required. Adjustments may be made on the inertial anti-icing system. Refer to RIGGING THE ENGINE INERTIAL ANTI-ICING SYSTEM, in this section.

A. Rigging the Engine Inertial Anti-Icing Systems The following procedure provides for the initial rigging of the inertial anti-icing system to ensure the proper adjustment of its various components with respect to one another throughout all phases of operation. Rather than rerigging the system in its entirety when only a portion of the system has been disassembled, or only a particular component is to be replaced, reference may be made to the applicable Steps within the procedure for the adjustment required. (1) Remove the cowling (Ref. COWLING REMOVAL, Chapter 71-10-00). (2) The rigging points for the normal dry air operation mode are identified as A, B, C, D, and E. The rigging points for the icing mode are identified as 1, 2, 3, 4, and 5 (Ref. Figure 203). (3) With the actuator disconnected at A, the bolt at B removed and the link installed at D and E, manually rotate the mechanism to the icing mode. (4) Position the link so that it is on center with the arm (Ref. Figure 204). (5) Energize the actuator to the full extend position and adjust the rod end, as necessary, to install the bolt at rigging point 1 (Ref. Figure 203). (6) Energize the actuator to the fully retracted position and inspect to assure free operation and adequate clearance for the inertial vane in the normal dry air position. (7) With the ice bypass door resting on the duct floor, adjust the clevis to install the bolt at B (Ref. Figure 204). (8) Energize the actuator to the fully extended position and inspect to assure free operation of the ice bypass door. (9) Reinstall the cowling (Ref. COWLING INSTALLATION, Chapter 71-10-00).

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Figure 201 Engine inlet Anti-Ice System (UB-61 and After, UC-1 and After and Airplanes With Kit No. 114-9014-1 S Installed)

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Figure 202 Inertial Anti-Icer System

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Figure 203 Rigging the Inertial Anti-Icing System

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Figure 204 Rigging the Overcenter Stop

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ICE AND RAIN PROTECTION WINDOWS AND WINDSHIELDS DESCRIPTION AND OPERATION

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1. GENERAL A. Windshield Anti-Ice The pilot's and copilot's windshields are protected against icing by internal heating elements. Separate switches for the PILOT and COPILOT are placarded WSHLD ANTI-ICE NORMAL - OFF - HI. The windshield deice switches are located on the subpanel, just to the right of the pilot's control column. Each windshield and its control switch are connected to the airplane electrical system through high heat and low heat relays and temperature controllers. The relays and temperature controller for the pilot are located on the aft side of the forward pressure bulkhead forward of the pilot's instruments. The relays and temperature controller for the copilot are situated on the aft side of the forward pressure bulkhead forward of the copilot's instruments. Each controller is protected by a 5-ampere circuit breaker located adjacent to the controller. Each system is protected by a 50-ampere limiter. The limiter for the pilot's windshield is located in the center of the LH nacelle on the electrical power distribution panel, in line with the wing leading edge, between Fuselage Stations 280.50 and 302.00. The copilot's windshield limiter is located in the center of the RH nacelle on the electrical power distribution panel, between Fuselage Stations 280.50 and 302.00.

B. Windshield Wiper The motor, arm assemblies, drive shafts, and converters of the windshield wiper installation are mounted forward of the instrument panel. The system is actuated by a control switch in the overhead panel and protected by a circuit breaker in the RH circuit breaker panel.

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ICE AND RAIN PROTECTION WINDOWS AND WINDSHIELD MAINTENANCE PRACTICES

200200

1. WINDSHIELD HEATING ELEMENTS A. Resistance Check (1) Remove all electrical power from the airplane and disconnect the battery. (2) Display warning notices prohibiting connection of the airplane’s electrical power. (3) Disconnect ground wire from terminal GRD on the windshield terminal block (Ref. Figure 201). (4) Using an ohmmeter, measure the resistance between terminals marked HP (High Power) and GRD (Ground). The reading should be 0.65 ohms ± 0.0975 ohms. (5) Measure the resistance between terminals LP (Low Power) and GRD (Ground). The reading should be 0.89 ohms ± 0.133 ohms. (6) Connect the ground wire to terminal GRD. (7) To verify sensor operation, remove any one wire from the terminal block marked S (Sensor). Using an ohmmeter, measure the resistance between the two S terminals. NOTE: The temperature sensor is a thermistor, which will change resistance with temperature. Sensor resistance should measure between 295 to 345 ohms for ambient temperatures between 60°F to 110°F. An open circuit will not allow the windshield to heat. (8) Connect wire removed from terminal S, in Step (7). (9) Connect battery and remove warning notices.

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Figure 201 Windshield Heating Elements

2. WINDSHIELD WIPER MOTOR A. Removal (1) To gain access to the motor, remove the attaching screws from the radio equipment panel, pull the panel out far enough to disconnect the electrical connectors, and remove the panel. (2) Remove the flexible drive shaft retaining nut and disconnect the electrical connector from the motor. (3) Remove the attaching screws and disengage the motor from the flexible drive shaft (Ref. Figure 202).

B. Installation (1) Connect the electrical connector to the motor and operate the motor to the PARK position. (2) Connect the flexible drive shaft to the motor with the retaining nut and secure the motor in place with the attaching screws (Ref. Figure 202). (3) Position the radio equipment panel, connect the electrical connectors and install the attaching screws to secure the panel.

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3. WINDSHIELD WIPER CONVERTER A. Removal (1) Perform the WINDSHIELD WIPER MOTOR REMOVAL procedure. NOTE: The motor must be removed before the LH converter can be removed, and both of the preceding units must be removed before the RH converter can be removed. (2) Rotate the adjustment bolt counterclockwise to relax tension on the wiper arm assembly (Ref. Figure 202). (3) Remove the wiper arm assembly. (4) Remove the screws securing the seal assembly to the fuselage. (5) Remove the flexible drive shaft from the side of the converter. (6) Remove the attaching bolts securing the converter in place. CAUTION: Avoid bending the ends of the flexible drive shafts when removing the converter.

B. Installation NOTE: First install the RH converter, then the LH converter, and finally the motor. (1) While viewing the RH converter from the wiper drive shaft side, rotate the converter’s internal drive shaft until the wiper drive shaft is in the extreme clockwise position (Ref. Figure 203). (2) Secure the RH converter in place with the attaching bolts. (3) Attach the flexible drive shaft to the RH converter. (4) While viewing the LH converter from the wiper drive shaft side, rotate the converter’s internal drive shaft until the wiper drive shaft is in the extreme counterclockwise position (Ref. Figure 203). (5) Connect the flexible drive shaft from the RH converter to the right side of the LH converter as it is placed into position. (6) Secure the LH converter in place with the attaching bolts. CAUTION: Make sure that the flexible drive shafts are properly engaged upon installation. (7) Secure the arm assembly drive shaft seal in place with the attaching screws. (8) Perform WINDSHIELD WIPER MOTOR INSTALLATION procedure.

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4. WINDSHIELD WIPER A. Adjustment NOTE: Never operate the windshield wiper on a dry windshield. Run the windshield wiper to the PARK position. If the wiper blades do not come to rest at approximately 2 inches from the center post of the windshield or if the tension of the wiper arm at the wiper attaching point is not within 4.5 to 5.5 pounds, adjust the windshield wiper in the following manner: (1) Lift the wiper arm away from the glass. Insert a 1/8 inch diameter pin through the hole on each side of the wiper arm. This will relieve blade tension against the glass. (2) Loosen the allen screw on the hub end of the wiper arm until the hub can be spread enough for removal (Ref. Figure 202). (3) Lift the wiper arm hub and reposition it two inches from the center of the windshield. An adjustment sleeve located on the converter shaft has both fine (inner teeth) and course (outer teeth) adjustments. (4) Install the arm assembly on the drive shaft, making sure that the arm is pulled up until the washer is in the counterbore of the hub. (5) Tighten the allen head screw until the wiper is anchored securely to the drive shaft. Safety wire the allen head screw with 0.032 inch safety wire. (6) Remove the pin from the holes in the wiper arm. (7) Set the wiper arm to a tension load of 4.5 to 5.5 pounds at the wiper attaching point. Rotate the spring adjusting screw on top of the wiper arm with the wrench (17, Table 7, Chapter 91-00-00). Turn the screw clockwise to decrease or counterclockwise to increase the adjusting spring tension.

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Figure 202 Windshield Wiper Installation

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A WIPER DRIVE SHAFT

WIPER DRIVE SHAFT

INTERNAL DRIVE SHAFT

EXTREME CLOCKWISE POSITION

INTERNAL DRIVE SHAFT

EXTREME COUNTER CLOCKWISE POSITION

RIGHT HAND CONVERTER

LEFT HAND CONVERTER

DETAIL

A

Figure 203 Windshield Wiper Converter Adjustment before Installation

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

ICE AND RAIN PROTECTION PROPELLER DESCRIPTION AND OPERATION

30-60-00 00

1. GENERAL A. Propeller Electric Deicer System The propeller electric deicer system includes: an electrically heated boot for each propeller blade, slip rings, brush assemblies, timer, on-off switch, and an ammeter (Ref. Figure 1). When the switch is turned on, the ammeter registers the amount of current (26 to 32 amperes) passing through the system. If the current rises beyond the switch limitation an integral circuit breaker will shut off power to the deicer timer. The current flows from the timer through the brush assemblies to the slip rings, where it is distributed to the individual propeller deicer boots. Heat from the boots reduces the grip of the ice, which is then removed by the centrifugal effect of the propeller rotation and the blast of the air stream. Two control switches are located on the LH subpanel, just to the right of the pilots control column and are placarded PROP-AUTO-MANUAL. When the AUTO switch is activated, power to the deice boots is cycled in 90-second phases. The first 90-second phase heats all the deicer boots on the RH propeller. The second phase heats all the deicer boots on the LH propeller. The deicer timer completes one full cycle every three minutes. A manual propeller deicer system is provided as a backup to the automatic system. When the MANUAL switch is activated power is supplied to the entire deice surface of both propellers. The manual override switch is of the momentary type, and must be held in place until the ice has been dislodged from the propeller surface. The load meters will indicate approximately a 0.05 increase of load when the manual propeller deicer system is in operation.

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Figure 1 Propeller Deicer Installation

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30-60-00 100100

ICE AND RAIN PROTECTION PROPELLER TROUBLESHOOTING

1. PROPELLER DEICER BOOT For troubleshooting of the propeller deicer systems, refer to the appropriate manual for B.F. Goodrich or Hartzell deicer systems in the Model 1900/1900C Airliner Series Component Maintenance Manual, P/N 114-590021-11 or subsequent.

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ICE AND RAIN PROTECTION PROPELLER MAINTENANCE PRACTICES

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1. PROPELLER DEICE AMMETER A. Calibration Should the ammeter be equipped with an adjusting screw on the back of the ammeter, the ammeter may be calibrated according to the following procedure: (1) Connect a voltmeter (accurate to 0.1-mv) to the positive and negative terminals on the back of the ammeter. (2) Turn the deice system on and measure the voltage to the ammeter. (3) Multiply the millivolt reading by 0.8; this product is what the ammeter should be indicating. (4) While lightly tapping the face of the ammeter, adjust the calibrating screw until the ammeter is indicating the correct value.

2. PROPELLER DEICER BOOT For maintenance of the propeller deicer systems, refer to the appropriate manual for B.F. Goodrich or Hartzell deicer systems in the Model 1900/1900C Airliner Series Component Maintenance Manual, P/N 114-590021-11 or subsequent.

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CHAPTER 32 - LANDING GEAR TABLE OF CONTENTS SUBJECT

PAGE

GENERAL INFORMATION 32-00-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1

MAIN LANDING GEAR 32-10-00 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Main Landing Gear Shock Absorber . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Servicing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Servicing - Inflation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Method 1 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Method 2 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Main Landing Gear . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Lubrication . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Main Landing Gear Torque Knee . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Main Landing Gear Drag Brace . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 201 201 201 202 202 202 202 202 203 205 205 205 206 206 207

NOSE LANDING GEAR 32-20-00 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Nose Landing Gear Shock Absorber . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Servicing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Servicing - Inflation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Method 1 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Method 2 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Nose Landing Gear Shimmy Damper . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Fluid Check (Manual Steering Only) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal (Manual Steering Only) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation (Manual Steering Only) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Nose Landing Gear . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Lubrication . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Nose Landing Gear Torque Knee . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Nose Landing Gear Drag Brace . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Bolt Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

32-CONTENTS

201 201 201 201 201 202 202 202 202 202 202 205 205 205 206 209 209 209 210 210 212 213

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LANDING GEAR EXTENSION AND RETRACTION 32-30-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Troubleshooting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 101 Time Delay Relay Printed Circuit Board (PCB) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 101 Functional Check. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 101 Hydraulic Power Pack Assembly . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 102 Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 102 Up-Pressure Switch . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 102 Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 102 Uplock Check Valve and Thermal Relief Valve . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 103 Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 103 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Landing Gear Component Access . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202 Hydraulic System Servicing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202 Hydraulic System Filling and Bleeding . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202 Landing Gear Hand Pump Cycling . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 206 Hydraulic System Leakage . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 207 Hydraulic Fittings Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 207 Hydraulic Line Filter Inspection and Cleaning (UC-143 and After and prior Airplanes with Kit No. 114-8022-1S Installed) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 209

LANDING GEAR HYDRAULIC ACCUMULATOR 32-30-01 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Servicing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 202 202 202

LANDING GEAR POWER PACK 32-30-02 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 201 202

LANDING GEAR POWER PACK - SPERRY VICKERS VALVE HOUSING AND CONTROLS 32-30-03 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Disassembly . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Assembly . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

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LANDING GEAR POWER PACK MOTOR 32-30-04 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 202 202

POWER PACK GEAR UP PRESSURE SWITCH 32-30-05 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 202 202

LANDING GEAR HYDRAULIC POWER PACK FILTERS 32-30-06 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .201 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202 Landing Gear Hydraulic Power Pack Bleed Air Filter . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 203 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 203 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .203 Cleaning . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 204

LANDING GEAR POWER PACK GEAR-UP AND GEAR-DOWN PORT FILTER SCREEN 32-30-07 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Cleaning . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 202 202 203

LANDING GEAR POWER PACK FLUID LEVEL SENSOR 32-30-08 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Functional Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 202 202 203

MAIN LANDING GEAR 32-30-09 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Rigging . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202

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MAIN LANDING GEAR ACTUATOR 32-30-10 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Main Landing Gear Actuator Shuttle Valve . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Functional Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Main Landing Gear - Actuator End Cap Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Preinspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Equipment Setup/Standardization . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Indication Evaluation Criteria . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Reporting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Post Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 202 202 204 204 207 207 208 209 210 210 210

MAIN LANDING GEAR ACTUATOR ORIFICE 32-30-11 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Cleaning . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 201 202 202

MAIN LANDING GEAR DOORS 32-30-12 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Rigging . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Hinge Wear Limits . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 202 203

NOSE LANDING GEAR 32-30-13 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Rigging . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202

NOSE LANDING GEAR ACTUATOR 32-30-14 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Shuttle Valve Functional Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Nose Landing Gear Actuator Ultrasonic Inspection. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

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NOSE LANDING GEAR DOOR 32-30-15 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Rigging . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201

HYDRAULIC LANDING GEAR SERVICE VALVE ASSEMBLY 32-30-16 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 202 202

EMERGENCY EXTENSION HAND PUMP ASSEMBLY 32-30-17 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Procedure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Bleeding . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 202 202 203

WHEELS AND BRAKES 32-40-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Main Gear Wheel Assemblies . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Nose Gear Wheel Assembly . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Tires . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Brake Assembly . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Hydraulic Brake System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2 Parking Brake Valves . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2 Troubleshooting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 101 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 101 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Tires . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Sealing Minor Leaks in Rim-Inflated Tubeless Tires . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Brake System Bleeding . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202 Pressure Bleeding . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202 Pressure Pot Bleeding Method . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202 Electric Bleeder Method . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 203 Gravity Bleeding Method . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 207 Brake Master Cylinder . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 207 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 207 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 207 Parking Brake Valve . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 208 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 208 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 208 Nose Wheel . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 211 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 211

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

CHAPTER 32 - LANDING GEAR TABLE OF CONTENTS (CONTINUED) SUBJECT PAGE Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 211 Main Wheel and Brake . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 212 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 212 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 213 Main Wheel . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 215 Removal (Using a Ramp) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 215 Installation (Using a Ramp) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 215 Brake Master Cylinder Linkage . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 218 Adjustment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 218 Parking Brake Control Cable . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 218 Adjustment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 218 Brake . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 219 Wear Limits . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 219 Fluid Reservoir Pressure Equalization Filter Cleaning . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 220 Fluid Reservoir Pressure Equalization Orifice Cleaning . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 221

ANTISKID BRAKES 32-41-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Manual Mode . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Antiskid Mode . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Servicing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Bleeding . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Gravity Bleeding Method . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202 Pressure Bleeding . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 203 Pressure Pot Bleeding Method . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 203 Electric Bleeder Method . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 204 Antiskid System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 205 Functional Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 205 Antiskid System Accumulator . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 209 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 209 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 209 Pressure Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 209 Antiskid Brake System Filter . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 210 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 210 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 210 Hydraulic Fittings . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 211 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 211 Solenoid Valve . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 211 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 211 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 212 Power Brake Valve . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 213 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 213 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 213 Pump and Motor . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 213 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 213 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 214

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CHAPTER 32 - LANDING GEAR TABLE OF CONTENTS (CONTINUED) SUBJECT

PAGE

Antiskid Brake System Wheel-Speed Transducer . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 216 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 216 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 216

BRAKE DEICE SYSTEM 32-42-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Distributor Manifold . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Shutoff Valve . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 204 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 204 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 204 Pressure Test and Functional Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 204 Overtemperature Warning Functional Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 205

MECHANICAL STEERING 32-50-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Mechanical Steering Mechanism . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202 Disconnect Actuator . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 203 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 203 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 203 Forward Steering Link Boot . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 204 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 204 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 204 (With Boot) Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 204 Mechanical Steering Nose Gear . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 205 Centering Adjustment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 205 Stop Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 205 Stop Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 205

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POWER STEERING (UA-1 AND AFTER; UB-1 AND AFTER) 32-51-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Troubleshooting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 101 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 101 Troubleshooting Notes (UA-1 and After; UB-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 101 Signal Amplifier (UA-1 and After; UB-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 102 System Control (UA-1 and After; UB-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 103 Breakout Unit (UA-1 and After; UB-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 103 System Calibration (UA-1 and After; UB-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 104 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 System Servicing (UA-1 and After; UB-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 System Filter Replacement (UA-1 and After; UB-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Hydraulic Fittings Installation (UA-1 and After; UB-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Power Steering Actuator . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Maintenance Checks (UA-1 and After; UB-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Removal (UA-1 and After; UB-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202 Installation (UA-1 and After; UB-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202 Repairing (UA-1 and After; UB-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 204 Servo Valve and Manifold (UA-1 and After; UB-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 204 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 204 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 204 Arming Solenoid (UA-1 and After; UB-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 205 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 205 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 205 Switch (UA-1 and After; UB-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 205 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 205 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 205 Pump and Motor (UA-1 and After; UB-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 209 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 209 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 209 Solenoid Valve (2-Position) (UA-1 and After; UB-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 209 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 209 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 209 Signal Amplifier (UA-1 and After; UB-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 210 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 210 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 210 Disconnect and Park Disconnect Switch (UA-1 and After; UB-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . 212 Adjustment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 212 Nose Wheel Centering Mechanism (UA-1 and After; UB-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . 215 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 215 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 215 Adjustment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 215 Nose Gear Position Switch Adjustment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 216

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POWER STEERING (UC-1 AND AFTER) 32-52-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Power Steering Amplifier (UC-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2 Troubleshooting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 101 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 101 Troubleshooting Notes (UC-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 101 Airplane Configuration (UC-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 102 System Configuration (UC-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 102 System Checks (UC-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 102 Required Test Equipment (UC-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 104 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Command Potentiometer . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Removal (UC-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Installation (UC-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202 Rigging (UC-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202 Command Potentiometer and Amplifier Adjustments (UC-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . 203 Hydraulic Filter . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 204 Servicing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 204 Supply and Return Filter Union (UC-143 and After; and earlier UC serials with Kit No. 114-8021-1S installed) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 204 Servicing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 204 Hydraulic Reservoir (UC-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 205 Servicing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 205 Actuator (UC-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 205 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 205 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 206 Servo Valve (UC-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 206 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 206 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 207 Arming Valve (UC-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 207 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 207 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 208 Manual Steering Fail Switch (UC-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 209 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 209 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 209 Feedback Potentiometer (UC-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 209 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 209 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 209 Amplifier AR103 (UC-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 210 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 210 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 210 Relay Panel A320 (UC-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 210 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 210 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 211 Hydraulic Pressure Switch (UC-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 211 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 211 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 211

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Pump Relief Valve (UC-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Pump and Motor (UC-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Hydraulic Reservoir (UC-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Nose Gear Centering Mechanism (UC-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Nose Wheel Centering (UC-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Adjustment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

211 211 211 212 212 212 212 212 212 213 213 213 214 214

LANDING GEAR POSITION AND WARNING 32-60-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 general . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Troubleshooting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 101 Procedure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 101 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Nose Gear Up-Position Indicator Switch . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Adjustment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Main Gear Up-Position Indicator Switch . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202 Adjustment (UA-3, UB-1 thru UB-45, Except UB-37, Without Kit No. 114-8005 Installed) . . . . . . . . . 202 Adjustment (UB-37, UB-46 and After, UC-1 and After and Prior Airplanes With Kit No. 114-8005 Installed) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202

NOSE GEAR DOWN-POSITION SWITCH 32-60-01 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Adjustment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201

MAIN GEAR DOWN-POSITION SWITCH 32-60-02 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Adjustment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201

NOSE GEAR ACTUATOR DOWNLOCK SWITCH 32-60-03 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Adjustment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201

Page 10 Nov 1/13

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

CHAPTER 32 - LANDING GEAR TABLE OF CONTENTS (CONTINUED) SUBJECT

PAGE

MAIN GEAR ACTUATOR DOWNLOCK SWITCH 32-60-04 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Adjustment (All Airight Actuators Except P/N 114-380041-1) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Adjustment (Phoenix Controls Actuators and Airight Actuator P/N 114-380041-1) . . . . . . . . . . . . . . . Adjustment (Frisby Actuators) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 201 202 203

MAIN GEAR SAFETY SWITCH 32-60-05 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Adjustment (444EN49-6 Switch) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Adjustment (39EN6-6 Switch) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 201 203

LANDING GEAR WARNING HORN 32-60-06 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Landing Gear Warning Horn . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Adjustment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

32-CONTENTS

201 201 201 202

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

LANDING GEAR GENERAL INFORMATION DESCRIPTION AND OPERATION

32-00-00 00

1. GENERAL WARNING: When an airplane has experienced abnormal landing gear procedures of any type, as a safety precaution, place the airplane on jacks prior to performing any inspection or maintenance. Ensure that all three landing gears are down and locked prior to removing the airplane from the jacks. CAUTION: Jacking of an airplane for the purpose of landing gear operation, inspection, servicing or maintenance, should be accomplished within an enclosed building or hangar. In the interest of safety, should it become necessary to jack the airplane in the open, wind velocity in any direction and terrain variations must be compensated for prior to jacking the airplane. The 1900 Airliner Series airplanes are equipped with a retractable hydraulic landing gear system. The nose gear and main gear assemblies incorporate air-oil type shock struts that are filled with both compressed air and hydraulic fluid. The nose gear strut is equipped with a single, wheel and tire, while each main gear strut is equipped with two wheels and tires. The landing gear is retracted and extended by the action of the individual actuators and drag brace assemblies connected to each landing gear. The nose gear is attached to two longitudinal fuselage members in the fuselage nose and pivots aft into the wheel well when retracted. The main landing gear are attached to the two main structural ribs of the nacelles and pivot forward into the wheel wells when retracted. The landing gear doors consist of one nose gear door and two sets of main gear doors. The nose gear door is hinged at the front and is connected to the nose gear brace with two push-pull links; when the landing gear is retracted, the door is pulled closed and when extended, the door is pushed open. The main gear doors are hinged at the sides and are spring-loaded to the open position. As the landing gear is retracted, two rollers on each main gear engage the door actuating cams and pull the doors closed. When the landing gear doors are closed, they cover the top braces of each landing gear. The nose and main landing gear assemblies are extended and retracted by a hydraulic system consisting of an actuator located in each wheel well, a hydraulic power pack located in the LH wing, and hydraulic plumbing. The hydraulic plumbing provides for normal extend, normal retract and emergency extend modes of operation. A landing gear control switch, placarded UP and DN, is located on the pilot's inboard subpanel. When the control handle is moved to the UP position, power is supplied to the pump motor and to the gear-up solenoid to allow system fluid under pressure to flow to the retract side of the system. The landing gear is held in the retracted position by positive hydraulic pressure from the pump. When the control handle is placed in the DN position, power is supplied to the pump motor and to the gear-down solenoid to allow system fluid under pressure to flow to the extend side of the system. When the actuator pistons are positioned to fully extend the landing gear, an internal mechanical lock in each actuator will lock the actuator pistons to hold the landing gear in the down position. Manual landing gear extension is provided through a manually powered hydraulic system. A hand pump, located on the floor between the pilot's seat and the pedestal, is used when emergency extension of the gear is required. To extend the landing gear with this system, the pump handle is removed from its securing clip and pumped up and down. As the handle is pumped, hydraulic fluid is drawn from the power pack into the pump and exited under pressure. Fluid under pressure from the pump is routed to each landing gear actuator. A service valve plumbed in the hydraulic system may be used in conjunction with the hand pump to raise the landing gear for maintenance purposes.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Visual indication of the landing gear position is provided by two red in-transit lights located in the landing gear control switch handle and a green gear-down indicator light assembly on the subpanel. The in-transit lights are wired through the automatic light dimming circuit. Illumination of the in-transit lights indicates when the landing gear is in transit; gear up is indicated when they go out. Illumination of the gear-down lights indicates when each landing gear is down and locked. Up-and-down position switches and actuator downlock switches are located in each wheel well to complete the electrical circuits necessary to illuminate the gear down and in-transit lights. Nose gear steering is accomplished by a mechanical linkage system extending from the rudder pedals to a bell crank mounted in the nose gear wheel well above the nose gear assembly. A spring mechanism in the steering linkage permits tow bar or braked steering angles greater than that provided by the rudder pedal system. An electrical actuator and a cam are incorporated in the steering linkage to remove the steering action from the rudder pedals while the airplane is in flight. An optional power steering system incorporates a hydraulically powered rotary actuator mounted on top of the nose gear assembly. Hydraulic power for the actuator is provided by a pump located in the LH wheel well. The power steering system uses the landing gear power pack as its fluid supply. Command potentiometers controlled by the pilot's rudder pedals provide the steering input signals for the electrical portion of the steering system. A servo valve in the actuator receives signals from the potentiometers to adjust the pressure inside the actuators to move the nose wheel left or right. The airplane is equipped with four hydraulically operated brake assemblies. Each main landing gear incorporates two multidisc, metallic-lined brake assemblies bolted together, one on each side of the strut. Each brake assembly contains two rotating discs, which are keyed to rotate with the wheel, a stationary disc, and a pressure plate and backplate which are attached to the brake housing. Braking action occurs when hydraulic pressure is applied to the pistons in the brake housing, forcing the disc stack together and creating friction between the rotating discs and the stationary components of the brake assembly. The hydraulic pressure to the brake assemblies is generated by the pilot's or copilot's master cylinders. An optional antiskid brake system incorporates a power brake valve, a pump and motor, a skid control unit, and a wheel speed transducer in each main gear axle. The electric motor-driven pump provides the hydraulic power for the system. When the antiskid system is off, the pressure at the brake assemblies is the same as the master cylinder pressure. When the system is in the antiskid mode, the master cylinder pressure is isolated from the brake assemblies and a braking pressure balanced to the master cylinder pressure is generated by the power brake valve. The objective of the antiskid function is to give optimum braking effectiveness. This is accomplished as the wheel transducers monitor the wheel RPM and relay the information to the skid control unit. When a skid condition is imminent, the appropriate signal is sent to the power brake valve to reduce the pressure to all four brake assemblies. As the wheel RPM begins to increase, the power brake valve increases the braking pressure; if the wheel transducers indicate that the wheels are still in a skid condition, the brake pressure will be decreased again. This cycle will continue until the tendency to skid has ended. Brake deicing is provided by an optional brake deice system which utilizes engine bleed air as the deicing agent. The brake deice system is composed of a combination of tubes and hoses plumbed from the engine bleed air duct forward of each firewall to a distributor manifold installed on each main gear shock strut. A shutoff valve in each main gear wheel well controls the flow of heated air to each distributor manifold. An overtemp warning line is routed in close proximity to the bleed air lines in each wheel well to notify the operator in the event of overheating of the deice system.

Page 2 Nov 1/09

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WARNING: Any time maintenance is performed on the landing gear system, always place the airplane on jacks. When jacking the airplane in an unsheltered area where winds in excess of 35 kts may be encountered, never jack more than one gear clear of the ground at a time. Any time the landing gear is only partially retracted during maintenance, always cycle the gear with the power pack through at least one complete cycle before removing the airplane from the jacks. As an added security measure, manual landing gear downlocks shown under Special Tools in Chapter 91-00-00 may be purchased or may be fabricated from Kit instructions no. 114-5015 S. The downlocks or the kit may be ordered through any Hawker Beechcraft outlet.

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LANDING GEAR MAIN LANDING GEAR MAINTENANCE PRACTICES

32-10-00 200200

1. PROCEDURES WARNING: When performing maintenance on a hydraulically operated landing gear system, be aware that any movement of a hydraulic actuator cylinder can displace hydraulic fluid and cause unanticipated movement of other actuator cylinders in the landing gear retraction system. Servicing of the landing gear hydraulic accumulator can also result in unanticipated movement of an actuator. Either action can result in an unsafe, unlocked landing gear system. Therefore, place the airplane on jacks prior to performing any inspection or maintenance. Cycle the landing gear and ensure that all three landing gears are down and locked prior to removing the aircraft from jacks.

2. MAIN LANDING GEAR SHOCK ABSORBER A. Servicing (1) To check the fluid level and fill the landing gear shock absorber, remove the valve cap and press the valve core to release the air. Allow the strut to fully compress. WARNING: Release the air pressure entirely before removing the valve core. (2) Remove the valve core and connect one end of a 1/4 inch I.D. hose to the valve stem. Submerge the other end of the hose in a container of clean hydraulic fluid (39, Table 1, Chapter 91-00-00). (3) Slowly extend the strut to draw hydraulic fluid into the strut. (4) When the strut is fully extended, slowly compress the strut to expel any excess hydraulic fluid and trapped air. (5) Cycle the strut as many times as necessary to expel all of the air. With the strut compressed, remove the hose and install the valve core. CAUTION: Never tow or taxi with a flat strut. Even brief towing or taxiing with a deflated strut can cause damage.

B. Servicing - Inflation WARNING: Anytime maintenance is to be performed on the landing gear system, always place the airplane on jacks. When jacking the airplane in an unsheltered area where winds in an excess of 35 kts may be encountered, never jack more than one gear clear of the ground at a time. Any time the landing gear is only partially retracted during maintenance, always cycle the gear with the power pack through at least one complete cycle before removing the airplane from the jacks. CAUTION: Loss of oil will result in struts appearing to need inflation. Low oil levels will damage the strut and could result in hard landing damage to the aircraft. If loss of oil is detected or suspected, check oil level using the MAIN LANDING GEAR SHOCK ABSORBER SERVICING procedure.

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C. Method 1 (1) With the airplane on the ground and empty, except for full fuel and oil, inflate the main strut with nitrogen or dry filtered air until the piston is extended 5.12 to 5.62 inches.

D. Method 2 (1) With the airplane on jacks, inflate the main landing gear with nitrogen or dry filtered air to 525 psig.

3. MAIN LANDING GEAR A. Lubrication Lubricate the main landing gear wheel bearings and grease fittings (Ref. Chapter 12-20-00, LUBRICATION SCHEDULE).

B. Removal (1) Perform the THREE-POINT JACKING (PREFERRED (Ref. Chapter 07-10-00). All tires must be clear of the floor.

PROCEDURE)

procedure

(2) Retract the main landing gear slightly to take the load off the drag brace and to unlock the actuator (Ref. Figure 201). NOTE: The nut next to the switch must be maintained in the same position to provide the correct adjustment when installing the switch. (3) Remove the down-position switch and wiring clamps from the drag leg. (4) Disconnect the drag leg from the main gear by removing the cotter pin, nut, washers, bolt (8) and bushings (7). After disconnecting the drag leg, install the bolt, bushings, washers and nut into the drag leg to facilitate correct installation. (5) Disconnect the hydraulic brake flex hose at the top of the gear and cap the hydraulic lines. (6) On installations equipped with brake deicing, remove the clamp securing the brake deicer hose to the upper end of the strut, then disconnect the hose at the union in the bracket just above the torque knees. (7) Disconnect the wiring (safety and downlock switch) from the receptacle at the upper rear of the wheel well. (8) Remove the access panels covering the landing gear hinge bolts. Remove the cotter pin, nut (6), washers (2, 3 and 5), hinge bolt (1) and bushing (4) and lower the main landing gear away from the airplane. NOTE: After removing the bolt (1), position the washers on the bolt in the same sequence as when installed.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

C. Installation (1) Coat the bushings (4 and 7) and bolts (1 and 8) with lubricating grease (61, Table 1, Chapter 91-00-00) at the time of installation (Ref. Figure 201). NOTE: Install the washers on the main gear hinge bolts in the same sequence as when removed. (2) Position the gear, install the bushings (4) and install the hinge bolts (1) through the strut and keel. Install the bolt heads toward the center of the strut. Install the washers (3) (minimum of one each side of the strut) between the keel and strut as required to provide a total clearance of 0.005 to 0.030 inch between the keel and washer surfaces. Apply a minimum torque of 800 inch-pounds to the hinge bolt nuts; tighten the nuts to the next castellation (but do not exceed 1200 inch-pounds) and install the cotter pin. One washer (5) may be used under the nut to obtain the required torque value. (3) Connect the hydraulic brake flex hose to the union on the landing gear. (4) On installations equipped with brake deicing, connect the brake deicer hose at the union in the bracket just above the torque knees. Install the clamp securing the hose to the upper end of the strut. (5) Slowly move the lower drag leg in to position between the MLG lugs without any side pressure being applied to the drag leg. If the drag leg does not align with the attachment lugs on the landing gear strut, perform the MAIN LANDING GEAR DRAG BRACE INSTALLATION procedure. (6) Connect the lower drag leg to the landing gear with the bolt (8), bushings (7), washers, nut and cotter pin. With a minimum of one washer under the nut and the head of the bolt (8), tighten the nut and bolt finger-tight, then tighten to the next slot of the castle nut to permit installation of the cotter pin. NOTE: The nut next to the switch must be maintained in the same position to provide the correct adjustment as when previously installed. (7) Install the downlock switch and wiring clamps on the drag leg. (8) Plug the down-position switch and safety switch wiring into the receptacle at the upper rear of the wheel well. (9) Install the landing gear hinge bolt access panels. (10) Perform the MAIN LANDING GEAR SHOCK ABSORBER SERVICING procedure contained in this section. (11) Bleed the brakes (Ref. 32-40-00, BRAKE SYSTEM BLEEDING or 32-41-00, BLEEDING). (12) Perform a retraction check (Ref. 32-30-00, LANDING GEAR HAND PUMP CYCLING). (13) Lubricate the main gear hinge points and drag brace joints with lubricating grease (61, Table 1, Chapter 91-00-00). (14) Perform the LOWERING THE AIRPLANE AFTER THREE-POINT JACKING procedure (Ref. Chapter 07-10-00).

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Figure 201 Main Landing Gear Assembly

Page 204 Nov 1/13

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4. MAIN LANDING GEAR TORQUE KNEE A. Removal (1) Perform the THREE-POINT JACKING (PREFERRED (Ref. Chapter 07-10-00). All tires must be clear of the floor.

PROCEDURE)

procedure

(2) Deflate the landing gear strut. NOTE: Cap the hydraulic brake line to prevent contamination of the hydraulic fluid. (3) Remove the wheels and brakes (Ref. 32-40-00, MAIN WHEEL AND BRAKE REMOVAL). (4) Remove the cotter pin, nut, washer and bolt (1) that attach the safety switch arm to the upper torque knee (Ref. Figure 202). (5) On airplanes with antiskid brakes, remove the clamps securing the wheel speed transducer wiring harness to the upper and lower torque knees. (6) Remove the cotter pin (10), nut (2), washers (3, 4, 5, 6, and 7), bushing (8) and bolt (9) that connect the upper and lower torque knees (13 and 14). Note the position of the washers to facilitate installation of the torque knees. (7) Cut the safety wire from the clevis pin (11) and remove the clevis pin from the upper torque knee pin (12). Tap the upper torque knee pin out and remove the upper torque knee (13). (8) Remove the cotter pin, washers and clevis pin (15) from the lower torque knee pin (16). Tap the lower torque knee pin (16) out and remove the lower torque knee (14). (9) Inspect the upper and lower torque knee bushings for wear and/or damage (Ref. Model 1900 Airliner Series Component Maintenance Manual, Chapter 32-10-00, MAIN LANDING GEAR MANUFACTURING AND WEAR TOLERANCES AND INSPECTION PROCEDURES).

B. Installation (1) Coat the torque knee pins (12 and 16), bushing (8), and bolt (9) with lubricating grease (61, Table 1, Chapter 91-00-00) at the time of installation (Ref. Figure 202). (2) Position the lower torque knee (14) in place. Tap the lower torque knee pin (16) in place and install the clevis pin (15), washer and cotter pin. (3) Position the upper torque knee (13) in place. Tap the upper torque knee pin (12) in place and install the clevis pin (11). Safety wire between the clevis pin and the safety wire hole in the forward side of the retainer ring. CAUTION: Do not create a tight joint between the upper and lower torque knees. The torque knee connection joint must move freely. (4) Connect the upper and lower torque knees as follows: (a) Align the upper torque knee (13) and the lower torque knee (14) and position the washer (5) between them. Insert the bushing (8) through the torque knees and washer (5). (b) Insert the bolt (9) through the washers (7, 6 and 5) and the bushing (8).

32-10-00

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (c) In the following Step, to obtain minimum bolt end play, use AN960-816 and /or AN960D816L washer(s) as required after washer (4). Use AN960-616 and/or AN960-616L washer(s) after washer (3) as required to align the cotter pin (10). (d) Install washers (5, 4 and 3), nut (2) and a new cotter pin (10). (5) On airplanes with antiskid brakes, install the clamps securing the wheel speed transducer wiring harness to the upper and lower torque knees. (6) Install the bolt (1), washer, nut and cotter pin that attach the safety switch arm to the upper torque knee. (7) Install the wheels and brakes (Ref. 32-40-00, MAIN WHEEL AND BRAKE INSTALLATION). (8) Lubricate the torque knee joints with lubricating grease (61, Table 1, Chapter 91-00-00). (9) Perform the LOWERING THE AIRPLANE AFTER THREE-POINT JACKING procedure (Ref. Chapter 07-10-00). (10) Perform the MAIN LANDING GEAR SHOCK ABSORBER SERVICING procedure in this section.

5. MAIN LANDING GEAR DRAG BRACE A. Removal (1) Perform the THREE-POINT JACKING (PREFERRED (Ref. Chapter 07-10-00). All tires must be clear of the floor.

PROCEDURE)

procedure

(2) Retract the gear slightly to take the load off the drag brace and to unlock the actuator (Ref. Figure 203). NOTE: To maintain switch adjustment do not permit the nut on the switch side of the rig plate (7) to turn. (3) Remove the wiring clamps and downlock switch (8). (4) Remove the cotter pin (13), nut (12), washers (10) and bolt (9) attaching the lower drag leg (15) to the landing gear lug and remove the lower drag leg (15). Leave the bushings (11) in the strut (14) and attach the bolt (9), washers (10) and nut (12) to the strut (14). (5) Remove the nut (1), washers (2) and bolt (3) attaching the actuator (4) rod end to the upper drag leg arm (25). (6) Remove the cotter pin (21), nut (22), washers (23) and bolt (24) securing the upper drag leg (5) to the wheel well attach fitting. NOTE: Note the number of washers (23) on each side of the upper drag leg (5) to facilitate installation of the drag brace in alignment with the attachment fitting on the landing gear strut (14).

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B. Installation NOTE: A shimming procedure must be completed prior to installing the drag brace assembly (Ref. Model 1900 Airliner Series Component Maintenance Manual, Chapter 32-10-00, MAIN LANDING GEAR DRAG BRACE SHIMMING PROCEDURE). If the upper and lower drag legs (5 and 15) were separated during overhaul make sure that two washers (6) are placed on each side of lower drag leg (15). Install the center pivot bolt (16), washers (17), bushings (18) and nut (19) through upper and lower drag legs (5 and 15) upon assembly. Torque the center pivot bolt (16) and nut (19) 200 to 400 inch-pounds and install cotter pin (20). The pivot joint of the drag brace should be free of binding. (1) Lubricate the bolts (9 and 24) and bushings (11) with lubricating grease (61, Table 1, Chapter 91-00-00) at the time of installation (Ref. Figure 203). (2) Insert the upper drag leg bolt (24) through the wheel well fitting and upper drag leg (5). Install the washers (23) between the upper drag leg (5) and the wheel well fitting in the same positions as when removed. With a minimum of one washer (23) under the head of the bolt (24) and the nut (22), tighten the nut and bolt finger-tight, then tighten to the next slot of the castle nut to permit installation of the cotter pin (21). NOTE: Washers (23) may be shifted from side to side of the wheel well fitting and upper drag leg (5) to align the lower drag leg (15) with the attachment point on the landing gear strut (14). A minimum of one washer (23) should be maintained between the upper drag leg (5) and the wheel well attachment fitting. (3) Attach the lower drag leg (15) to the landing gear lug with the attach bolt (9), bushings (11), washers (10), nut (12) and cotter pin (13). With a minimum of one washer (10) under the nut (12) and the head of the bolt (9), tighten the nut and bolt finger-tight, then tighten to the next slot of the castle nut to permit installation of the cotter pin (13). (4) Lubricate bolt (3) with grease (80, Table 1, Chapter 91-00-00) at time of installation. (5) Attach the actuator (4) rod end to the upper drag leg arm (25) with the bolt (3), washers (2) and nut (1). NOTE: To maintain switch adjustment, do not permit the nut on the switch side of the rig plate (7) to turn. (6) Install the downlock switch (8) and attach the wiring clamps to the drag leg. NOTE: No side force may be applied to the lower drag brace to facilitate alignment with the attach point on the landing gear strut. (7) Lubricate the drag brace joints with lubricating grease (61, Table 1, Chapter 91-00-00). (8) Perform a rigging check (Ref. 32-30-00, MAIN LANDING GEAR RIGGING). (9) Perform the LOWERING THE AIRPLANE AFTER THREE-POINT JACKING procedure (Ref. Chapter 07-10-00).

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Figure 202 Main Landing Gear Torque Knee Assembly

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Figure 203 Main Landing Gear Drag Brace Assembly

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LANDING GEAR NOSE LANDING GEAR MAINTENANCE PRACTICES

32-20-00 200200

1. PROCEDURES WARNING: When performing maintenance on a hydraulically operated landing gear system, be aware that any movement of a hydraulic actuator cylinder can displace hydraulic fluid and cause unanticipated movement of other actuator cylinders in the landing gear retraction system. Servicing of the landing gear hydraulic accumulator can also result in unanticipated movement of an actuator. Either action can result in an unsafe, unlocked landing gear system. Therefore, place the airplane on jacks prior to performing any inspection or maintenance. Cycle the landing gear and ensure that all three landing gears are down and locked prior to removing the airplane from jacks.

2. NOSE LANDING GEAR SHOCK ABSORBER A. Servicing (1) To check the fluid level and fill the nose landing gear shock absorber, remove the valve cap and depress the valve core to release the air. Allow the strut to fully compress. WARNING: Release the air pressure entirely before removing the valve core. (2) Remove the valve core and connect one end of a 1/4 inch I.D. hose to the valve stem. Submerge the other end of the hose in a container of clean hydraulic fluid (39, Table 1, Chapter 91-00-00). (3) Slowly extend the strut to draw hydraulic fluid into the strut. (4) When the strut is fully extended, slowly compress the strut to expel any excess hydraulic fluid and trapped air. (5) Cycle the strut as many times necessary to expel all of the air. With the strut compressed, remove the hose and install the valve core. CAUTION: Damage to the torque knees may result if the struts are inflated while the airplane is on jacks. Never tow or taxi with a flat strut. Even brief towing or taxiing with a deflated strut can cause severe damage.

B. Servicing - Inflation WARNING: Anytime maintenance is to be performed on the landing gear system, always place the airplane on jacks. When jacking the airplane in an unsheltered area where winds in an excess of 35 kts may be encountered, never jack more than one gear clear of the ground at a time. Any time the landing gear is only partially retracted during maintenance, always cycle the gear with the power pack through at least one complete cycle before removing the airplane from the jacks.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL CAUTION: Loss of oil will result in struts appearing to need inflation. Low oil levels will damage the strut and could result in hard landing damage to the aircraft. If loss of oil is detected or suspected, check oil level using the NOSE LANDING GEAR SHOCK ABSORBER SERVICING procedure.

C. Method 1 (1) With the airplane on the ground and empty, except for full fuel and oil, inflate the nose strut with nitrogen or dry filtered air until the piston is extended 5.25 to 5.75 inches.

D. Method 2 (1) With the airplane on jacks, inflate the nose landing gear with nitrogen or dry filtered air to 95 psig.

3. NOSE LANDING GEAR SHIMMY DAMPER A. Fluid Check (Manual Steering Only) NOTE: The shimmy damper fluid level can be checked without removing the damper from the airplane by inserting a thin wire through the aft end of the piston rod assembly and into the hole of the aft floating piston. To determine if the wire is inserted in the hole of the floating piston, insert the wire several times noting each insertion depth. When the wire is in the floating piston hole, the depth will be approximately 0.44 inch greater than when against the face of the floating piston. (1) Check the shimmy damper hydraulic fluid level as follows (Ref. Figure 201): (a) Insert a 1/16 inch diameter wire into the hole of the aft floating piston (2), mark the wire at the piston rod assembly - aft end (3), remove and measure the distance from inserted end of the wire (Dimension A). (2) If dimension A is greater than six inches, service the shimmy damper. Refer to Chapter 32 of the Model 1900 Airliner Series Component Maintenance Manual for servicing instructions.

B. Removal (Manual Steering Only) (1) Remove nut (1), washer (2) and bolt (3) attaching the shimmy damper (4) to steering bell crank (5) (Ref. Figure 202). (2) Remove nut (6) washer (7), bolt (9) and bushing (8) connecting the shimmy damper (4) to the airplane structure. (3) Remove the shimmy damper (4) from the airplane.

C. Installation (Manual Steering Only) (1) Position the nose landing gear (13) so the steering stop assembly (14) is centered between the steering stop ears (15) on the aft side of the nose landing gear (Ref. Figure 202). (2) Prior to installing the nose landing gear shimmy damper (1), adjust the aft end of piston rod assembly (3) to measure 3.75 inches from the aft end of barrel (4) (Ref. Figure 201). (3) Position the aft end of shimmy damper (10) in the airplane structure (Ref. Figure 202).

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(4) Insert bushing (8) and secure with bolt (9), washer (7) and nut (6). (5) With the nose wheels centered, adjust rod end (11) until the bolt (3) will fit through the rod end and steering bellcrank (5) without binding. Install washer (2) and nut (1). (6) Tighten jam nut (12) on rod end (11).

DIMENSION "A" 1

4

3

2 FWD 3.75 INCHES

1. SHIMMY DAMPER 2. AFT FLOATING PISTON 3. PISTON ROD ASSEMBLY-AFT END 4. BARREL-AFT END UC32B 074213AA.AI

Figure 201 Nose Landing Gear Shimmy Damper Manual Steering Only

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9 8 12

10

3

4 7

11 6

A

5

2 1

1. NUT 2. WASHER 3. BOLT 4.SHIMMY DAMPER 5. BELLCRANK 6. NUT 7. WASHER 8. BUSHING 9. BOLT 10. SHIMMY DAMPER AFT END 11. ROD END 12. JAM NUT

B

DETAIL

A

UC32B 070471AA.AI

Figure 202 (Sheet 1 of 2) Shimmy Damper Installation

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13 13. NOSE LANDING GEAR ASSEMBLY 14. STEERING STOP ASSEMBLY 15. STEERING STOP EARS

15

15 14

DETAIL

B

UC32B 070472AA.AI

Figure 202 (Sheet 2 of 2) Shimmy Damper Installation

4. NOSE LANDING GEAR A. Lubrication Lubricate the nose landing gear wheel bearings and grease fittings (Ref. Chapter 12-20-00, Table 202, LUBRICATION SCHEDULE).

B. Removal (1) Perform the THREE-POINT JACKING (PREFERRED (Ref. Chapter 07-10-00). All tires must be clear of the floor.

PROCEDURE)

procedure

(2) Retract the nose landing gear slightly to take the load off the drag brace and to unlock the actuator. (3) Disconnect the push-pull links, which operate the nose landing gear door from the strut by removing the nuts, washers, and bolts that attach the rod ends to the strut. (4) Remove the bolt, nut and washer to disconnect the actuator clevis from the yoke. (5) Disconnect the electrical wiring for the taxi light. CAUTION: As the hydraulic hoses are disconnected, plug all openings to prevent entry of foreign material into the lines or hoses. (6) If power steering is installed on the airplane, disconnect the electrical wiring from the power steering actuator and disconnect the flexible hydraulic hoses from the bracket mounted on the LH keel of the nose gear wheel well.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (7) Remove the cotter pin, nut (13) and washer (11) from lower bolt (9) and drive the bolt out. Note the location of the washers (10) and bushings (12) for proper installation (Ref. Figure 203). (8) Remove the nut, washers, bushings and bolt securing the upper and lower drag legs together. (9) Place a block under the wheel to remove part of the load on the landing gear trunnion bolts and to support the strut when the bolts are removed. NOTE: Note the position of the washers to facilitate installation. (10) Remove the cotter pins, nuts (8), washers (2, 3, 4, 6 and 7), bushings (5) and nose landing gear trunnion bolts (1) and lower the strut away from the airplane.

C. Installation (1) Lubricate bolts (1 and 9) and bushings (5 and 12) with lubricating grease (61, Table 1, Chapter 91-00-00) (Ref. Figure 203). (2) Position the strut in the wheel well and install the landing gear trunnion bolts (1) with two 100951-X-031-UC washers (3 and 4) between the strut and keel. Install washer (2) on the bolt with the countersink towards the bolt heads. Install bushings (5) over the bolt and into the strut casting after the strut is in position. Install nut (8) on the bolt and torque the trunnion bolts 400 to 700 in.-lbs and secure with cotter pins. Use an AN960-1016 and/or AN960-1016L washer (7) under the nut as required to obtain proper torque and engagement of the nut and cotter pin. NOTE: Position the washers in the same sequence as when previously installed. (3) Place the lower ends of the lower drag legs through the strut web and into position to be attached. Attach the lower drag legs to the upper drag leg with the bolt, bushings, washers and nut. Torque the attaching nut 200 to 400 in.-lbs and install the cotter pin. Use AN960-1016 or AN960-1016L washers under the bolt head and nut to obtain the proper torque and engagement of the nut and cotter pin. (4) Secure the lower drag legs to the strut with bolt (9), being careful to install bushings (12) and washers (10 and 11) in the same positions from which they were removed. Torque the attaching nut (13) 200 to 400 inch-pounds and install the cotter pin. Use an AN960-1016 and/or AN960-1016L washer (11) under the nut as required to obtain the proper torque and engagement of the nut and cotter pin. (5) Extend the actuator and align the actuator clevis bolt hole with the bolt in the yoke, then secure with the bolt, washer, nut, and cotter pin. (6) Connect the electrical wiring to the taxi light. (7) Attach the push-pull links, that operate the nose gear door, to the strut by installing the nuts, washers, and bolts that attach the rod ends to the strut. (8) Lubricate the strut and drag brace hinge joints with lubricating grease (61, Table 1, Chapter 91-00-00). (9) Inflate the nose landing gear strut to 30 psi minimum with nitrogen or dry filtered air. (10) Perform a retraction check (Ref. 32-30-00, LANDING GEAR HAND PUMP CYCLING). (11) Install the nose landing gear trunnion bolt access covers. Page 206 Nov 1/13

32-20-00

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (12) Perform the LOWERING THE AIRPLANE AFTER THREE-POINT JACKING procedure (Ref. Chapter 07-10-00). (13) Perform the NOSE LANDING GEAR SHOCK ABSORBER SERVICING - INFLATION procedure in this section.

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Figure 203 Nose Landing Gear Assembly

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5. NOSE LANDING GEAR TORQUE KNEE A. Removal (1) Perform the THREE-POINT JACKING (PREFERRED (Ref. Chapter 07-10-00). All tires must be clear of the floor.

PROCEDURE)

procedure

(2) Deflate the nose landing gear strut by removing the valve cap and depressing the valve core to release the air. Allow the strut to fully compress. (3) Remove the cotter pin, nut, washers (1 and 2), bushing (3) and bolt (4) that connect upper torque knee (5) and lower torque knee (10). Note the position of the washers on the attach bolt to facilitate installation (Ref. Figure 204). (4) Remove the cotter pin, washer and tow limit pin (6) from upper torque knee pin (7). Tap the upper torque knee pin out and remove torque knee (5). (5) Remove the cotter pin, washer and clevis pin (8) from the lower torque knee pin (9). Tap the lower torque knee pin out and remove the torque knee (10). (6) Inspect the upper and lower torque knee bushings for wear and/or damage (Ref. Model 1900 Airliner Series Component Maintenance Manual, Chapter 32-20-00, NOSE LANDING GEAR DISASSEMBLY Chart 1).

B. Installation (1) Position lower torque knee (10) in place and tap the lower torque knee pin (9) in place. Install clevis pin (8), washer and cotter pin (Ref. Figure 204). (2) Position upper torque knee (5) in place and tap the upper torque knee pin (7) in place. Install tow limit pin (6), washer and cotter pin. (3) Install the bolt (4), washers (1 and 2), nut, and cotter pin that attach upper torque knee (5) to lower torque knee (10). Install washers (1 and 2) in the same positions as when removed. (4) Lubricate the torque knee joints with lubricating grease (61, Table 1, Chapter 91-00-00). (5) Inflate the nose landing gear to 30 psi minimum with nitrogen or dry filtered air. (6) Perform the LOWERING THE AIRPLANE AFTER THREE-POINT JACKING procedure (Ref. Chapter 07-10-00). (7) Perform the NOSE LANDING GEAR SHOCK ABSORBER SERVICING - INFLATION procedure in this section.

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Figure 204 Nose Landing Gear Torque Knee Assembly

6. NOSE LANDING GEAR DRAG BRACE WARNING: When performing maintenance on a hydraulically operated landing gear system, be aware that any movement of a hydraulic actuator cylinder can displace hydraulic fluid and cause unanticipated movement of other actuator cylinders in the landing gear retraction system. Servicing of the landing gear hydraulic accumulator can also result in unanticipated movement of an actuator. Either action can result in an unsafe, unlocked landing gear system. Therefore, place the airplane on jacks prior to performing any inspection or maintenance. Cycle the landing gear and ensure that all three landing gears are down and locked prior to removing the aircraft from jacks.

A. Removal WARNING: Any time maintenance is to be performed on the landing gear system, place the airplane on jacks. When jacking the airplane in an unsheltered area where winds in excess of 35 kts will be encountered, never jack more than one gear at a time clear of the ground.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Any time the landing gear is only partially retracted during maintenance, always cycle the gear with the power pack through at least one complete cycle before removing the airplane from the jacks. For safety reasons, pull the 2-ampere control circuit breaker on the pilot's inboard subpanel and place a note on the circuit breaker panel that LANDING GEAR MAINTENANCE IS IN PROCESS during maintenance on the landing gear. The landing gear control handle must never be moved from the down-and-locked position while the airplane is on the ground. It is recommended that the area be roped off during extension or retraction of the landing gear. CAUTION: The landing gear must not be cycled with the power pack if low on fluid or if the landing gear system is not properly rigged. Use the emergency extension hand pump, TK229/939 hydraulic hand pump or TK229-1/939 air-driven hydraulic pump to extend and retract the landing gear for maintenance and rigging (Ref. Chapter 91-00-00). (1) Perform the THREE-POINT JACKING (PREFERRED (Ref. Chapter 07-10-00). All tires must be clear of the floor.

PROCEDURE)

procedure

(2) Perform the REMOVING GROUND POWER procedure (Ref. Chapter 24-40-00) and disconnect the battery. (3) Retract the nose gear until the load is removed from the drag brace and the actuator (3) is unlocked (Ref. Figure 205). NOTE: Note the position and thickness of all washers for installation purposes. (4) Remove the nuts, washers and bolts attaching the push-pull link rod ends to the nose gear door bracket. Tie the nose gear door away from the work area. (5) Remove the clamps securing the down-position switch wiring to the upper drag brace (21). NOTE: To maintain down-position switch (33) adjustment, do not allow the down-position switch nut on the top of the upper drag brace (21) to turn. (6) Remove the down-position switch (33) from the upper drag brace (21). (7) Remove cotter pin, nut, washer, bushing and bolt (34) attaching the yoke to the upper drag brace (21). (8) Remove the nut (5), washer (6) and bolt (7) attaching the actuator clevis to the yoke (8). NOTE: Note the location and thickness of the washers and bushings for installation purposes. (9) Remove the cotter pin (9), nut (10), washers (11 and 12), bushings (13) and bolt (14) attaching the lower drag brace legs (1 and 2) to the strut (4). (10) If necessary, retract the actuator to avoid interference when removing the lower drag brace legs (1 and 2). (11) Remove the cotter pin (22), nut (23), washers (24 and 25), bushings (26) and bolt (27) connecting the upper drag brace (21) and lower drag brace legs (1 and 2).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (12) Remove the two nuts (28), washers (29) and bolts (30) securing the two pins (31) in the upper drag brace (21). (13) Remove the two pins (31) and washers (32) from the upper drag brace using vise-grips or equivalent pliers. Retain the washers (32) located between the upper drag brace and the attachment fitting. Remove the upper drag brace assembly.

B. Installation (1) Coat the pins (31) and washers (32) with grease (61, Table 1, Chapter 91-00-00) (Ref. Figure 205). (2) Install one AN960-1016L washer (32) on each side of the upper drag brace (21) between the upper drag brace and attachment beam. (3) Tap the two pins (31) into the attachment beam and the upper drag brace (21). Align the holes in the pin with the drag brace bolt holes. NOTE: Ensure the end of each pin is 0.25 inch from the outboard side of the attachment beam (Ref. Figure 205, Sheet 2 of 2). (4) Install the two bolts (30), through the upper drag brace (21) and the pins. Install the washers (29) and nuts (28). (5) Coat the hinge bolt (27) and bushings (26) with grease (61, Table 1, Chapter 91-00-00). (6) Align the upper drag brace (21) with the lower drag brace legs (1 and 2). (7) Install the bushings (26) in the same LH or RH lower drag brace legs (1 or 2) as noted during removal. (8) Connect the lower drag brace legs (1 and 2) to the upper drag brace (21) with the attaching bolt (27), washers (24 and 25) and nut (23). NOTE: Use an AN960-1016 and/or AN960-1016L washers (24) under the nut as required to obtain proper torque. (9) Torque the nut (23) 200 to 400 inch-pounds and install a new cotter pin (22). (10) Coat the bolt (14) and bushings (13) with grease (61, Table 1, Chapter 91-00-00). (11) Place the lower ends of the lower drag brace legs (1 and 2) through the web in the strut (4). (12) Attach the lower drag brace legs to the nose gear strut with the bolt (14), washers (11 and 12) and nut (10). NOTE: Use AN960-1016 or AN960-1016L washers (11) under the bolt head (14) and nut (10) to obtain proper torque. (13) Torque the nut (10) 200 to 400 inch-pounds and install a new cotter pin (9). (14) Align the actuator (3) until the actuator clevis bolt hole aligns with the bolt hole in the yoke (8) and install the attaching bolt (7), washer (6), nut (5).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL NOTE: To maintain adjustment of the down-position switch (33), do not allow the nut on the top of the upper drag brace (21) to turn. A minimum of one AN960-616 or AN960-616L washer must be installed on each side of the yoke. (15) Manually pull the drag brace to the extended position and extend the actuator. Install the AN960-616 and/or AN960-616L washers, as required on each side of the yoke to obtain correct alignment (Ref. 32-30-13, NOSE LANDING GEAR RIGGING). CAUTION: Ensure the yoke bolt head is facing the shimmy damper. Failure to do so will cause damage to the shimmy damper upon retraction. (16) Install the stop bolt (34), bushing, washer, nut and cotter pin. (17) Install the down-position switch (33) and wiring clamps to the upper drag brace (21). (18) Connect the nose gear door push-pull links to the nose gear door with the attaching nuts, washers and bolts. Refer to the 32-30-15, NOSE GEAR DOOR ASSEMBLY illustration in the Maintenance Practices section. (19) Lubricate the drag brace grease fittings with grease (61, Table 1, Chapter 91-00-00). Use an Alemite Z-737 nozzle to lubricate the drag brace hinge bolt and pins (14, 15, 27 and 31). (20) Check the landing gear rigging (Ref. Chapter 32-30-09, MAIN LANDING GEAR RIGGING). (21) Cycle the landing gear and check for proper operation of the in-transit and gear-down lights. If necessary, adjust the down-position, up-position, and actuator down-position switches (Ref. Chapter 32-60-00).

C. Bolt Inspection (1) Perform the THREE-POINT JACKING (PREFERRED (Ref. Chapter 07-10-00). All tires must be clear of the floor.

PROCEDURE)

procedure

(2) Perform the REMOVING GROUND POWER procedures (Ref. Chapter 24-40-00) and disconnect the battery. (3) Pull up the red knob on top of the service valve. Using the alternate extension hand pump, slightly retract the nose landing gear until the nose gear actuator is unlocked and the load is removed from the drag brace. Refer to the LANDING GEAR HAND PUMP CYCLING procedure in Chapter 32-30-00. (4) Lubricate the drag brace grease fittings with grease (61, Table 1, Chapter 91-00-00). (5) Remove the cotter pin (9) and loosen the nut (10) two turns on the lower bolt (14) attaching the lower drag brace legs (1 and 2) to the strut (4) (Ref. Figure 201). (6) Remove the cotter pin (22) and loosen the nut (23) two turns on the center bolt (27) attaching the lower drag brace legs (1 and 2) to the upper drag brace (21). (7) With a torque wrench on the bolt head, check the torque needed to turn the lower bolt (14). Repeat for the center bolt (27). The torque required to turn either bolt is 400 inch-pounds maximum.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (8) If more than 400 inch-pounds torque is needed to turn the bolts, determine the cause and replace parts as necessary. Install parts removed and repeat Steps (4) thru (7). (9) When the bolt torques meet requirements, tighten nut (23) on the center bolt (27) and torque 200 to 400 inch-pounds. Install a new cotter pin (22). (10) Tighten nut (10) on the lower bolt (14) and torque 200 to 400 inch-pounds. Install a new cotter pin (9). (11) Disconnect the nose landing gear actuator clevis from the yoke (8) by removing the nut (5), washer (6) and clevis bolt (7). (12) Retract the actuator or move the drag brace so the yoke (8) can be rotated. Check for free rotation of the yoke around the pivot/stop bolt (34). (13) If the yoke does not rotate freely on the pivot/stop bolt (34), determine the cause and replace parts as necessary. The yoke is not symmetrical, so if removal of the yoke is necessary, note the position of the short side and the thickness and number of the washers on each side of the yoke. Install parts removed and repeat Step (12). (14) If the yoke was removed, position the short side and install the washers as noted at removal. If necessary, check alignment of the yoke with the drag brace (Ref. 32-30-13). (15) Connect the nose landing gear actuator clevis to the yoke (8) by installing the clevis bolt (7), washer (6), and nut (5). Tighten the nut to a recommended torque of 56 to 78 inch-pounds but do not exceed the maximum torque of 135 inch-pounds. (16) Push down the service valve red knob. Place the cockpit landing gear control handle in the DN position. (17) Use the alternate extension hand pump to move the landing gear to the down and locked position. (18) Perform the APPLYING GROUND POWER procedures (Ref. Chapter 24-40-00). (19) Cycle the landing gear with the power pack through at least one cycle and ensure that all three landing gears are down-and-locked. (20) Perform the REMOVING GROUND POWER procedures (Ref. Chapter 24-40-00). Connect the battery. (21) Perform the LOWERING THE AIRPLANE AFTER THREE-POINT JACKING procedures (Ref. Chapter 07-10-00).

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Figure 205 (Sheet 1 of 2) Nose Landing Gear Drag Brace

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Figure 205 (Sheet 2 of 2) Nose Landing Gear Drag Brace

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LANDING GEAR LANDING GEAR EXTENSION AND RETRACTION DESCRIPTION AND OPERATION

32-30-00 00

1. GENERAL WARNING: When an airplane has experienced abnormal landing gear procedures of any type, as a safety precaution, place the airplane on jacks prior to performing any inspection or maintenance. Ensure that all three landing gears are down and locked prior to removing the airplane from jacks. When performing maintenance on a hydraulically operated landing gear system, be aware that any movement of a hydraulic actuator cylinder can displace hydraulic fluid and cause unanticipated movement of other actuator cylinders in the landing gear retraction system. Servicing of the landing gear hydraulic accumulator can also result in unanticipated movement of an actuator. Either action can result in an unsafe, unlocked landing gear system. Therefore, place the airplane on jacks prior to performing any inspection or maintenance. Cycle the landing gear and ensure that all three landing gears are down and locked prior to removing the airplane from jacks. Any time maintenance is to be performed on the landing gear system, place the airplane on jacks. When jacking the airplane in an unsheltered area where winds in excess of 35 kts will be encountered, never jack more than one gear at a time clear of the ground. Any time the landing gear is only partially retracted during maintenance, always cycle the gear with the power pack through at least one complete cycle before removing the airplane from the jacks. For safety reasons, pull the 2-ampere control circuit breaker on the pilot's inboard subpanel and place a note on the circuit breaker panel that LANDING GEAR MAINTENANCE IS IN PROCESS during maintenance on the landing gear. The landing gear control handle must never be moved from the down-and-locked position while the airplane is on the ground. It is recommended that the area be roped off during extension or retraction of the landing gear. CAUTION: Jacking of an airplane for the purpose of landing gear operation, inspection, servicing or maintenance, should be accomplished within an enclosed building or hangar. In the interest of safety, should it become necessary to jack the airplane in the open, wind velocity in any direction and terrain variations, must be compensated for prior to jacking the airplane. The landing gear must not be cycled with the power pack if low on fluid or if the landing gear system is not properly rigged. Use the emergency extension hand pump, TK229/939 hydraulic hand pump or TK229-1/939 air-driven hydraulic pump to extend and retract the landing gear for maintenance and rigging (Ref. Chapter 91-00-00).

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The nose and main landing gear assemblies are extended and retracted by a hydraulic power pack, located inboard of the LH nacelle and forward of the main spar. The hydraulic power pack consists primarily of a 28 vdc motor, pump, two-section reservoir, filter, four-way selector valve, up-and-down selector solenoids, gear-up pressure switch and low fluid level sensor (Ref. Figure 1). To prevent cavitation of the pump, engine bleed air, regulated to 18 to 20 psi, is plumbed into the power pack reservoir and the system fill can. A capped tee, adjacent to the manual bleed valve, is plumbed into the bleed air line as a connection point from a source of pressurized air when the engines are not operating. Associated plumbing for a normal extend mode, normal retract mode, and emergency extend mode is routed from the power pack to each main landing gear actuator and the nose landing gear actuator (Ref. Figures 2, 3 and 4). The plumbing for the normal extend mode and the emergency extend mode is fitted as separate plumbing to the shuttle valve in each actuator. The shuttle valves are spring-loaded to a position which allows hydraulic fluid from the normal extend plumbing to flow into and out of the actuator cylinder. The retract mode plumbing is fitted to the opposite end of each actuator. A landing gear control switch handle, placarded UP DN is located on the pilot's inboard subpanel (Ref. Figure 5). A solenoid operated downlock latch prevents the landing gear control handle from being raised while the airplane is on the ground. The landing gear safety switch releases the latch when the plane leaves the ground. If necessary, the latch can be manually overridden by pressing down on the red button placarded DN LOCK REL. To prevent accidental gear retraction on the ground, a safety switch on the right main strut breaks the control circuit whenever the strut is compressed. CAUTION: Never rely on the safety switch to keep the gear down while taxiing or on landing or take-off roll. Always check the position of the landing gear switch. The landing gear control circuit is protected by a two ampere circuit breaker located on the pilot's inboard subpanel. Power for the pump motor is supplied through the landing gear motor relay and the 200 ampere relay limiter, both of which are located in the LH nacelle power distribution panel. The motor relay is energized by current from the time delay relay Printed Circuit Board (PCB) located under the center aisle floorboard, just aft of the main spar. The time delay relay will interrupt current to the motor and trip the landing gear control circuit breaker if continuous motor operation exceeds the maximum time limit. A circuit to the time delay relay from the landing gear control switch is completed through the RH safety switch and the gear-up pressure switch when the control handle is placed in the UP position and through the actuator downlock switches when the control handle is placed in the DN position. When the landing gear control handle is moved to the UP position, the gear-up solenoid, mounted on the valve body end of the power pack, is energized to actuate the gear selector valve to allow system fluid under pressure from the pump to flow to the retract side of the system. On airplanes equipped with power steering, the nose landing gear position switch, mounted on the nose landing gear straightener link, will prevent retraction of the landing gear until the nose wheel is centered. The landing gear will begin to retract when 200 to 400 psi of hydraulic pressure at the retract port of the actuator unlocks the internal mechanical locking mechanism. As the actuator pistons move to retract the landing gear, the fluid in the actuator exits through the normal extend port of the actuators and is carried back to the power pack through the normal extend plumbing. When the hydraulic fluid enters the power pack, the gear selector valve directs the return fluid to the primary reservoir. The landing gear is held in the retracted position by positive hydraulic pressure. When the system pressure reaches the high pressure limit, the gear-up pressure switch, mounted on the power pack assembly, will interrupt current to the pump motor. This same pressure switch will activate the pump motor should the system pressure drop to the low pressure limit. An accumulator, precharged to 800 ± 50 psi, is located in the LH wheel well and is designed to aid in maintaining the system pressure in the gear-up mode. When the landing gear control handle is moved to the DN position, the gear-down solenoid is energized to actuate the selector valve to allow the system fluid under pressure from the pump to flow to the extend side of the system. As the actuator pistons move to extend the landing gear, the fluid in the actuators exits through

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL the normal retract port of the actuators and is carried back to the power pack through the normal retract plumbing. Fluid from the pump opens a pressure check valve in the power pack to allow the return fluid to flow into the primary reservoir. When the actuator pistons are positioned to fully extend the landing gear, an internal mechanical lock in each actuator will lock the actuator pistons to hold the gear in the down position. In this position, the internal locking mechanism in each actuator will actuate the actuator downlock switch to interrupt current to the pump motor. The motor will continue to run until all three landing gears are down and locked. A yellow HYD FLUID LOW annunciator located in the CAUTION/ADVISORY panel will illuminate in the event the hydraulic fluid level in the landing gear power pack becomes critically low. When low fluid level is indicated, the landing gear should not be extended or retracted using the hydraulic power pack; however, the landing gear can be extended using the emergency extension hand pump. A sensing unit mounted on the motor end of the power pack provides the necessary switching circuitry to illuminate the low fluid light. The optically operated sensing unit has an integrated self-test circuit. The integral self-test circuit is energized by a switch on the instrument panel and functionally tests the sensing unit's internal circuitry. Manual landing gear extension is provided through a manually powered hydraulic system. A hand pump, placarded LANDING GEAR ALTERNATE EXTENSION, is located on the floor between the pilot's seat and the pedestal. The pump is used when emergency extension of the gear is required. To extend the gear with this system, pull the landing gear control circuit breaker on the pilot's inboard subpanel and place the landing gear control handle in the DN position. Remove the pump handle from the securing clip and pump the handle up and down to extend the gear. As the handle is pumped, hydraulic fluid is drawn from the hand pump suction port of the power pack into the pump and exited under pressure. Fluid under pressure from the pump is routed to the power pack hand pump pressure port and to the shuttle valve in each actuator. Fluid pressure at the shuttle valves will position the valves to allow fluid to flow into the actuator cylinders. As the actuator pistons move to extend the landing gear, the fluid in the actuators exits through the normal retract port of the actuators and is carried back to the power pack through the normal retract plumbing. The fluid routed to the power pack hand pump pressure port from the hand pump unseats the internal hand pump dump valve to allow the return fluid to flow into the primary reservoir. Continue to pump the handle up and down until the green GEAR DOWN indicator lights on the pilot's inboard subpanel illuminate. Ensure that the pump handle is in the full down position prior to placing the pump handle in the securing clip. When the pump handle is stowed, an internal relief valve is actuated to relieve the hydraulic pressure in the pump. WARNING: After an EMERGENCY landing gear extension has been made, do not move any landing gear controls or reset any switches or circuit breakers until the cause of the malfunction has been determined and corrected. A service valve located inboard of the pump and motor assembly may be used, in conjunction with the hand pump, to raise and lower the gear for maintenance purposes. The service valve is accessible through a door in the wing skin just inboard of the LH nacelle. With the airplane on jacks and an external power source attached, unlatch the hinged retainer and pull up on the red knob located on top of the service valve. The hand pump can then be pumped to raise the gear to the desired position. After the required maintenance has been performed, push the red knob down and use the hand pump to lower the gear. CAUTION: If the red knob on the service valve is pushed down while the landing gear is retracted and the electrical power is on and the landing gear control handle is in the down position, the landing gear will extend immediately. A fill can, located just inboard of the LH nacelle and forward of the main spar, contains a cap and dipstick assembly, marked FILL WARM - COLD, for convenience of maintaining system fluid level. Prior to removing the fill can lid, the knob on the manual bleed valve must be depressed to relieve the air pressure in the system.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

LINE CODE TO OVERBOARD VENT

LIGHTS BEACON STROBE LDG GEAR CONTROL

NOSE LANDING GEAR ACTUATOR

REGULATED ENGINE BLEED AIR LINE

TO CROSS TO NOSE, LH AND RH MAIN LANDING GEAR ACTUATORS

NORMAL EXTEND LINE

UP GEAR DOWN

DN DN LOCK REL

NORMAL RETRACT LINE

TAIL FLOOD

NOSE

HDL LT TEST

L

HAND PUMP SUCTION LINE

R

WARN HORN

SERVICE VALVE

CHECK VALVE

OFF LANDING GEAR

HAND PUMP PRESSURE LINE CHECK VALVE

FILTER

EMERGENCY EXTEND LINE

2A SILENCE

RELAY

FROM ENGINE BLEED AIR MANIFOLD

PILOTS INBOARD SUBPANEL

A

DETAIL

EMERGENCY EXTENSION HAND PUMP

A

FLOW DIRECTION

GEAR UP AND DOWN SOLENOIDS VENT PORT

FROM HAND PUMP

GEAR UP PRESSURE SWITCH

GEAR DOWN PORT HAND PUMP PRESSURE PORT

LH NACELLE ELECTRICAL DISTRIBUTION PANEL ACCUMULATOR

LANDING GEAR MOTOR RELAY

LH MAIN LANDING GEAR ACTUATOR

B

AIR REGULATOR BLEED AIR FROM ENGINES

FILL CAN

SERVICE VALVE POWER PACK ASSEMBLY

200 AMP LIMITER

MAIN SPAR

TIME DELAY RELAY PCB

HAND PUMP SUCTION PORT

FLUID LEVEL SENSOR

TO LH MAIN LANDING GEAR ACTUATOR MANUAL BLEED VALVE

RH MAIN LANDING GEAR ACTUATOR

MOTOR

TO NOSE LANDING GEAR ACTUATOR GEAR UP PORT TO TEE TO NOSE AND LH MAIN LANDING GEAR ACTUATORS

TO RH MAIN LANDING GEAR ACTUATOR

TO OVERBOARD VENT

FILL CAN

TO RH MAIN LANDING GEAR ACTUATOR

TO OVERBOARD VENT DETAIL

B

FILL PORT

TO HAND PUMP

SEAL DRAIN PORT UC32B 063184AA.AI

Figure 1 Hydraulic Landing Gear System

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Figure 2 Hydraulic Landing Gear Schematic (Normal Extend Mode)

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Figure 3 Hydraulic Landing Gear Schematic (Normal Retract Mode)

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Figure 4 Hydraulic Landing Gear Schematic (Emergency Extend Mode)

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Figure 5 Hydraulic Landing Gear Schematic (Hand Pump Retract Mode)

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LANDING GEAR EXTENSION AND RETRACTION TROUBLESHOOTING

100100

WARNING: When an airplane has experienced abnormal landing gear procedures of any type, as a safety precaution, place the airplane on jacks prior to performing any inspection or maintenance. Ensure that all three landing gears are down and locked prior to removing the airplane from jacks. CAUTION: Jacking of an airplane for the purpose of landing gear operation, inspection, servicing or maintenance, should be accomplished within an enclosed building or hangar. In the interest of safety, should it become necessary to jack the airplane in the open, wind velocity in any direction and terrain variations, must be compensated for prior to jacking the airplane. Refer to Charts in Figures 102, 103 and 104 for data on troubleshooting the landing gear system.

1. TIME DELAY RELAY PRINTED CIRCUIT BOARD (PCB) A. Functional Check In the normally operating landing gear control circuit, the time delay assembly will apply ground, through a current limiting resistor, to the landing gear control power circuit in 16 ± 0.5 seconds, causing the two ampere gear control circuit breaker to open, preventing further damage to the system. Should the time delay circuit for the landing gear malfunction for any reason and should the uplock pressure switch fail to interrupt power to the landing gear power pack motor, the motor would continue to run and serious damage to the power pack assembly could result. A test unit has been designed for use in functionally checking the time delay assembly for proper operation. The following procedure outlines the method of checking the operation of the time delay assembly. (1) Assemble the test unit (Ref. Figure 101). (2) Remove the time delay PCB from the airplane. (3) Connect the test unit to the time delay assembly. (4) Apply power to the time delay assembly. Lamps L1 and L3 will illuminate. (5) After a time delay of 16 ± 0.5 seconds, lamps L1 and L3 should extinguish and lamps L2 and L4 will illuminate. If not, the time delay assembly must be replaced. (6) Remove the test unit and install the time delay PCB in the airplane.

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Figure 101 Time Delay Relay Functional Check

2. HYDRAULIC POWER PACK ASSEMBLY The following test equipment will permit troubleshooting the hydraulic power pack assembly:

A. Test (1) Hand pump capable of delivering 3,600 psig. (2) Pressure gage with a range of 4,000 psig. (3) Continuity checker. The hand pump should have a reservoir, or be equipped with fittings which will permit connection to a reservoir. The test equipment needs valves capable of controlling the application and release of hydraulic fluid to the unit under test. The test equipment should be capable of holding a pressure of at least 3,600 psig without leaks.

3. UP-PRESSURE SWITCH A. Test (1) Connect the hand pump and gage to the gear-up port of the power pack. NOTE: Connect the hand pump and gage to the gear-up port of the power pack only. Do not connect the hand pump at any other location in the hydraulic system. (2) Connect the continuity checker to the pins B and C of the electrical connector of the pressure switch. The continuity checker should indicate that there is continuity between the pins.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (3) Slowly apply pressure with the hand pump while monitoring the continuity checker and the pressure gage. The pressure switch should open, as indicated by a loss of continuity, at a pressure of 2,720 to 2,830 psig. (4) Decrease the pressure slowly while monitoring the continuity checker and the pressure gage. The pressure switch should close, as indicated by regaining continuity, at a pressure of 2,320 to 2,530 psig. (5) Disconnect the test equipment.

4. UPLOCK CHECK VALVE AND THERMAL RELIEF VALVE A. Test (1) Connect the hand pump and the pressure gage to the gear-up port of the power pack. NOTE: Connect the hand pump and the pressure gage to the gear-up port of the power pack only, do not connect the hand pump at any other location in the hydraulic system. (2) Apply a pressure of 2,800 psig. Allow the pressure gage to stabilize. The gage should not indicate a loss of pressure due to an internal leakage. (3) Slowly increase the pressure while monitoring the pressure gage. The thermal relief valve should relieve at a pressure of 3,500 ± 50 psig. Stop pumping after the thermal relief valve has cracked, and monitor the pressure gage as the pressure decreases. The pressure indication should stabilize at a minimum of 3,150 psig, indicating the pressure at which the thermal relief valve has closed. (4) Release the pressure and disconnect the equipment.

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Figure 102 Troubleshooting - Landing Gear Extension Landing Gear Will Not Extend

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Figure 103 (Sheet 1 of 2) Troubleshooting - Landing Gear Retraction Landing Gear Will Not Retract

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

IS POWER AT PIN 4 OF A316 PCB? (UB-1 THRU UB-35 ONLY W/O KIT 114-3036-1 INSTALLED)

IS POWER AT PIN 3 OF A316 PCB?

REPAIR JUMPER WIRE BETWEEN PINS 3 AND 4 OF A316 PCB

IS SERVICE VALVE PROPERLY STOWED?

IS SAFETY SWITCH CLOSED?

REPAIR OPEN CIRCUIT BETWEEN PIN U OF J107 AND PIN A OF J206.

REPAIR OR RESET CIRCUIT BREAKER OR REPAIR OPEN CIRCUIT BETWEEN PIN 4 OF P200 AND CIRCUIT BREAKER.

ADJUST OR REPLACE SWITCH.

STOW VALVE

IS POWER AT LANDING GEAR CONTROL CIRCUIT BREAKER?

IS POWER AT PIN A OF SERVICE VALVE PLUG J206?

REPAIR OR REPLACE A316 PCB.

IS POWER AT PIN T OF RIGHT WHEEL WELL PLUG J107?

IS POWER AT PIN 4 OF LANDING GEAR HANDLE PLUG P200. (UB-1 THRU UB-35 ONLY)

IS POWER AT PIN B OF SERVICE VALVE PLUG J206? IS SERVICE VALVE SWITCH CLOSED? CHECK FOR CONTINUITY ACROSS PINS B AND C AT PLUG A188P1.

REPAIR OPEN CIRCUIT BETWEEN PIN C OF J206 AND PIN 4 OF A316 PCB (UB-1 THRU UB-35 ONLY W/O KIT 114-3036-1 INSTALLED)

IS POWER AT PIN 19 OF LANDING GEAR HANDLE PLUG P200? (UB-36 AND AFTER ONLY)

REPAIR OR REPLACE GEAR HANDLE.

REPAIR OPEN CIECUIT BETWEEN PIN 7 OF P200 AND PIN T OF J107.

IS LANDING GEAR HANDLE SWITCH CLOSED? CHECK FOR CONTINUITY ACROSS PINS 4 AND 7 OF PLUG P200.

REPAIR OR REPLACE GEAR HANDLE.

IS POWER PACK PRESSURE SWITCH CLOSED?

REPAIR OR REPLACE SWITCH OR SERVICE VALVE.

REPAIR OPEN CIRCUIT BETWEEN PIN C OF J206 AND PIN 3 OF A316 PCB.

IS POWER AT PIN B OF POWER PACK PRESSURE SWITCH PLUG P208? (UB-36 AND AFTER ONLY)

RESTORE POWER TO CIRCUIT BREAKER.

REPLACE SWITCH.

REPAIR OPEN CIRCUIT BETWEEN PIN B OF J208 AND SPLICE IN WIRE G8D22.

IS LANDING GEAR HANDLE SWITCH CLOSED? CHECK FOR CONTINUITY ACROSS PINS 18 AND 19 OF PLUG P200.

REPAIR OPEN CIRCUIT BETWEEN PIN 18 OF P200 AND PIN B OF J206.

REPAIR OPEN CIRCUIT BETWEEN PIN 19 OF P200 AND PIN C OF P208. (UB-36 AND AFTER ONLY)

REPAIR OPEN CIRCUIT BETWEEN P208 PIN B AND J206 PIN B. CHECK DIODE CR339. (UB-1 THRU UB-35 ONLY) UC32B 024203AA.AI

Figure 103 (Sheet 2 of 2) Troubleshooting - Landing Gear Retraction Landing Gear Will Not Retract

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Figure 104 Troubleshooting - Landing Gear Extension And Retraction Landing Gear Cycles Too Often (1)

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LANDING GEAR LANDING GEAR EXTENSION AND RETRACTION MAINTENANCE PRACTICES

200200

1. PROCEDURES WARNING: When an airplane has experienced abnormal landing gear procedures of any type, as a safety precaution, place the airplane on jacks prior to performing any inspection or maintenance. Ensure that all three landing gears are down and locked prior to removing the airplane from jacks. When performing maintenance on a hydraulically operated landing gear system, be aware that any movement of a hydraulic actuator cylinder can displace hydraulic fluid and cause unanticipated movement of other actuator cylinders in the landing gear retraction system. Servicing of the landing gear hydraulic accumulator can also result in unanticipated movement of an actuator. Either action can result in an unsafe, unlocked landing gear system. Therefore, place the airplane on jacks prior to performing any inspection or maintenance. Cycle the landing gear and ensure that all three landing gears are down and locked prior to removing the airplane from jacks. Any time maintenance is to be performed on the landing gear system, place the airplane on jacks. When jacking the airplane in an unsheltered area where winds in excess of 35 kts will be encountered, never jack more than one gear at a time clear of the ground. Any time the landing gear is only partially retracted during maintenance, always cycle the gear with the power pack through at least one complete cycle before removing the airplane from the jacks. For safety reasons, pull the 2-ampere control circuit breaker on the pilot's inboard subpanel and place a note on the circuit breaker panel that LANDING GEAR MAINTENANCE IS IN PROCESS during maintenance on the landing gear. The landing gear control handle must never be moved from the down-and-locked position while the airplane is on the ground. It is recommended that the area be roped off during extension or retraction of the landing gear. CAUTION: Jacking of an airplane for the purpose of landing gear operation, inspection, servicing or maintenance, should be accomplished within an enclosed building or hangar. In the interest of safety, should it become necessary to jack the airplane in the open, wind velocity in any direction and terrain variations, must be compensated for prior to jacking the airplane. The landing gear must not be cycled with the power pack if low on fluid or if the landing gear system is not properly rigged. Use the emergency extension hand pump, TK229/939 hydraulic hand pump or TK229-1/939 air-driven hydraulic pump to extend and retract the landing gear for maintenance and rigging (Ref. Chapter 91-00-00). NOTE: This hydraulic system is self bleeding. Always cycle the landing gear a minimum of five times after maintenance is performed to move fluid through the system and work out trapped air that may have been introduced.

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A. Landing Gear Component Access A hinged door for the hydraulic fill can and an access panel for the service valve are provided in the upper wing panel (Ref. Figure 201). Any maintenance on the components other than changing the position of the service valve or filling the system will require removing the upper and lower wing panel assemblies. Remove the upper and lower wing panel assemblies to gain access to the landing gear power pack and related components as follows: (1) Remove the screws securing the forward wing fillet to the fuselage and remove the fillet from the airplane. (2) Remove the leading edge. (3) Remove the remaining screws from the perimeter of the upper or lower panel. NOTE: Install the panels in the following sequence: upper and lower panel, leading edge. The upper panel has a guard attached to the bottom side of the service valve access panel to prevent installation of the panel if the red knob on the service valve is not pushed down. If any difficulty is experienced installing the upper panel, check to ensure that the knob on the service valve is pushed down and that the hinged retainer is in place.

B. Hydraulic System Servicing Servicing the hydraulic landing gear system consists of maintaining the correct fluid level and correct accumulator precharge. A fill can, located just inboard of the left nacelle and forward of the main spar, contains a cup and dipstick assembly marked FILL WARM - COLD.

C. Hydraulic System Filling and Bleeding WARNING: Always place the airplane on jacks before performing any maintenance on the landing gear hydraulic system. CAUTION: The landing gear must not be cycled with the power pack if low on fluid or if the landing gear system is not properly rigged. Use the emergency extension hand pump to extend and retract the landing gear for maintenance and rigging. Any time the fill can is opened, make sure to shut off the air pressure supply and depressurize the system by depressing the button on the manual bleed valve. (1) If less than eight ounces of fluid is being added to the hydraulic system, perform the following Steps. Otherwise, proceed Step (2). (a) Perform Steps (2) thru (7). (b) Perform the LANDING GEAR HAND PUMP CYCLING procedure twice. (c) Perform Steps (16) thru (22). (2) Perform APPLYING GROUND POWER procedure (Ref. Chapter 24-40-00). (3) Perform the THREE-POINT JACKING (PREFERRED (Ref. Chapter 07-10-00). All tires must be clear of the floor.

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PROCEDURE)

procedure

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (4) Remove the forward wing fillet, wing leading edge, and upper wing panel (Ref. Figure 201). (5) Perform the ACCUMULATOR SERVICING procedure (Ref. 32-30-01). (6) Press the button on the manual bleed valve to release any air pressure in the power pack reservoir. (7) Fill the power pack by adding hydraulic fluid (39, Table 1, Chapter 91-00-00) to the fill can up to the mark on the dipstick. NOTE: When filling the hydraulic system, the air being displaced by the hydraulic fluid will need to be relieved through the manual bleed valve. Occasionally Press the button on the manual bleed valve to relieve this air pressure and to allow the system to fill faster. (8) Connect a regulated supply of 18 psi dry air to a pressure pot filled with hydraulic fluid and connect the pressure pot to the capped tee that is located adjacent to the manual bleed valve. NOTE: The hand pump handle must be in the stowed position while bleeding air out of the system. (9) With the system pressurized by the pressure pot, crack the B nuts at the following locations to bleed air from the system: (a) All three lines at the LH main landing gear actuator. (b) All three lines at the RH main landing gear actuator. (c) All three lines at the nose landing gear actuator. (d) The line at the accumulator. (e) The gear-up line at the service valve. (10) Shut off the air pressure supply and relieve the air pressure through the manual bleed valve in the power pack reservoir. Check the fluid level and refill the reservoir as required. (11) Pressurize the system with air, unlatch the hinged retainer and pull up on the red knob on the service valve to place the valve in the gear-up position. Using the hand pump, raise the landing gear to the retracted position. NOTE: Since the normal extend ports of the main landing gear actuators will be inaccessible when the landing gear is in the retracted position, crack the extend line at the bulkhead fitting on the forward side of the main spar. (12) Bleed the normal extend lines at each landing gear actuator. The normal extend ports of the actuators can be identified by a letter “P” stamped on the actuator adjacent to the port. (13) Push down the red knob on the service valve and use the hand pump to extend the landing gear. Holding maximum pressure with the hand pump handle, check the system for leaks. (14) Perform the LANDING GEAR HAND PUMP CYCLING procedure contained in this section to complete a full cycle of the landing gear. (15) Shut off the air supply and relieve the air pressure in the power pack reservoir through the manual bleed valve. Check the fluid level and refill the reservoir as required.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL CAUTION: To prevent serious damage to the pump, never operate the power pack when the engines are not operating without supplying 18 psi of regulated dry air to the power pack reservoir. (16) With the red knob on the service valve pushed down, the hand pump handle in the stowed position and the system pressurized with air, operate the landing gear a minimum of ten complete cycles using the hydraulic power pack. While cycling the landing gear, check for any signs of hydraulic fluid leakage. Every third or fourth cycle of the landing gear, check the fluid level and refill as required until the fluid level stabilizes and the retraction time of the landing gear stabilizes at approximately five seconds. CAUTION: Do not operate the power steering pump when the engines are not operating without supplying 18 psi of regulated dry air to the power pack reservoir. (17) On airplanes equipped with power steering, operate the power steering system electrically for short periods of time and check for leaks. Shut off the air supply and relieve the air pressure in the reservoir through the manual bleed valve. Check the fluid level and refill the reservoir as required. Since the power steering system is a continuous flow system using the power pack reservoir for the fluid source, the system is self-bleeding. (18) Shut off the air supply and relieve the air pressure in the reservoir through the manual bleed valve. Disconnect the air supply line from the tee and replace the cap on the tee. (19) Install the upper wing panel, wing leading edge and forward wing fillet (Ref. Figure 201). (20) Perform REMOVING GROUND POWER procedure (Ref. Chapter 24-40-00). (21) Before removing the airplane from the jacks, make sure that the hand pump handle is in the stowed position, that the red knob on the service valve is pushed down with the hinged retainer in place, and that the landing gear is in the down and locked position. (22) Perform the LOWERING THE AIRPLANE AFTER THREE-POINT JACKING procedure (Ref. Chapter 07-10-00).

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Figure 201 Landing Gear Component Access

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D. Landing Gear Hand Pump Cycling WARNING: Any time maintenance is to be performed on the landing gear system, place the airplane on jacks. When jacking the airplane in an unsheltered area where winds in excess of 35 kts will be encountered, never jack more than one gear at a time clear of the ground. Any time the landing gear is only partially retracted during maintenance, always cycle the gear with the power pack through at least one complete cycle before removing the airplane from the jacks. For safety reasons, pull the 2 ampere control circuit breaker on the pilot's inboard subpanel and place a note on the circuit breaker panel that LANDING GEAR MAINTENANCE IS IN PROGRESS during maintenance on the landing gear. The landing gear control handle must never be moved from the down-and-locked position while the airplane is on the ground. It is recommended that the area be roped off during extension or retraction of the landing gear. A service valve in the left hand wing, adjacent to the power pack is provided for use with the emergency extension hand pump to raise and lower the landing gear. The service valve is normally in a mode to extend the landing gear (red knob pushed down). The service valve is accessible through a door in the wing skin just inboard of the LH nacelle. The red knob cannot be pulled up until the hinged retainer covering it is unlatched and moved out of the way. (1) Perform REMOVING GROUND POWER procedure (Ref. Chapter 24-40-00). (2) Perform BATTERY DISCONNECTION procedure (Ref. Chapter 24-31-00). (3) Perform the THREE-POINT JACKING (PREFERRED (Ref. Chapter 07-10-00). All tires must be clear of the floor.

PROCEDURE)

procedure

(4) Unlatch the service valve knob hinged retainer. (5) Pull up on the red knob on the service valve. This places the knob in the gear-up position. (6) Pump the emergency extension hand pump until the landing gear is fully retracted. (7) Push the red knob on the service valve down. (8) Pump the emergency extension hand pump until the landing gear is fully extended. NOTE: If no other maintenance is to be performed on the landing gear system, proceed with Step (11). (9) Perform BATTERY CONNECTION procedure (Ref. Chapter 24-31-00). (10) Perform APPLYING GROUND POWER procedure (Ref. Chapter 24-40-00). WARNING: Before removing the airplane from the jacks, make sure that the landing gear emergency extend hand pump handle is in the stowed position, the plunger on the service valve is pushed down with the hinged retainer in place, the landing gear control handle is in the DOWN position, the landing gear is down and locked and the accumulator is charged to 800 ± 50 psi.

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32-30-00

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (11) Perform the LOWERING THE AIRPLANE AFTER THREE-POINT JACKING procedure (Ref. Chapter 07-10-00).

E. Hydraulic System Leakage During all landing gear tests, monitor the hydraulic system components for external leakage. Any external leakage other than a slight wetting insufficient to form a drop through any seal except in those areas where leakage is specified shall be cause for rejection. Perform the HYDRAULIC SYSTEM FILLING AND BLEEDING procedure in this section to verify system leakage following any component changes or replacement of hydraulic lines.

F. Hydraulic Fittings Installation To reduce the likelihood of leaks in the hydraulic system, care should be taken when installing the hydraulic fittings. Anytime a fitting is loosened or removed, discard the packing and install the fitting with a new packing. To prevent damage to the packing, coat the packing and the threads of the fitting with hydraulic fluid (39, Table 1, Chapter 91-00-00) before installing the packing on the fitting. Some fittings in the landing gear hydraulic system employ both a packing and an aluminum ring. When this type of fitting is loosened or removed, both the packing and the aluminum ring must be replaced when the fitting is installed (Ref. Figure 202).

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FITTING

COAT THE RING, PACKING, AND THE MALE THREADS OF THE FITTING WITH MIL-H-5606 HYDRAULIC FLUID AND ASSEMBLE AS SHOWN. WORK THE RING INTO THE COUNTERBORE OF THE NUT AND TURN THE NUT DOWN UNTIL THE PACKING IS PUSHED FIRMLY AGAINST THE LOWER THREADED PORTION OF THE FITTING.

NUT

RING PACKING STEP 1

INSTALL THE FITTING, TURNING THE NUT WITH THE FITTING, UNTIL THE PACKING CONTACTS THE SURFACE. WITH THE FITTING IN THIS POSITION, PUT A WRENCH ON THE NUT TO PREVENT IT FROM TURNING, AND TURN THE FITTING IN 1 1/2 TURNS. POSITION THE FITTING BY TURNING IT IN NOT MORE THAN ONE ADDITIONAL TURN.

FITTING

NUT

RING PACKING

STEP 2

WHILE HOLDING THE FITTING TO PREVENT MOVEMENT, TURN THE NUT DOWN AGAINST THE SURFACE. SLIGHT EXTRUSION OF THE RING IS PERMISSIBLE.

FITTING

NUT RING PACKING

STEP 3

UC32B 063194AA.AI

Figure 202 Hydraulic Fitting Installation

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G. Hydraulic Line Filter Inspection and Cleaning (UC-143 and After and prior Airplanes with Kit No. 114-8022-1S Installed) WARNING: Any time maintenance is to be performed on the landing gear system, place the airplane on jacks. When jacking the airplane in an unsheltered area where winds in excess of 35 kts will be encountered, never jack more than one gear at a time clear of the ground. Any time the landing gear is only partially retracted during maintenance, always cycle the gear with the power pack through at least one complete cycle before removing the airplane from the jacks. For safety reasons, pull the 2-ampere control circuit breaker on the pilot's inboard subpanel and place a note on the circuit breaker panel that LANDING GEAR MAINTENANCE IS IN PROCESS during maintenance on the landing gear. The landing gear control handle must never be moved from the down-and-locked position while the airplane is on the ground. It is recommended that the area be roped off during extension or retraction of the landing gear. CAUTION: The landing gear must not be cycled with the power pack if low on fluid or if the landing gear system is not properly rigged. Use the emergency extension hand pump, TK229/939 hydraulic hand pump or TK229-1/939 air-driven hydraulic pump to extend and retract the landing gear for maintenance and rigging (Ref. Chapter 91-00-00). (1) Perform the THREE-POINT JACKING (PREFERRED (Ref. Chapter 07-10-00). All tires must be clear of the floor.

PROCEDURE)

procedure

(2) Perform the REMOVING GROUND POWER procedure (Ref. Chapter 24-40-00). (3) Perform the BATTERY DISCONNECTION procedure (Ref. Chapter 24-31-00). CAUTION: If any of the filters are missing contact Hawker Beechcraft Corporation. (4) Carefully clean the hydraulic line filters and the filter connections in the two main gear and the nose gear wheel wells (Ref. Figure 203). (a) Disconnect the hydraulic line filters from the hydraulic lines. Immediately cap the lines to avoid system contamination. If any of the filters are missing, contact Hawker Beechcraft Corporation. NOTE: It is recommended that the line filter be agitated in an ultrasonic cleaner and flushed with clean solvent. (b) Clean the line filters with solvent (2, Table 1, Chapter 91-00-00). Blow dry with dry filtered compressed air. (c) Check line filters for condition. If the filter is deformed, flattened or collapsed it must be replaced. 1 Filters longer than 2.12 inches are deforming and should be replaced. (5) Remove the caps from the lines.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (6) Install the filters and connect the lines. (7) Check operation, Perform HYDRAULIC SYSTEM FILLING AND BLEEDING procedure in this section. WARNING: Before removing the airplane from the jacks, make sure that the landing gear emergency extend hand pump handle is in the stowed position, the plunger on the service valve is pushed down with the hinged retainer in place, the landing gear control handle is in the DOWN position, the landing gear is down and locked and the accumulator is charged to 800 ± 50 psi. (8) Perform the LOWERING THE AIRPLANE AFTER THREE-POINT JACKING procedure (Ref. Chapter 07-10-00). (9) Perform BATTERY CONNECTION procedure (Ref. Chapter 24-31-00).

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Figure 203 Hydraulic Line Filters

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LANDING GEAR LANDING GEAR HYDRAULIC ACCUMULATOR MAINTENANCE PRACTICES

32-30-01 200200

1. PROCEDURES WARNING: When an airplane has experienced abnormal landing gear procedures of any type, as a safety precaution, place the airplane on jacks prior to performing any inspection or maintenance. Ensure that all three landing gears are down and locked prior to removing the airplane from jacks. When performing maintenance on a hydraulically operated landing gear system, be aware that any movement of a hydraulic actuator cylinder can displace hydraulic fluid and cause unanticipated movement of other actuator cylinders in the landing gear retraction system. Servicing of the landing gear hydraulic accumulator can also result in unanticipated movement of an actuator. Either action can result in an unsafe, unlocked landing gear system. Therefore, place the airplane on jacks prior to performing any inspection or maintenance. Cycle the landing gear and ensure that all three landing gears are down and locked prior to removing the airplane from jacks. Any time maintenance is to be performed on the landing gear system, place the airplane on jacks. When jacking the airplane in an unsheltered area where winds in excess of 35 kts will be encountered, never jack more than one gear at a time clear of the ground. Any time the landing gear is only partially retracted during maintenance, always cycle the gear with the power pack through at least one complete cycle before removing the airplane from the jacks. For safety reasons, pull the 2-ampere control circuit breaker on the pilot's inboard subpanel and place a note on the circuit breaker panel that LANDING GEAR MAINTENANCE IS IN PROCESS during maintenance on the landing gear. The landing gear control handle must never be moved from the down-and-locked position while the airplane is on the ground. It is recommended that the area be roped off during extension or retraction of the landing gear. CAUTION: Jacking of an airplane for the purpose of landing gear operation, inspection, servicing or maintenance, should be accomplished within an enclosed building or hangar. In the interest of safety, should it become necessary to jack the airplane in the open, wind velocity in any direction and terrain variations, must be compensated for prior to jacking the airplane. The landing gear must not be cycled with the power pack if low on fluid or if the landing gear system is not properly rigged. Use the emergency extension hand pump, TK229/939 hydraulic hand pump or TK229-1/939 air-driven hydraulic pump to extend and retract the landing gear for maintenance and rigging (Ref. Chapter 91-00-00). NOTE: This hydraulic system is self bleeding. Always cycle the landing gear a minimum of five times after maintenance is performed to move fluid through the system and work out trapped air that may have been introduced.

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A. Servicing (1) Perform the THREE-POINT JACKING (PREFERRED PROCEDURE) procedure (Ref. Chapter 07-10-00). All tires must be clear of the floor. Strut limiters will reduce strut extension (Ref. Chapter 91-00-00). (2) Charge the landing gear hydraulic accumulator to 800 ± 50 psi using dry compressed air or bottled nitrogen. A charging gage is mounted on the accumulator (Ref. Figure 201). (3) Perform the LOWERING THE AIRPLANE AFTER THREE-POINT JACKING procedure (Ref. Chapter 07-10-00).

B. Removal (1) Perform the THREE-POINT JACKING (PREFERRED PROCEDURE) procedure (Ref. Chapter 07-10-00). All tires must be clear of the floor. Strut limiters will reduce strut extension (Ref. Chapter 91-00-00). (2) Release all pressure in the accumulator located in the LH wheel well. (3) Disconnect the hydraulic line connected to the forward end of the accumulator. Immediately cap the open line and union. (4) Remove the bolts, washers and clamps attaching the accumulator to the accumulator support bracket. Remove the accumulator from the airplane.

C. Installation (1) Position the accumulator attaching clamps on the accumulator. (2) Install the bolts and washers attaching the clamps and the accumulator to the accumulator support bracket in the LH wheel well. (3) Remove the protective caps and connect the hydraulic line to the forward end of the accumulator. Immediately cap the open line and union. (4) Recharge the accumulator as instructed under the heading LANDING GEAR ACCUMULATOR SERVICING in this section. (5) Perform the LOWERING THE AIRPLANE AFTER THREE-POINT JACKING procedure (Ref. Chapter 07-10-00).

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32-30-01

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

1

2

1. VALVE CAP 2. SWIVEL NUT 3. VALVE BODY 4. PACKING

3

4 UC27B 083931AA.AI

Figure 201 Landing Gear Accumulator Filler Valve

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LANDING GEAR LANDING GEAR POWER PACK MAINTENANCE PRACTICES

32-30-02 200200

1. PROCEDURES WARNING: When an airplane has experienced abnormal landing gear procedures of any type, as a safety precaution, place the airplane on jacks prior to performing any inspection or maintenance. Ensure that all three landing gears are down and locked prior to removing the airplane from jacks. When performing maintenance on a hydraulically operated landing gear system, be aware that any movement of a hydraulic actuator cylinder can displace hydraulic fluid and cause unanticipated movement of other actuator cylinders in the landing gear retraction system. Servicing of the landing gear hydraulic accumulator can also result in unanticipated movement of an actuator. Either action can result in an unsafe, unlocked landing gear system. Therefore, place the airplane on jacks prior to performing any inspection or maintenance. Cycle the landing gear and ensure that all three landing gears are down and locked prior to removing the airplane from jacks. Any time maintenance is to be performed on the landing gear system, place the airplane on jacks. When jacking the airplane in an unsheltered area where winds in excess of 35 kts will be encountered, never jack more than one gear at a time clear of the ground. Any time the landing gear is only partially retracted during maintenance, always cycle the gear with the power pack through at least one complete cycle before removing the airplane from the jacks. For safety reasons, pull the 2-ampere control circuit breaker on the pilot's inboard subpanel and place a note on the circuit breaker panel that LANDING GEAR MAINTENANCE IS IN PROCESS during maintenance on the landing gear. The landing gear control handle must never be moved from the down-and-locked position while the airplane is on the ground. It is recommended that the area be roped off during extension or retraction of the landing gear. CAUTION: Jacking of an airplane for the purpose of landing gear operation, inspection, servicing or maintenance, should be accomplished within an enclosed building or hangar. In the interest of safety, should it become necessary to jack the airplane in the open, wind velocity in any direction and terrain variations, must be compensated for prior to jacking the airplane. The landing gear must not be cycled with the power pack if low on fluid or if the landing gear system is not properly rigged. Use the emergency extension hand pump, TK229/939 hydraulic hand pump or TK229-1/939 air-driven hydraulic pump to extend and retract the landing gear for maintenance and rigging (Ref. Chapter 91-00-00).

A. Removal (1) Perform the THREE-POINT JACKING (PREFERRED (Ref. Chapter 07-10-00). All tires must be clear of the floor.

PROCEDURE)

procedure

(2) Perform REMOVING GROUND POWER procedure (Ref. Chapter 24-40-00).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (3) Perform BATTERY DISCONNECTION procedure (Ref. Chapter 24-31-00). (4) Remove the forward wing fillet, wing leading edge, and the upper and lower wing panels (Ref. Figure 201). (5) Remove and identify the electrical wiring from the motor, fluid level sensor, gear up pressure switch, and the gear up and gear down solenoids (Ref. Figure 202). (6) Press down on the button on the manual bleed valve to release the air pressure in the power pack reservoir. CAUTION: As the hydraulic lines are disconnected from the power pack, plug or cap all openings to prevent entry of foreign material into the lines or power pack. (7) Remove the following: (a) The line from the fill can to the fill port of the power pack. (b) The line from the hand pump pressure port of the power pack to the tee in the nose landing gear emergency extend line. (c) The swivel fitting at the auxiliary return port of the power pack, if power steering is installed. (8) Disconnect the following from the power pack: (a) The line to the hand pump suction port of the power pack. (b) The line to the gear up port of the power pack. (c) The line to the gear down port of the power pack. (d) The two lines to the tee in the vent port of the power pack. (e) The line to the seal drain port of the power pack. (9) Remove the bonding jumper from the motor end of the power pack. (10) Supporting the power pack from underneath the wing, cut the safety wire and remove the three bolts and washers attaching each end of the power pack to the mounting brackets. Airplanes with power steering have only two bolts securing the aft end of the power pack. The power pack can now be lowered away from the airplane.

B. Installation (1) Install the power pack through the underside of the wing and position between the mounting brackets. Install one AN960-416 and one AN960-416L washer under each bolt head and install three bolts in each end of the power pack and safety wire. Airplanes with power steering have only two bolts securing the aft end of the power pack (Ref. Figure 202). (2) Secure the bonding jumper to the power pack with the bolt and washers. (3) Connect the following to the power pack: (a) The line to the hand pump suction port of the power pack.

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32-30-02

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (b) The line to the gear up port of the power pack. (c) The line to the gear down port of the power pack. (d) The two lines to the tee in the vent port of the power pack. (e) The line to the seal drain port of the power pack. (4) Install the following: (a) The line from the fill can to the fill port of the power pack. (b) The line from the hand pump pressure port of the power pack to the tee in the emergency extend line to the nose landing gear actuator. (c) The swivel fitting at the auxiliary return port of the power pack, if power steering is installed. (5) Connect the electrical wiring to the motor, fluid level sensor, gear-up pressure switch, and the gear up and down solenoids. (6) Perform the HYDRAULIC SYSTEM FILLING AND BLEEDING procedure to fill and bleed the landing gear hydraulic system (Ref. 32-30-00). (7) Install the upper and lower wing panels, wing leading edge and the forward wing fillet (Ref. Figure 201). (8) Perform the LOWERING THE AIRPLANE AFTER THREE-POINT JACKING procedure (Ref. Chapter 07-10-00). (9) Perform BATTERY CONNECTION procedure (Ref. Chapter 24-31-00). (10) Perform APPLYING GROUND POWER procedure (Ref. Chapter 24-40-00).

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Figure 201 Landing Gear Component Access

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

SWIVEL FITTING AUXILIARY RETURN PORT (POWER STEERING ONLY)

BOLT AN960-416L WASHER AN960-416 WASHER (TYPICAL 6 PLACES) MOUNTING BRACKET

HAND PUMP PRESSURE PORT

GEAR UP AND DOWN SOLENOIDS

AIRPLANES WITH POWER STEERING HAVE THIS BOLT OMITTED

GEAR DOWN PORT POWER PACK ASSEMBLY

VENT PORT

NOSE GEAR EMERGENCY EXTENDED LINE

GEAR-UP PRESSURE SWITCH

GEAR-UP PORT MOTOR FLUID LEVEL SENSOR

HAND PUMP SUCTION PORT FILL PORT

MANUAL BLEED VALVE

FILL CAN

WASHER

BOLT

BONDING JUMPER

LOCK WASHER

WASHER

UC32B 073704AA.AI

Figure 202 Hydraulic Power Pack Assembly

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32-30-03 200200

LANDING GEAR LANDING GEAR POWER PACK VALVE HOUSING AND CONTROLS MAINTENANCE PRACTICES 1. PROCEDURES WARNING: When performing maintenance on a hydraulically operated landing gear system, be aware that any movement of a hydraulic actuator cylinder can displace hydraulic fluid and cause unanticipated movement of other actuator cylinders in the landing gear retraction system. Servicing of the landing gear hydraulic accumulator can also result in unanticipated movement of an actuator. Either action can result in an unsafe, unlocked landing gear system. Therefore, place the airplane on jacks prior to performing any inspection or maintenance. Cycle the landing gear and ensure that all three landing gears are down and locked prior to removing the aircraft from jacks.

A. Disassembly NOTE: The assembly area must be void of materials and equipment not directly associated with assembly of these units. Normal room temperature/humidity conditions prevail. Ventilation cannot allow outside small particle debris to settle in the area. (1) Perform the THREE POINT JACKING (PREFERRED (Ref. Chapter 07-10-00). All tires must be clear of the floor.

PROCEDURE)

procedure

(2) Remove the filter bowl (1) from the valve housing. Remove the packing (2) from the filter bowl (Ref. Figure 201). (3) Remove the filter element (3) from the valve housing. Remove the packing (4) from the filter element. (4) Remove the screws (5 and 6) and washers (7) attaching the selector valve (8) to the valve housing. (5) Remove the selector valve and packings (9) from the valve housing. (6) Remove the gear-up and hand pump pressure-operated check valves (10). Remove the packings (11, 12 and 13) and the packing retainers (14 and 15) from the check valves.

B. Assembly NOTE: The assembly area must be void of materials and equipment not directly associated with assembly of these units. Normal room temperature/humidity conditions prevail. Ventilation cannot allow outside small particle debris to settle in the area. (1) Clean and lubricate all parts in clean hydraulic fluid (39, Table 1, Chapter 91-00-00) just prior to assembly to provide initial lubrication and, in the case of screws to assure more accurate torque readings. (2) Install the packings (11, 12, and 13) and packing retainers (14 and 15) on the gear-up and hand pump pressure operated check valves (10) (Ref. Figure 201). (3) Install the gear-up and hand pump check valves (10) in the valve housing. Tighten the check valves 55 to 60 inch-pounds. Secure check valves to each other with lockwire.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (4) Install the packing (9) for the selector valve (8) in the valve housing. (5) Position the selector valve (8) on the valve housing. Secure with screws (5 and 6) and washers (7). (6) Torque the screws 45 to 50 inch-pounds. (7) Install a new packing (4) on a new filter element (3) and install the new filter element in the valve housing. (8) Install a new packing (2) on the filter bowl (1). (9) Install the filter bowl (1) in the valve housing (hand tight). (10) Fill and bleed the hydraulic fluid system as instructed under the heading LANDING GEAR HYDRAULIC SYSTEM FILLING AND BLEEDING in 32-30-00. (11) Perform the LOWERING THE AIRPLANE AFTER THREE-POINT JACKING procedure (Ref Chapter 7-10-00).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 201 Landing Gear Power Pack Valve Housing

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LANDING GEAR LANDING GEAR POWER PACK MOTOR MAINTENANCE PRACTICES

32-30-04 200200

1. PROCEDURES WARNING: When an airplane has experienced abnormal landing gear procedures of any type, as a safety precaution, place the airplane on jacks prior to performing any inspection or maintenance. Ensure that all three landing gears are down and locked prior to removing the airplane from jacks. When performing maintenance on a hydraulically operated landing gear system, be aware that any movement of a hydraulic actuator cylinder can displace hydraulic fluid and cause unanticipated movement of other actuator cylinders in the landing gear retraction system. Servicing of the landing gear hydraulic accumulator can also result in unanticipated movement of an actuator. Either action can result in an unsafe, unlocked landing gear system. Therefore, place the airplane on jacks prior to performing any inspection or maintenance. Cycle the landing gear and ensure that all three landing gears are down and locked prior to removing the airplane from jacks. Any time maintenance is to be performed on the landing gear system, place the airplane on jacks. When jacking the airplane in an unsheltered area where winds in excess of 35 kts will be encountered, never jack more than one gear at a time clear of the ground. Any time the landing gear is only partially retracted during maintenance, always cycle the gear with the power pack through at least one complete cycle before removing the airplane from the jacks. For safety reasons, pull the 2-ampere control circuit breaker on the pilot's inboard subpanel and place a note on the circuit breaker panel that LANDING GEAR MAINTENANCE IS IN PROCESS during maintenance on the landing gear. The landing gear control handle must never be moved from the down-and-locked position while the airplane is on the ground. It is recommended that the area be roped off during extension or retraction of the landing gear. CAUTION: Jacking of an airplane for the purpose of landing gear operation, inspection, servicing or maintenance, should be accomplished within an enclosed building or hangar. In the interest of safety, should it become necessary to jack the airplane in the open, wind velocity in any direction and terrain variations, must be compensated for prior to jacking the airplane. The landing gear must not be cycled with the power pack if low on fluid or if the landing gear system is not properly rigged. Use the emergency extension hand pump, TK229/939 hydraulic hand pump or TK229-1/939 air-driven hydraulic pump to extend and retract the landing gear for maintenance and rigging (Ref. Chapter 91-00-00).

32-30-04

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A. Removal (1) Perform the THREE-POINT JACKING (PREFERRED PROCEDURE) procedure (Ref. Chapter 07-10-00). All tires must be clear of the floor. Strut limiters will reduce strut extension (Ref. Chapter 91-00-00). (2) Remove the forward wing fillet, wing leading edge, and the upper wing panel (Ref. Figure 201). (3) Disconnect the electrical wiring from the motor. (4) Remove the two bolts securing the motor to the power pack (Ref. Figure 202). (5) Remove the motor from the power pack.

B. Installation (1) Fill the motor drive shaft coupling chamber one-third full and coat the coupling shaft splines with grease (83, Table 1, 91-00-00) (Ref. Figure 202). (2) Install the motor to the drive shaft coupling. Alignment pins on the motor are provided to ensure correct positioning of the motor. Install the two bolts securing the motor to the power pack and torque 72 to 84 inch-pounds. (3) Connect the electrical wiring to the motor. (4) Install the upper wing panel, wing leading edge, and forward wing fillet (Ref. Figure 201). WARNING: Before removing the airplane from the jacks, make sure that the landing gear emergency-extend hand pump handle is in the stowed position, the plunger on the service valve is pushed down with the hinged retainer in place, the landing gear control handle is in the DOWN position, the landing gear is down and locked and the accumulator charged to 800 ± 50 psi. (5) Perform the LOWERING (Ref. Chapter 07-10-00).

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32-30-04

AIRPLANE

AFTER

THREE-POINT

JACKING

procedure

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 201 Landing Gear Component Access

32-30-04

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 202 Power Pack Motor Assembly

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

LANDING GEAR POWER PACK GEAR UP PRESSURE SWITCH MAINTENANCE PRACTICES

32-30-05 200200

1. PROCEDURES WARNING: When an airplane has experienced abnormal landing gear procedures of any type, as a safety precaution, place the airplane on jacks prior to performing any inspection or maintenance. Ensure that all three landing gears are down and locked prior to removing the airplane from jacks. When performing maintenance on a hydraulically operated landing gear system, be aware that any movement of a hydraulic actuator cylinder can displace hydraulic fluid and cause unanticipated movement of other actuator cylinders in the landing gear retraction system. Servicing of the landing gear hydraulic accumulator can also result in unanticipated movement of an actuator. Either action can result in an unsafe, unlocked landing gear system. Therefore, place the airplane on jacks prior to performing any inspection or maintenance. Cycle the landing gear and ensure that all three landing gears are down and locked prior to removing the airplane from jacks. Any time maintenance is to be performed on the landing gear system, place the airplane on jacks. When jacking the airplane in an unsheltered area where winds in excess of 35 kts will be encountered, never jack more than one gear at a time clear of the ground. Any time the landing gear is only partially retracted during maintenance, always cycle the gear with the power pack through at least one complete cycle before removing the airplane from the jacks. For safety reasons, pull the 2-ampere control circuit breaker on the pilot's inboard subpanel and place a note on the circuit breaker panel that LANDING GEAR MAINTENANCE IS IN PROCESS during maintenance on the landing gear. The landing gear control handle must never be moved from the down-and-locked position while the airplane is on the ground. It is recommended that the area be roped off during extension or retraction of the landing gear. CAUTION: Jacking of an airplane for the purpose of landing gear operation, inspection, servicing or maintenance, should be accomplished within an enclosed building or hangar. In the interest of safety, should it become necessary to jack the airplane in the open, wind velocity in any direction and terrain variations, must be compensated for prior to jacking the airplane. The landing gear must not be cycled with the power pack if low on fluid or if the landing gear system is not properly rigged. Use the emergency extension hand pump, TK229/939 hydraulic hand pump or TK229-1/939 air-driven hydraulic pump to extend and retract the landing gear for maintenance and rigging (Ref. Chapter 91-00-00).

32-30-05

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

A. Removal (1) Perform the THREE-POINT JACKING (PREFERRED PROCEDURE) procedure (Ref. Chapter 07-10-00). All tires must be clear of the floor. Strut limiters will reduce strut extension (Ref. Chapter 91-00-00). (2) Perform REMOVING GROUND POWER procedure (Ref. Chapter 24-40-00). (3) Perform BATTERY DISCONNECTION procedure (Ref. Chapter 24-31-00). (4) Remove the forward wing fillet, wing leading edge, and upper wing panel (Ref. Figure 201). (5) Disconnect the electrical wiring from the gear up pressure switch (Ref. Figure 202). NOTE: Shop towels will be needed to absorb the system fluid which will escape from the port when the switch is removed. (6) Unscrew the gear up pressure switch from the power pack.

B. Installation (1) Install a new packing on the gear up pressure switch (Ref. Figure 202). (2) Install the gear up pressure switch by screwing it into the power pack. Torque the pressure switch 50 to 60 inch-pounds. (3) Connect the electrical wiring to the switch. (4) Perform the HYDRAULIC SYSTEM FILLING AND BLEEDING procedure to fill and bleed the landing gear hydraulic system (Ref. 32-30-00). (5) Install the upper wing panel, wing leading edge and forward wing fillet (Ref. Figure 201). (6) Perform LOWERING THE (Ref. Chapter 07-10-00).

AIRPLANE

AFTER

THREE-POINT

JACKING

(7) Perform BATTERY CONNECTION procedure (Ref. Chapter 24-31-00). (8) Perform APPLYING GROUND POWER procedure (Ref. Chapter 24-40-00).

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procedure

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 201 Landing Gear Component Access

32-30-05

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SWIVEL FITTING AUXILIARY RETURN PORT (POWER STEERING ONLY)

BOLT AN960-416L WASHER AN960-416 WASHER (TYPICAL 6 PLACES) MOUNTING BRACKET

HAND PUMP PRESSURE PORT

GEAR UP AND DOWN SOLENOIDS

AIRPLANES WITH POWER STEERING HAVE THIS BOLT OMITTED

GEAR DOWN PORT POWER PACK ASSEMBLY

VENT PORT

NOSE GEAR EMERGENCY EXTENDED LINE

GEAR-UP PRESSURE SWITCH

GEAR-UP PORT MOTOR FLUID LEVEL SENSOR

HAND PUMP SUCTION PORT FILL PORT

FILL CAN

MANUAL BLEED VALVE

WASHER

BOLT

BONDING JUMPER

LOCK WASHER

WASHER

Figure 202 Hydraulic Power Pack Assembly

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UC32B 073704AA.AI

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32-30-06 200200

LANDING GEAR LANDING GEAR HYDRAULIC POWER PACK FILTERS MAINTENANCE PRACTICES 1. PROCEDURES WARNING: When an airplane has experienced abnormal landing gear procedures of any type, as a safety precaution, place the airplane on jacks prior to performing any inspection or maintenance. Ensure that all three landing gears are down and locked prior to removing the airplane from jacks. When performing maintenance on a hydraulically operated landing gear system, be aware that any movement of a hydraulic actuator cylinder can displace hydraulic fluid and cause unanticipated movement of other actuator cylinders in the landing gear retraction system. Servicing of the landing gear hydraulic accumulator can also result in unanticipated movement of an actuator. Either action can result in an unsafe, unlocked landing gear system. Therefore, place the airplane on jacks prior to performing any inspection or maintenance. Cycle the landing gear and ensure that all three landing gears are down and locked prior to removing the airplane from jacks. Any time maintenance is to be performed on the landing gear system, place the airplane on jacks. When jacking the airplane in an unsheltered area where winds in excess of 35 kts will be encountered, never jack more than one gear at a time clear of the ground. Any time the landing gear is only partially retracted during maintenance, always cycle the gear with the power pack through at least one complete cycle before removing the airplane from the jacks. For safety reasons, pull the 2-ampere control circuit breaker on the pilot's inboard subpanel and place a note on the circuit breaker panel that LANDING GEAR MAINTENANCE IS IN PROCESS during maintenance on the landing gear. The landing gear control handle must never be moved from the down-and-locked position while the airplane is on the ground. It is recommended that the area be roped off during extension or retraction of the landing gear. CAUTION: Jacking of an airplane for the purpose of landing gear operation, inspection, servicing or maintenance, should be accomplished within an enclosed building or hangar. In the interest of safety, should it become necessary to jack the airplane in the open, wind velocity in any direction and terrain variations, must be compensated for prior to jacking the airplane. The landing gear must not be cycled with the power pack if low on fluid or if the landing gear system is not properly rigged. Use the emergency extension hand pump, TK229/939 hydraulic hand pump or TK229-1/939 air-driven hydraulic pump to extend and retract the landing gear for maintenance and rigging (Ref. Chapter 91-00-00).

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A. Removal The system filter mounted on the power pack assembly will be replaced in accordance with the detailed inspection program. (1) Perform REMOVING GROUND POWER procedure (Ref. Chapter 24-40-00). (2) Perform BATTERY DISCONNECTION procedure (Ref. Chapter 24-31-00). (3) Perform the THREE-POINT JACKING (PREFERRED PROCEDURE) procedure (Ref. Chapter 07-10-00). All tires must be clear of the floor. Strut limiters will reduce strut extension (Ref. Chapter 91-00-00). (4) Remove the forward wing fillet, wing leading edge, and lower wing panel. NOTE: Hydraulic fluid will escape when the filter element is removed from the power pack. To prevent excessive leakage, plug the port until the new filter element is installed. Use shop rags and a container to catch the released hydraulic fluid. (5) Remove safety wire and unscrew the filter housing (1) from the power pack (Ref. Figure 201). (6) Remove the filter element (3) from the power pack.

B. Installation (1) Install a new packing (4) on the filter element (Ref. Figure 201). (2) Install the hydraulic system filter element into the power pack. Make sure the filter element is properly mated to the power pack. (3) Install a new packing (2) on the filter housing (1) and screw the housing into the power pack. Safety wire the filter housing to the power pack. (4) Perform the HYDRAULIC SYSTEM FILLING AND BLEEDING procedure to fill and bleed the landing gear hydraulic system (Ref. 32-30-00). (5) Install the forward wing fillet, wing leading edge, and lower wing panel. WARNING: Before removing the airplane from the jacks, make sure that the landing gear emergency-extend hand pump handle is in the stowed position, the plunger on the service valve is pushed down with the hinged retainer in place, the landing gear control handle is in the DOWN position, the landing gear is down and locked and the accumulator charged to 800 ± 50 psi. (6) Perform the LOWERING THE AIRPLANE AFTER THREE-POINT JACKING procedure (Ref. Chapter 07-10-00). (7) Perform BATTERY CONNECTION procedure (Ref. Chapter 24-31-00). (8) Perform APPLYING GROUND POWER procedure (Ref. Chapter 24-40-00).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

1. FILTER HOUSING 2. PACKING 3. FILTER ELEMENT 4. PACKING

4

3

2

1

UC32B 063192AA.AI

Figure 201 Hydraulic System Filter

2. LANDING GEAR HYDRAULIC POWER PACK BLEED AIR FILTER A. Removal (1) Remove the wing leading edge panel 55 (Ref. Chapter 06-50-00, WING ACCESS PANELS). (2) Remove the clamp (2) by removing the attaching screw (Ref. Figure 202). (3) Disconnect the bleed air filter (1) from the check valve (4) and the bleed air line.

B. Installation (1) Install a new packing (3) on the power pack check valve flare fitting (4) (Ref. Figure 202). (2) Position and connect the filter (1) to the bleed air line and the check valve (4). (3) Position and secure the clamp (2) with the attaching screw. (4) Install the wing leading edge panel 55 (Ref. Chapter 06-50-00, WING ACCESS PANELS).

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C. Cleaning (1) Perform POWER PACK BLEED AIR FILTER REMOVAL procedure in this section. (2) Blow shop air through filter (maximum 18 psi). (3) If the filter’s condition will not allow low pressure air (5 psi or less) to flow freely, the filter must be replaced. (4) Perform POWER PACK BLEED AIR FILTER INSTALLATION procedure in this section.

1 2

3

4

1. FILTER 2. CLAMP 3. PACKING 4. CHECK VALVE UC32B 001460AB.AI

Figure 202 Landing Gear Hydraulic Power Pack Bleed Air Filter

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32-30-07 200200

LANDING GEAR LANDING GEAR POWER PACK GEAR-UP AND GEAR-DOWN PORT FILTER SCREEN MAINTENANCE PRACTICES 1. PROCEDURES WARNING: When an airplane has experienced abnormal landing gear procedures of any type, as a safety precaution, place the airplane on jacks prior to performing any inspection or maintenance. Ensure that all three landing gears are down and locked prior to removing the airplane from jacks. When performing maintenance on a hydraulically operated landing gear system, be aware that any movement of a hydraulic actuator cylinder can displace hydraulic fluid and cause unanticipated movement of other actuator cylinders in the landing gear retraction system. Servicing of the landing gear hydraulic accumulator can also result in unanticipated movement of an actuator. Either action can result in an unsafe, unlocked landing gear system. Therefore, place the airplane on jacks prior to performing any inspection or maintenance. Cycle the landing gear and ensure that all three landing gears are down and locked prior to removing the airplane from jacks. Any time maintenance is to be performed on the landing gear system, place the airplane on jacks. When jacking the airplane in an unsheltered area where winds in excess of 35 kts will be encountered, never jack more than one gear at a time clear of the ground. Any time the landing gear is only partially retracted during maintenance, always cycle the gear with the power pack through at least one complete cycle before removing the airplane from the jacks. For safety reasons, pull the 2-ampere control circuit breaker on the pilot's inboard subpanel and place a note on the circuit breaker panel that LANDING GEAR MAINTENANCE IS IN PROCESS during maintenance on the landing gear. The landing gear control handle must never be moved from the down-and-locked position while the airplane is on the ground. It is recommended that the area be roped off during extension or retraction of the landing gear. CAUTION: Jacking of an airplane for the purpose of landing gear operation, inspection, servicing or maintenance, should be accomplished within an enclosed building or hangar. In the interest of safety, should it become necessary to jack the airplane in the open, wind velocity in any direction and terrain variations, must be compensated for prior to jacking the airplane. The landing gear must not be cycled with the power pack if low on fluid or if the landing gear system is not properly rigged. Use the emergency extension hand pump, TK229/939 hydraulic hand pump or TK229-1/939 air-driven hydraulic pump to extend and retract the landing gear for maintenance and rigging (Ref. Chapter 91-00-00).

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A. Removal (1) Perform the THREE-POINT JACKING (PREFERRED PROCEDURE) procedure (Ref. Chapter 07-10-00). All tires must be clear of the floor. Strut limiters will reduce strut extension (Ref. Chapter 91-00-00). (2) Perform REMOVING GROUND POWER procedure (Ref. Chapter 24-40-00). (3) Perform BATTERY DISCONNECTION procedure (Ref. Chapter 24-31-00). (4) Remove the forward wing fillet, wing leading edge, and the upper and lower wing panels (Ref. Figure 202). (5) Disconnect and cap the hydraulic lines connected to the gear-up and gear-down ports of the power pack (Ref. Figure 201). (6) Remove the unions (1) from the gear-up and gear-down ports. (7) Using an Allen wrench, unscrew the permanent screens (2) from the gear-up and gear-down ports.

B. Installation (1) Screw the permanent filter screens (2) into the gear-up and gear-down ports of the power pack (Ref. Figure 201). (2) Install new packings (3) on the unions (1) and install the unions in the gear-up and gear-down ports of the power pack. (3) Remove the caps and connect the hydraulic lines to the gear-up and gear-down ports of the power pack. (4) Perform the HYDRAULIC SYSTEM FILLING AND BLEEDING procedure to fill and bleed the landing gear hydraulic system (Ref. 32-30-00). (5) Install the upper and lower wing panels, wing leading edge and the forward wing fillet (Ref. Figure 202). WARNING: Before removing the airplane from the jacks, make sure that the landing gear emergency-extend hand pump handle is in the stowed position, the plunger on the service valve is pushed down with the hinged retainer in place, the landing gear control handle is in the DOWN position, the landing gear is down and locked and the accumulator charged to 800 ± 50 psi. (6) Perform the LOWERING THE AIRPLANE AFTER THREE-POINT JACKING procedure (Ref. Chapter 07-10-00). (7) Perform BATTERY CONNECTION procedure (Ref. Chapter 24-31-00). (8) Perform APPLYING GROUND POWER procedure (Ref. Chapter 24-40-00).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

C. Cleaning (1) Perform the LANDING GEAR POWER PACK GEAR-UP AND GEAR-DOWN PORT FILTER SCREEN REMOVAL procedure in this section. (2) Clean the permanent filter screens with solvent (2, Table 1, Chapter 91-00-00). Blow dry with clean dry air. (3) Perform the LANDING GEAR POWER PACK GEAR-UP AND GEAR-DOWN PORT FILTER SCREEN INSTALLATION procedure in this section.

1 3 2

GEAR DOWN PORT

1. UNION 2. PERMANENT SCREEN 3. PACKING

2 3 1

GEAR UP PORT UC32B 063193AA.AI

Figure 201 Power Pack Filter Screens

32-30-07

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Figure 202 Landing Gear Component Access

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32-30-08 200200

LANDING GEAR LANDING GEAR POWER PACK FLUID LEVEL SENSOR MAINTENANCE PRACTICES 1. PROCEDURES WARNING: When an airplane has experienced abnormal landing gear procedures of any type, as a safety precaution, place the airplane on jacks prior to performing any inspection or maintenance. Ensure that all three landing gears are down and locked prior to removing the airplane from jacks. When performing maintenance on a hydraulically operated landing gear system, be aware that any movement of a hydraulic actuator cylinder can displace hydraulic fluid and cause unanticipated movement of other actuator cylinders in the landing gear retraction system. Servicing of the landing gear hydraulic accumulator can also result in unanticipated movement of an actuator. Either action can result in an unsafe, unlocked landing gear system. Therefore, place the airplane on jacks prior to performing any inspection or maintenance. Cycle the landing gear and ensure that all three landing gears are down and locked prior to removing the airplane from jacks. Any time maintenance is to be performed on the landing gear system, place the airplane on jacks. When jacking the airplane in an unsheltered area where winds in excess of 35 kts will be encountered, never jack more than one gear at a time clear of the ground. Any time the landing gear is only partially retracted during maintenance, always cycle the gear with the power pack through at least one complete cycle before removing the airplane from the jacks. For safety reasons, pull the 2-ampere control circuit breaker on the pilot's inboard subpanel and place a note on the circuit breaker panel that LANDING GEAR MAINTENANCE IS IN PROCESS during maintenance on the landing gear. The landing gear control handle must never be moved from the down-and-locked position while the airplane is on the ground. It is recommended that the area be roped off during extension or retraction of the landing gear. CAUTION: Jacking of an airplane for the purpose of landing gear operation, inspection, servicing or maintenance, should be accomplished within an enclosed building or hangar. In the interest of safety, should it become necessary to jack the airplane in the open, wind velocity in any direction and terrain variations, must be compensated for prior to jacking the airplane. The landing gear must not be cycled with the power pack if low on fluid or if the landing gear system is not properly rigged. Use the emergency extension hand pump, TK229/939 hydraulic hand pump or TK229-1/939 air-driven hydraulic pump to extend and retract the landing gear for maintenance and rigging (Ref. Chapter 91-00-00).

32-30-08

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

A. Removal (1) Perform the THREE-POINT JACKING (PREFERRED PROCEDURE) procedure (Ref. Chapter 07-10-00). All tires must be clear of the floor. Strut limiters will reduce strut extension (Ref. Chapter 91-00-00). (2) Perform REMOVING GROUND POWER procedure (Ref. Chapter 24-40-00). (3) Perform BATTERY DISCONNECTION procedure (Ref. Chapter 24-31-00). (4) Remove the forward wing fillet, wing leading edge, and upper wing panel (Ref. Figure 202). (5) Disconnect the electrical wiring from the fluid level sensor (Ref. Figure 201). NOTE: A suitable container will be needed to catch hydraulic fluid which will drain from the port when the hydraulic fluid level sensor is removed. (6) Remove the fluid level sensor from the power pack. Immediately cap the open port.

B. Installation (1) Install a new packing on the fluid level sensor (Ref. Figure 201). (2) Install the hydraulic fluid level sensor into the power pack and torque 30 to 50 inch-pounds. (3) Connect the electrical wiring to the fluid level sensor. (4) Perform the HYDRAULIC SYSTEM FILLING AND BLEEDING procedure to fill and bleed the landing gear hydraulic system (Ref. 32-30-00). WARNING: Before removing the airplane from the jacks, make sure that the landing gear emergency-extend hand pump handle is in the stowed position, the plunger on the service valve is pushed down with the hinged retainer in place, the landing gear control handle is in the DOWN position, the landing gear is down and locked and the accumulator charged to 800 ± 50 psi. (5) Install the upper wing panel, wing leading edge and forward wing fillet (Ref. Figure 202). (6) Perform the LOWERING THE AIRPLANE AFTER THREE-POINT JACKING procedure (Ref. Chapter 07-10-00). (7) Perform BATTERY CONNECTION procedure (Ref. Chapter 24-31-00). (8) Perform APPLYING GROUND POWER procedure (Ref. Chapter 24-40-00).

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C. Functional Test (1) Perform the POWER PACK FLUID LEVEL SENSOR REMOVAL procedure in this section. (2) Connect the electrical connector to the fluid level sensor. (3) Perform BATTERY CONNECTION procedure (Ref. Chapter 24-31-00). (4) Perform APPLYING GROUND POWER procedure (Ref. Chapter 24-40-00). (5) Select the BATT switch to the ON position. (6) Select the EXT PWR switch to the EXT PWR position. (7) Position one person in the flight compartment and one at the fluid level sensor. (8) Observe the HYD FLUID LOW annunciator is illuminated. NOTE: The container should be a solid container that will block light with a lid that has a suitable hole that will allow just the sensor to fit. (9) Submerge the sensor into a container with at least two inches of hydraulic fluid (39, Table 1, Chapter 91-00-00). (10) The HYD FLUID LOW annunciator should extinguish within a few seconds. (11) With the sensor submerged, press and hold the PRESS TO TEST switch on the annunciator. The HYD FLUID LOW annunciator should illuminate within a few seconds. (12) Release the PRESS TO TEST switch. The annunciator should extinguish within a few seconds. (13) While observing the annunciator remove the sensor from the container. The annunciator should illuminate within a few seconds. (14) If any of the above requirements are not met, troubleshoot the system. (15) If all the above checks are satisfactory perform the POWER PACK FLUID LEVEL SENSOR INSTALLATION procedure in this section. (16) Select the EXT PWR switch to the OFF position. (17) Select the BATT switch to the OFF position. (18) Perform REMOVING GROUND POWER procedure (Ref. Chapter 24-40-00).

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SWIVEL FITTING AUXILIARY RETURN PORT (POWER STEERING ONLY)

BOLT AN960-416L WASHER AN960-416 WASHER (TYPICAL 6 PLACES) MOUNTING BRACKET

HAND PUMP PRESSURE PORT

GEAR UP AND DOWN SOLENOIDS

AIRPLANES WITH POWER STEERING HAVE THIS BOLT OMITTED

GEAR DOWN PORT POWER PACK ASSEMBLY

VENT PORT

NOSE GEAR EMERGENCY EXTENDED LINE

GEAR-UP PRESSURE SWITCH

GEAR-UP PORT MOTOR FLUID LEVEL SENSOR

HAND PUMP SUCTION PORT FILL PORT

FILL CAN

MANUAL BLEED VALVE

WASHER

BOLT

BONDING JUMPER

LOCK WASHER

WASHER

Figure 201 Hydraulic Power Pack Assembly

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UC32B 073704AA.AI

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 202 Landing Gear Component Access

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LANDING GEAR MAIN LANDING GEAR MAINTENANCE PRACTICES

32-30-09 200200

1. PROCEDURES WARNING: When an airplane has experienced abnormal landing gear procedures of any type, as a safety precaution, place the airplane on jacks prior to performing any inspection or maintenance. Ensure that all three landing gears are down and locked prior to removing the airplane from jacks. When performing maintenance on a hydraulically operated landing gear system, be aware that any movement of a hydraulic actuator cylinder can displace hydraulic fluid and cause unanticipated movement of other actuator cylinders in the landing gear retraction system. Servicing of the landing gear hydraulic accumulator can also result in unanticipated movement of an actuator. Either action can result in an unsafe, unlocked landing gear system. Therefore, place the airplane on jacks prior to performing any inspection or maintenance. Cycle the landing gear and ensure that all three landing gears are down and locked prior to removing the airplane from jacks. Any time maintenance is to be performed on the landing gear system, place the airplane on jacks. When jacking the airplane in an unsheltered area where winds in excess of 35 kts will be encountered, never jack more than one gear at a time clear of the ground. Any time the landing gear is only partially retracted during maintenance, always cycle the gear with the power pack through at least one complete cycle before removing the airplane from the jacks. For safety reasons, pull the 2-ampere control circuit breaker on the pilot's inboard subpanel and place a note on the circuit breaker panel that LANDING GEAR MAINTENANCE IS IN PROCESS during maintenance on the landing gear. The landing gear control handle must never be moved from the down-and-locked position while the airplane is on the ground. It is recommended that the area be roped off during extension or retraction of the landing gear. CAUTION: Jacking of an airplane for the purpose of landing gear operation, inspection, servicing or maintenance, should be accomplished within an enclosed building or hangar. In the interest of safety, should it become necessary to jack the airplane in the open, wind velocity in any direction and terrain variations, must be compensated for prior to jacking the airplane. The landing gear must not be cycled with the power pack if low on fluid or if the landing gear system is not properly rigged. Use the emergency extension hand pump, TK229/939 hydraulic hand pump or TK229-1/939 air-driven hydraulic pump to extend and retract the landing gear for maintenance and rigging (Ref. Chapter 91-00-00).

32-30-09

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A. Rigging A service valve in the left hand wing, adjacent to the power pack, is provided for use with the emergency extension hand pump to raise and lower the landing gear while the landing gear is being rigged. The service valve is normally in a mode to extend the landing gear (red knob pushed down). The service valve is accessible through a door in the wing skin just inboard of the LH nacelle. The red knob cannot be pulled up until the hinged retainer covering it is unlatched and moved out of the way. The landing gear must not be cycled electrically until the landing gear is properly rigged. With the airplane on jacks and an external power supply connected and adjusted to 28 ± 0.25 volts, the landing gear is retracted by pulling the red knob on the service valve up and then operating the emergency extension hand pump. When the red knob is pulled up, the two switches mounted on the service valve are actuated to open the circuit to the power pack motor and to complete a circuit from the 2-ampere control circuit breaker to the retract solenoid on the gear selector valve. The same action positions the service valve to route the hand pump pressure fluid to the normal retract mode plumbing to the actuators. Remove the pump handle from the securing clip and pump the handle up and down to retract the landing gear. After the landing gear is retracted, position the emergency extend hand handle in the stowed position and turn off the electrical power. To extend the landing gear push the red knob on the service valve down and extend the landing gear using the emergency extension hand pump; the electrical power does not have to be on to extend the landing gear. For safety reasons, pull the 2-ampere control circuit breaker on the pilot's inboard subpanel. CAUTION: If the red knob on the service valve is pushed down while the landing gear is retracted, the electrical power is on and the landing gear control handle is in the down position, the landing gear will extend immediately. To prevent serious damage to the pump, never operate the power pack when the engines are not operating without supplying 18 psi of regulated dry air to pressurize the power pack reservoir in place of the engine bleed air. When cycling the landing gear with the power pack, do not exceed three cycles in the first six minutes of operation with approximately a two-minute cooling period between cycles, then with a five-minute cooling interval between each cycle. Before removing the airplane from the jacks, make sure that the emergency extension hand pump handle is in the stowed position and that the red knob on the service valve is pushed down with the hinged retainer in place. Retract and extend the landing gear, using the hydraulic power pack to ensure the landing gear is in the down-and-locked position before removing the airplane from the jacks. CAUTION: Do not fully cycle the landing gear electrically until the gear is properly rigged. Use of the emergency hand pump is recommended for initial rigging. (1) Perform the THREE-POINT JACKING (PREFERRED PROCEDURE) procedure (Ref. Chapter 07-10-00). All tires must be clear of the floor. Strut limiters will reduce strut extension (Ref. Chapter 91-00-00). (2) Perform APPLYING GROUND POWER procedure (Ref. Chapter 24-40-00). (3) Disconnect the door actuation cams from the doors by removing the cotter pins, nuts, washers, and bolts from the lower attaching links. Secure the doors out of the way with safety wire. NOTE: Use the emergency extension hand pump to retract or extend the actuator during this rigging procedure. (4) Retract the gear slightly to take the load off the drag brace and to unlock the actuator.

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32-30-09

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (5) Remove the nut, bolt, and washers attaching the actuator rod end to the upper drag leg arm. Remove the cotter pin, nut, washers, bushings, and bolt attaching the lower drag leg to the strut (Ref. Figure 201). (6) Check the alignment of the lower drag leg with the attachment lug on the landing gear strut; there shall be no side loads on the drag leg when the landing gear is in the extended or retracted position. Correct misalignment by shifting AN960-1016L washers from side to side of the wheel well attach fitting and upper drag legs. A minimum of one washer must be maintained between the upper drag legs and the wheel well attach fitting. After the proper alignment is obtained, connect the lower drag leg to the landing gear strut. NOTE: Verify continuity of the left main actuator switch by checking continuity between pins A and E then H and K, of connector P197. Verify continuity of the right main actuator switch by checking continuity between pins A and E then H and K, of connector P198. (7) With the actuator fully retracted and locked, and the landing gear fully extended, check the alignment of the actuator rod end with the upper drag leg arm. The actuator rod end must slip into the upper drag leg arm without causing any side loads on the actuator piston rod. With the actuator fully extended, manually push the landing gear to the retracted position and again check for any side loads on the actuator piston rod. If any misalignment is found in either position, shift the AN960-716L washers as required to correct the misalignment. A minimum of one AN960-716L washer is required between the actuator and the support structure. (8) Cut the safety wire and loosen the jam nut on the actuator rod end. With the actuator fully retracted and the landing gear fully extended, adjust the actuator rod end to align with nut attaching hole in the upper drag leg arm. Tighten the jam nut. (9) Lubricate the actuator rod end bolt with grease (80, Table 1, Chapter 91-00-00) at time of installation. Attach the actuator rod end to the upper drag leg arm with the nut, bolt, and washers. (10) With the landing gear fully extended, apply moderate pressure to the drag brace assembly to bottom out any end play in the actuator. Measure the amount of clearance between the upper drag leg and the rig plate. If a clearance of 0.005 +0.005/- 0.004 inch at the closest point does not exist, loosen the jam nut and lengthen the actuator rod end to the next keying position by rotating the actuator piston rod. Tighten the jam nut and safety wire to the key washer. (11) Retract the landing gear fully and measure the clearance between the landing gear piston and the main spar. If the clearance is not 2.28 +0.25/-0.09 inches, remove the actuator from the airplane and adjust the actuator stroke as outlined in Step (12). (12) The stroke length of the actuator (6.23 ± 0.12 inches) is preset by the supplier and should not require adjustment. If the clearance measured in Step (9) is not within tolerance, remove the actuator from the airplane and adjust the actuator stroke length as follows: (a) Remove the safety wire and loosen the end cap jam nut. Rotate the end cap one full turn at a time until the proper adjustment is obtained. When the end cap is rotated out, the stroke length is increased and when rotated in, the stroke length is decreased. Each full turn of the end cap will change the stroke length by 0.06 inch.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL NOTE: To ensure proper alignment of the hydraulic plumbing to the shuttle valve located in the end cap, the end cap must be rotated in 360° rotations only. If more than two complete rotations of the end cap are required to obtain the proper stroke adjustment, structural damage may be present and the airplane structure must be inspected. (b) After the stroke length is properly adjusted, tighten the jam nut and safety wire. (c) Perform the MAIN LANDING GEAR ACTUATOR INSTALLATION procedure in this section (Ref. Chapter 32-30-10). (d) Perform the HYDRAULIC SYSTEM FILLING AND BLEEDING procedure to fill and bleed the landing gear hydraulic system (Ref. 32-30-00). (13) Lubricate the landing gear and drag brace hinge points with the proper lubricant (Ref. Chapter 12-20-00, MAIN LANDING GEAR LUBRICATION). (14) Cycle the landing gear and check for proper operation of the in-transit and gear-down lights. If necessary, adjust the downlock, up-position, and actuator downlock switches (Ref. 32-60-00). CAUTION: When cycling the landing gear with the power pack, do not exceed three cycles in the first six minutes of operation with approximately a two minute cooling period between cycles, then with a five minute cooling interval between each cycle. (15) Connect the landing gear doors. If necessary, perform the MAIN LANDING GEAR DOOR RIGGING procedure (Ref. Chapter 32-30-12). WARNING: Before removing the airplane from the jacks, make sure that the landing gear emergency-extend hand pump handle is in the stowed position, the plunger on the service valve is pushed down with the hinged retainer in place, the landing gear control handle is in the DOWN position, the landing gear is down and locked and the accumulator charged to 800 ± 50 psi. (16) Perform REMOVING GROUND POWER procedure (Ref. Chapter 24-40-00). (17) Perform the LOWERING THE AIRPLANE AFTER THREE-POINT JACKING procedure (Ref. Chapter 07-10-00).

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Figure 201 Main Landing Gear Rigging

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LANDING GEAR MAIN LANDING GEAR ACTUATOR MAINTENANCE PRACTICES

32-30-10 200200

1. PROCEDURES WARNING: When an airplane has experienced abnormal landing gear procedures of any type, as a safety precaution, place the airplane on jacks prior to performing any inspection or maintenance. Ensure that all three landing gears are down and locked prior to removing the airplane from jacks. When performing maintenance on a hydraulically operated landing gear system, be aware that any movement of a hydraulic actuator cylinder can displace hydraulic fluid and cause unanticipated movement of other actuator cylinders in the landing gear retraction system. Servicing of the landing gear hydraulic accumulator can also result in unanticipated movement of an actuator. Either action can result in an unsafe, unlocked landing gear system. Therefore, place the airplane on jacks prior to performing any inspection or maintenance. Cycle the landing gear and ensure that all three landing gears are down and locked prior to removing the airplane from jacks. Any time maintenance is to be performed on the landing gear system, place the airplane on jacks. When jacking the airplane in an unsheltered area where winds in excess of 35 kts will be encountered, never jack more than one gear at a time clear of the ground. Any time the landing gear is only partially retracted during maintenance, always cycle the gear with the power pack through at least one complete cycle before removing the airplane from the jacks. For safety reasons, pull the 2-ampere control circuit breaker on the pilot's inboard subpanel and place a note on the circuit breaker panel that LANDING GEAR MAINTENANCE IS IN PROCESS during maintenance on the landing gear. The landing gear control handle must never be moved from the down-and-locked position while the airplane is on the ground. It is recommended that the area be roped off during extension or retraction of the landing gear. CAUTION: Jacking of an airplane for the purpose of landing gear operation, inspection, servicing or maintenance, should be accomplished within an enclosed building or hangar. In the interest of safety, should it become necessary to jack the airplane in the open, wind velocity in any direction and terrain variations, must be compensated for prior to jacking the airplane. The landing gear must not be cycled with the power pack if low on fluid or if the landing gear system is not properly rigged. Use the emergency extension hand pump, TK229/939 hydraulic hand pump or TK229-1/939 air-driven hydraulic pump to extend and retract the landing gear for maintenance and rigging (Ref. Chapter 91-00-00).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

A. Removal CAUTION: To prevent damage to the landing gear actuator rod end ball bearing, inspect the actuator rod end ball bearing for contamination from oil, grease, solvents, degreaser, anti-ice fluid or other contaminants. The actuator rod end ball bearing is coated with a dry-film permanent lubricant and contaminants could wash away the dry-film lubricant on the rod end ball bearing. During servicing of the gear and/or where cleaning is required, protect the actuator rod end ball bearing from all agents sprayed or applied. (1) Perform THREE-POINT JACKING (PREFERRED PROCEDURE) procedure (Ref. Chapter 07-10-00). All tires must be clear of the floor. Strut limiters will reduce strut extension (Ref. Chapter 91-00-00). (2) Perform REMOVING GROUND POWER procedure (Ref. Chapter 24-40-00). (3) Perform BATTERY DISCONNECTION procedure (Ref. Chapter 24-31-00). (4) Disconnect the door actuation cams from the doors by removing the cotter pins, nuts, washers, and bolts from the lower attaching links. Secure the doors out of the way with safety wire. CAUTION: As the hydraulic hoses are disconnected, plug or cap all openings to prevent entry of foreign material into the hoses or actuator. (5) Remove the three hydraulic hoses from the actuator and identify to facilitate installation (Ref. Figure 201). (6) Remove the nut, bolt (1), and washers attaching the actuator rod end to the upper drag leg arm (2). (7) Disconnect the actuator downlock switch receptacle plug located in the upper rear of the wheel well. (8) Remove the nuts, bolts (3), bushings (4), and washers (5 and 6) attaching the actuator to the support structure. To maintain alignment of the actuator to the drag brace, retain the washers (5) between the actuator and support structure in the same position as when removed. (9) Remove the actuator downward out of the wheel well.

B. Installation NOTE: It is recommended to use the same manufacturer’s actuators on any one airplane, but it is permissible to intermix actuator part numbers and manufacturers on any one airplane. (1) Position the actuator in the wheel well and install the bushings (4), washers (5 and 6), bolts (3), and nuts to secure the actuator to the support structure. Install the same number of AN960-716L washers (5) on each side of the actuator (between the actuator and the support structure) that were removed with the actuator (Ref. Figure 201). (2) Connect the hydraulic hoses to the actuator. If the unions between the actuator and hose ends were removed or loosened, install new packings. (3) Connect the actuator downlock switch wiring in the upper rear of the wheel well.

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32-30-10

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (4) With the actuator fully retracted and the landing gear fully extended, check the alignment of the actuator rod end with the upper drag leg arm. The actuator rod end must slip into the upper drag leg arm without causing any side loads on the actuator piston rod. With the actuator fully extended, manually push the landing gear to the retracted position and again check for any side loads on the actuator piston rod. If any misalignment is found in either position, shift the AN960-716L washers (5) as required to correct the misalignment. A minimum of one AN960-716L washer is required between the actuator and the support structure. NOTE: To extend and retract the actuator piston rod, use the Emergency Extension Hand Pump. Refer to LANDING GEAR HAND PUMP CYCLING procedure (Ref. 32-30-00). (5) Lubricate bolt (1) with grease (80, Table 1, Chapter 91-00-00) at time of installation. (6) With the actuator fully retracted and the landing gear fully extended, check the alignment of the actuator rod end with the attaching hole in the upper drag leg arm. If the holes align, attach the actuator rod end to the upper drag leg arm with the nut, bolt (1) and washers. NOTE: If the actuator rod end will not align with the hole in the upper drag leg arm, if the actuator stroke length or rod end length has been changed, or if the actuator being installed is not the one that was removed, rig the landing gear (Ref. 32-30-09, MAIN LANDING GEAR RIGGING). (7) Install the bolts, washers, nuts, and cotter pins connecting the lower attaching links to the door actuation cams. (8) Perform HYDRAULIC SYSTEM FILLING AND BLEEDING procedure to fill and bleed the landing gear hydraulic system (Ref. 32-30-00). (9) Perform BATTERY CONNECTION procedure (Ref. Chapter 24-31-00). (10) Perform APPLYING GROUND POWER procedure (Ref. Chapter 24-40-00). CAUTION: To prevent serious damage to the pump, never operate the power pack when the engines are not running without supplying 18 psi of regulated dry air to the manual bleed valve to pressurize the reservoir. When cycling the landing gear with the power pack, do not exceed three cycles in the first six minutes of operation with approximately a two minute cooling period between cycles, then with a five minute cooling interval between each cycle. (11) Cycle the landing gear with the power pack, and check for proper operation of the in-transit and gear down lights. If necessary, adjust the downlock, up position, and actuator downlock switches (Ref. 32-60-00). WARNING: Before removing the airplane from the jacks, make sure that the landing gear emergency-extend hand pump handle is in the stowed position, the plunger on the service valve is pushed down with the hinged retainer in place, the landing gear control handle is in the DOWN position, the landing gear is down and locked and the accumulator is charged to 800 ± 50 psi. (12) Perform REMOVING GROUND POWER procedure (Ref. Chapter 24-40-00). (13) Perform LOWERING THE (Ref. Chapter 07-10-00).

AIRPLANE

AFTER

THREE-POINT

JACKING

32-30-10

procedure

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2. MAIN LANDING GEAR ACTUATOR SHUTTLE VALVE A. Functional Test This procedure may be performed with the actuator on the bench or installed in the airplane. If the procedure is to be performed with the actuator on the bench, perform the MAIN LANDING GEAR ACTUATOR REMOVAL procedure in this section then proceed to Step (4) of this procedure. (1) Perform THREE-POINT JACKING (PREFERRED PROCEDURE) procedure (Ref. Chapter 07-10-00). All tires must be clear of the floor. Strut limiters will reduce strut extension (Ref. Chapter 91-00-00). (2) Perform REMOVING GROUND POWER procedure (Ref. Chapter 24-40-00). (3) Perform BATTERY DISCONNECTION procedure (Ref. Chapter 24-31-00). CAUTION: As the hydraulic hoses are disconnected, plug or cap all openings to prevent entry of foreign material into the hoses or actuator. (4) Tag the three hydraulic hoses for identification to facilitate installation. Remove the three hydraulic hoses from the unions (7) and orifice (10) (Ref. Figure 201). (5) Connect a hand pump and pressure gauge to the main landing gear hydraulic actuator’s secondary retract (main landing gear extend) 1/4 inch port (12) (Ref. Chapter 91-00-00, SPECIAL TOOLS AND EQUIPMENT). Note that when the actuator retracts, the main landing gear will extend. (6) The hydraulic actuator’s primary retract 3/8 inch port (11) and extend 3/8 inch port (13) must be open (Ref. Figure 201). (7) Slowly increase pressure to the secondary retract (main landing gear extend) 1/4 inch port (12) while observing the pressure gauge. The shuttle valve will move to close the primary retract 3/8 inch port (11). The actuator rod will retract. (8) When the rod is fully retracted, increase pressure to 650 psi. (9) With 650 psi applied to the secondary 1/4 inch port (12), maximum leakage from the primary retract 3/8 inch port (11) shall not exceed 10 drops per minute (20 drops equal one milliliter). (10) Slowly decrease pressure from the hand pump and disconnect it from the actuator. (11) If there is excessive shuttle valve leakage, the actuator must be overhauled or replaced. (12) If the actuator has been removed from the airplane, perform MAIN LANDING GEAR ACTUATOR INSTALLATION procedure in this section then proceed to Step (13) of this procedure. (13) If the unions (7) or orifice (10) were removed or loosened, install new packings (8) between the actuator and the union, or orifice. (14) Connect the three hydraulic hoses to the actuator with respect to identification tags. (15) Perform HYDRAULIC SYSTEM FILLING AND BLEEDING procedure in this section to fill and bleed the landing gear hydraulic system (Ref. 32-30-00). (16) Perform BATTERY CONNECTION procedure (Ref. Chapter 24-31-00). Page 204 Nov 1/09

32-30-10

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (17) Perform APPLYING GROUND POWER procedure (Ref. Chapter 24-40-00). CAUTION: To prevent serious damage to the pump, never operate the power pack when the engines are not running without supplying 18 psi of regulated dry air to the manual bleed valve to pressurize the reservoir. When cycling the landing gear with the power pack, do not exceed three cycles in the first six minutes of operation with a two minute (approximately) cooling period between cycles. After the first six minute of operation, a five minute cooling interval between each cycle. (18) Cycle the gear with the power pack through at least 3 complete cycles before removing the airplane from the jacks. (19) Perform REMOVING GROUND POWER procedure (Ref. Chapter 24-40-00). WARNING: Before removing the airplane from the jacks, make sure that the landing gear emergency-extend hand pump handle is in the stowed position, the plunger on the service valve is pushed down with the hinged retainer in place, the landing gear control handle is in the DOWN position, the landing gear is down and locked and the accumulator is charged to 800 ± 50 psi. (20) Perform LOWERING THE (Ref. Chapter 07-10-00).

AIRPLANE

AFTER

THREE-POINT

JACKING

32-30-10

procedure

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

3 6

4

A

5 6

A

1. BOLT 2. UPPER DRAG LEG ARM 3. BOLT 4. BUSHING 5. WASHER 6. WASHER 7. UNION 8. PACKING 9. HYDRAULIC HOSE 10. ORIFICE 11. PRIMARY RETRACT 3/8 INCH PORT 12. SECONDARY RETRACT 1/4 INCH PORT 13. EXTEND 3/8 INCH PORT

8 7 9

B 2

6

8 10 9 1 DETAIL

A

13

12

DETAIL 11

Figure 201 Main Landing Gear Actuator Assembly

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B UB32B 016530AB.A

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

3. MAIN LANDING GEAR - ACTUATOR END CAP INSPECTION WARNING: When an airplane has experienced abnormal landing gear procedures of any type, as a safety precaution, place the airplane on jacks prior to performing any inspection or maintenance. Ensure that all three landing gears are down and locked prior to removing the airplane from jacks. When performing maintenance on a hydraulically operated landing gear system, be aware that any movement of a hydraulic actuator cylinder can displace hydraulic fluid and cause unanticipated movement of other actuator cylinders in the landing gear retraction system. Servicing of the landing gear hydraulic accumulator can also result in unanticipated movement of an actuator. Either action can result in an unsafe, unlocked landing gear system. Therefore, place the airplane on jacks prior to performing any inspection or maintenance. Cycle the landing gear and ensure that all three landing gears are down and locked prior to removing the airplane from jacks. Any time maintenance is to be performed on the landing gear system, place the airplane on jacks. When jacking the airplane in an unsheltered area where winds in excess of 35 kts will be encountered, never jack more than one gear at a time clear of the ground. Any time the landing gear is only partially retracted during maintenance, always cycle the gear with the power pack through at least one complete cycle before removing the airplane from the jacks. For safety reasons, pull the 2-ampere control circuit breaker on the pilot's inboard subpanel and place a note on the circuit breaker panel that LANDING GEAR MAINTENANCE IS IN PROCESS during maintenance on the landing gear. The landing gear control handle must never be moved from the down-and-locked position while the airplane is on the ground. It is recommended that the area be roped off during extension or retraction of the landing gear. CAUTION: Jacking of an airplane for the purpose of landing gear operation, inspection, servicing or maintenance, should be accomplished within an enclosed building or hangar. In the interest of safety, should it become necessary to jack the airplane in the open, wind velocity in any direction and terrain variations, must be compensated for prior to jacking the airplane. The landing gear must not be cycled with the power pack if low on fluid or if the landing gear system is not properly rigged. Use the emergency extension hand pump, TK229/939 hydraulic hand pump or TK229-1/939 air-driven hydraulic pump to extend and retract the landing gear for maintenance and rigging (Ref. Chapter 91-00-00).

A. Preinspection (1) Determine the MLG actuator part number. If the part number is 114-380041-11, -13 or -15 (Manufactured by Triumph Actuation Systems or Frisby Airborne Hydraulics), proceed to the next Step. If the actuator part number is not one of those listed above, no further action is required. (2) Perform THREE POINT JACKING (PREFERRED (Ref. Chapter 07-10-00). All tires must be clear of the floor.

PROCEDURE)

procedure

(3) Perform REMOVING GROUND POWER procedure (Ref. Chapter 24-40-00).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (4) Perform BATTERY DISCONNECTION procedure (Ref. Chapter 24-31-00). (5) The actuator(s) may be removed from the airplane for the ultrasonic inspection. If an on-plane inspection is to be performed, Steps (5) (a) thru (5) (d) are recommended. If removal of the actuator is required perform the MAIN LANDING GEAR ACTUATOR REMOVAL procedure in this section. (a) Disconnect the actuator down position switch wiring at the receptacle plug, located in the upper rear of the wheel well. (b) Remove the two bolts that attach the actuator to the support structure. To maintain alignment upon installation, note the position of the washers between each leg of the actuator and the structure (Ref. Figure 204). (c) Remove the two bolts that attach the upper drag leg arm to the upper drag legs. (d) Move the upper drag leg arm down and position the actuator for inspection. NOTE: This Inspection procedure specifies the requirements and instructions for ultrasonic angle beam inspection of the Main Landing Gear (MLG) actuator assembly for cracks propagating from the inside radius of the end cap. Personnel shall be qualified and certified at a minimum level II in accordance with NAS 410. Equipment GE USN 60, Olympus Sonic 1200 ultrasonic instrument with A-scan display. The ultrasonic unit shall be capable of meeting performance characteristics as described in ASTM E-317 and AMS-STD-2154 for Horizontal Limit and Linearity; Vertical Limit and Linearity; Attenuator/Decade Switch Accuracy; Sensitivity and Noise; Resolution - Entry Surface and Back Surface. 45° (steel) 5MHz angle beam search unit shall be used, Panametrics P/N A5020 or equivalent. The search unit’s angle beam wedge or casing (if transducer and angled wedge are built as one unit) shall be no larger than 0.4 x 0.25 inch in the “X” and “Y” dimensions. Ultrasonic Couplant, Exosen 20, Mfg. Krautkramer, Lewiston, PA or equivalent. Aluminum IIW block Type I. (6) Clean the surfaces to be inspected using a shop cloth dampened with solvent (24, 30 or 54 Table 1, Chapter 91-00-00). (7) Ensure the inspection surface is free from grease, oil, sealer, loose or flaking paint or any other substance that would prohibit the coupling of the search unit to the part to be inspected.

B. Equipment Setup/Standardization NOTE: Periodically during the inspection and following the completion of all inspections, verify the standardization by using the calibration standard (IIW block) to insure the instrument remains within the calibration limits. The time between standardization shall not exceed 20 minutes. If the original calibration requirements are not met, all inspections performed since the last successful calibration shall be re-inspected.

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32-30-10

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL The following instrument settings are for a typical A-scan presentation ultrasonic instrument and are meant as a guide; however, sensitivity requirements shall be met. (1) Connect the probe to the cable and the cable to the instrument and turn the instrument on. (2) Set the instrument to the initial settings in Table 201 below. Table 201 Initial Instrument Settings Description

Settings

Gain

40 dB (Decibels)

Range

1.25 inches

Delay

0 in/μs (Micro Seconds)

Velocity

0.1230 in/μs (Micro Seconds)

Pulse

50 ns (Nano Seconds)

Damp

50 ohms

Mode

Single

Gate 1

Positive

Gate Position

0.475 inch

Gate Width

0.304 inch

Gate Amplitude

30% FSH

Display

Full Wave

Frequency

2-25 MHz BB (Broad Band)

Reject

Off

(3) Apply ultrasonic couplant and couple the transducer to the IIW block and adjust the gain to achieve an 80% Full Screen Height (FSH) signal from the 0.060 inch diameter Sensitivity Hole then add 6 dB. (4) Adjust the horizontal sweep to position the signal from the 0.060 inch diameter hole at 8.0 on the horizontal baseline. (5) Adjust the gate controls to position the gate at 30% FSH at 4.0 to 6.0 on the horizontal baseline.

C. Inspection (1) Apply couplant to the inspection zone of the part. Couple the transducer to the part with the sound beam directed at the end of the actuator cap as shown in Figures 205, 206 and 207. (2) Scan 180° of the circumference of the part in the area of interest shown in Figures 205, 206 and 207. Index no greater than 0.10 inch for 100% coverage of the inspection zone. Use a typical “Z” scan path while angling the transducer to maximize any crack response received between 4.0 to 6.0 on the horizontal baseline.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (3) As the transducer is scanned across the inspection zone toward the end cap of the actuator, the reflected signal from a crack on the third leg of the “V” path will first appear at 10.0 on the horizontal baseline and peak at 9.0 on the horizontal baseline. As the transducer is scanned toward the end of the actuator cap, the reflected signal from the crack on the first leg of the “V” path will appear at 6.0 on the horizontal baseline and peak at 5.0 on the horizontal baseline. Inspections shall be performed on the first leg of the “V” path as not to confuse the non-relevant signal at 8.0 on the horizontal baseline as a relevant crack signal (Ref. Figure 202). (4) Confirm instrument calibration at completion of inspection. If the original calibration requirements are not met, All Inspections performed since the last successful calibration shall be re-inspected.

D. Indication Evaluation Criteria (1) Any indication between 4.0 and 6.0 on the horizontal baseline shall be re-scanned to determine if the indication is false, caused by excessive couplant, part geometry, surface condition, or a defect. (2) Any repeatable crack response of 30% FSH or greater within the gated zone shall result in rejection of part.

E. Reporting (1) Mark cracks and report location and size to Hawker Beechcraft Corporation Airline Technical Support.

F. Post Inspection (1) Clean ultrasonic couplant from the actuator prior to installation. (2) If any cracks are found, replace with serviceable actuator(s). (3) Check rod end nut for security. If loose, remove safety wire, tighten nut and install new safety wire. (4) If an on-plane inspection was performed, complete the following Steps: (a) Move the upper drag leg arm up and position the actuator (Ref. Figure 204). (b) Install the two bolts that attach the upper drag leg arm to the upper drag legs. (c) Install the two bolts that attach the actuator to the support structure with washers positioned as noted during removal. (d) Connect the actuator down position switch wiring at the receptacle plug, located in the upper rear of the wheel well. (5) If actuator was removed from the airplane, perform the MAIN LANDING GEAR ACTUATOR INSTALLATION procedure in this section. (6) Perform BATTERY CONNECTION procedure (Ref. Chapter 24-31-00). (7) Perform APPLYING GROUND POWER procedure (Ref. Chapter 24-40-00). (8) Perform HYDRAULIC SYSTEM FILLING AND BLEEDING procedure to fill and bleed the landing gear hydraulic system (Ref. 32-30-00).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (9) Perform LOWERING THE (Ref. Chapter 07-10-00).

AIRPLANE

AFTER

THREE-POINT

JACKING

procedure

Figure 202 Non-relevant Indication

32-30-10

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 203 Relevant Indication Peaked at 5.0 on the Horizontal Baseline

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 204 Main Landing Gear Actuator

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 205 Couple Transducer

Figure 206 MLG Actuator Assembly Inspection Area

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 207 Scan Plan

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LANDING GEAR MAIN LANDING GEAR ACTUATOR ORIFICE MAINTENANCE PRACTICES

32-30-11 200200

1. PROCEDURES WARNING: When an airplane has experienced abnormal landing gear procedures of any type, as a safety precaution, place the airplane on jacks prior to performing any inspection or maintenance. Ensure that all three landing gears are down and locked prior to removing the airplane from jacks. When performing maintenance on a hydraulically operated landing gear system, be aware that any movement of a hydraulic actuator cylinder can displace hydraulic fluid and cause unanticipated movement of other actuator cylinders in the landing gear retraction system. Servicing of the landing gear hydraulic accumulator can also result in unanticipated movement of an actuator. Either action can result in an unsafe, unlocked landing gear system. Therefore, place the airplane on jacks prior to performing any inspection or maintenance. Cycle the landing gear and ensure that all three landing gears are down and locked prior to removing the airplane from jacks. Any time maintenance is to be performed on the landing gear system, place the airplane on jacks. When jacking the airplane in an unsheltered area where winds in excess of 35 kts will be encountered, never jack more than one gear at a time clear of the ground. Any time the landing gear is only partially retracted during maintenance, always cycle the gear with the power pack through at least one complete cycle before removing the airplane from the jacks. For safety reasons, pull the 2-ampere control circuit breaker on the pilot's inboard subpanel and place a note on the circuit breaker panel that LANDING GEAR MAINTENANCE IS IN PROCESS during maintenance on the landing gear. The landing gear control handle must never be moved from the down-and-locked position while the airplane is on the ground. It is recommended that the area be roped off during extension or retraction of the landing gear. CAUTION: Jacking of an airplane for the purpose of landing gear operation, inspection, servicing or maintenance, should be accomplished within an enclosed building or hangar. In the interest of safety, should it become necessary to jack the airplane in the open, wind velocity in any direction and terrain variations, must be compensated for prior to jacking the airplane. The landing gear must not be cycled with the power pack if low on fluid or if the landing gear system is not properly rigged. Use the emergency extension hand pump, TK229/939 hydraulic hand pump or TK229-1/939 air-driven hydraulic pump to extend and retract the landing gear for maintenance and rigging (Ref. Chapter 91-00-00).

A. Removal (1) Perform the THREE-POINT JACKING (PREFERRED PROCEDURE) procedure (Ref. Chapter 07-10-00). All tires must be clear of the floor. Strut limiters will reduce strut extension (Ref. Chapter 91-00-00).

32-30-11

Page 201 Nov 1/09

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (2) Perform REMOVING GROUND POWER procedure (Ref. Chapter 24-40-00). (3) Perform BATTERY DISCONNECTION procedure (Ref. Chapter 24-31-00). NOTE: The actuator orifice (10) is located in the port along side the actuator down-position switch (the orifice fitting closest to the actuator-to-keel attach point) (Ref. Figure 201). (4) Disconnect the hydraulic line (9) from the actuator fitting. Cap the open hydraulic line. NOTE: The actuator orifice valve will have pressure on the hydraulic line fitting and care must be taken to prevent loss of the poppet and spring. (5) Remove the fitting from the actuator port.

B. Installation (1) Install a new packing (8) on the hydraulic line fitting (Ref. Figure 201). (2) Insert a pin punch, slightly smaller than the orifice poppet hole, through the flared end of the hydraulic line fitting. (3) Install the orifice poppet on the end of the pin punch. (4) Install the orifice spring in the port on the actuator housing. CAUTION: The orifice poppet must be kept in the same alignment as the actuator port or damage to the poppet requires replacement of the actuator housing. (5) Install the orifice poppet and fitting in the port on the actuator. (6) Using a pin punch slightly larger than the orifice poppet hole, press lightly on the orifice poppet. The orifice poppet must have free movement. WARNING: Before removing the airplane from the jacks, make sure that the landing gear emergency-extend hand pump handle is in the stowed position, the plunger on the service valve is pushed down with the hinged retainer in place, the landing gear control handle is in the DOWN position, the landing gear is down and locked and the accumulator is charged to 800 ± 50 psi. (7) Perform LOWERING THE (Ref. Chapter 07-10-00).

AIRPLANE

AFTER

THREE-POINT

JACKING

procedure

C. Cleaning Clean the main gear actuator orifice (10) as follows (Ref. Figure 201): (1) Clean the actuator orifice poppet and spring with solvent (2, Table 1, Chapter 91-00-00). (2) Blow dry using clean shop air. (3) Perform MAIN LANDING GEAR ACTUATOR ORIFICE INSTALLATION procedure above.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

3 6

4

A

5 6

A

1. BOLT 2. UPPER DRAG LEG ARM 3. BOLT 4. BUSHING 5. WASHER 6. WASHER 7. UNION 8. PACKING 9. HYDRAULIC HOSE 10. ORIFICE 11. PRIMARY RETRACT 3/8 INCH PORT 12. SECONDARY RETRACT 1/4 INCH PORT 13. EXTEND 3/8 INCH PORT

8 7 9

B 2

6

8 10 9 1 DETAIL

A

13

12

DETAIL 11

B UB32B 016530AB.A

Figure 201 Main Landing Gear Actuator

32-30-11

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

LANDING GEAR MAIN LANDING GEAR DOORS MAINTENANCE PRACTICES

32-30-12 200200

1. PROCEDURES WARNING: When an airplane has experienced abnormal landing gear procedures of any type, as a safety precaution, place the airplane on jacks prior to performing any inspection or maintenance. Ensure that all three landing gears are down and locked prior to removing the airplane from jacks. When performing maintenance on a hydraulically operated landing gear system, be aware that any movement of a hydraulic actuator cylinder can displace hydraulic fluid and cause unanticipated movement of other actuator cylinders in the landing gear retraction system. Servicing of the landing gear hydraulic accumulator can also result in unanticipated movement of an actuator. Either action can result in an unsafe, unlocked landing gear system. Therefore, place the airplane on jacks prior to performing any inspection or maintenance. Cycle the landing gear and ensure that all three landing gears are down and locked prior to removing the airplane from jacks. Any time maintenance is to be performed on the landing gear system, place the airplane on jacks. When jacking the airplane in an unsheltered area where winds in excess of 35 kts will be encountered, never jack more than one gear at a time clear of the ground. Any time the landing gear is only partially retracted during maintenance, always cycle the gear with the power pack through at least one complete cycle before removing the airplane from the jacks. For safety reasons, pull the 2-ampere control circuit breaker on the pilot's inboard subpanel and place a note on the circuit breaker panel that LANDING GEAR MAINTENANCE IS IN PROCESS during maintenance on the landing gear. The landing gear control handle must never be moved from the down-and-locked position while the airplane is on the ground. It is recommended that the area be roped off during extension or retraction of the landing gear. CAUTION: Jacking of an airplane for the purpose of landing gear operation, inspection, servicing or maintenance, should be accomplished within an enclosed building or hangar. In the interest of safety, should it become necessary to jack the airplane in the open, wind velocity in any direction and terrain variations, must be compensated for prior to jacking the airplane. The landing gear must not be cycled with the power pack if low on fluid or if the landing gear system is not properly rigged. Use the emergency extension hand pump, TK229/939 hydraulic hand pump or TK229-1/939 air-driven hydraulic pump to extend and retract the landing gear for maintenance and rigging (Ref. Chapter 91-00-00).

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A. Rigging Prior to rigging the main gear doors, the landing gear must be properly rigged. Refer to MAIN LANDING GEAR RIGGING procedure. While rigging the doors, use the emergency extension hand pump to raise and lower the landing gear (Ref. 32-30-00). The following adjustments are to be done in the sequence as given. Any adjustment after Step (7) will require the return to Step (7) and repeat the subsequent Steps. (1) Perform the THREE-POINT JACKING (PREFERRED PROCEDURE) procedure (Ref. Chapter 07-10-00). All tires must be clear of the floor. Strut limiters will reduce strut extension (Ref. Chapter 91-00-00). (2) Perform APPLYING GROUND POWER procedure (Ref. Chapter 24-40-00). (3) Disconnect the door actuating cams from the doors by removing the cotter pin, nut, washer, and bolt from the upper end of the link. Position the doors so the landing gear will clear during retraction. Check the door cams and the rollers on the landing gear to ensure free movement (Ref. Figure 201). (4) Position the adjusting blocks so the radius of the adjusting block approximately matches the radius of the cam lower finger and secure with the attaching bolts. (5) Partially retract the landing gear until the rollers are positioned next to the lower finger of the actuating cams. Adjust the actuating cam stop bolt to a clearance of 0.02 to 0.06 inch between the actuating cam lower finger and the roller. (6) Retract the landing gear. Push the doors up by hand and check for fit and clearance. (7) Working with one door at a time, manually close the door and adjust the link assembly as required so the door closes snugly, front and rear. After adjustment is complete, tighten the jam nut on the link. (8) After both doors are properly adjusted, extend the landing gear and connect the links to the actuating cams with the bolts, washers, nuts, and cotter pins. (9) Place the upper edge of a straight edge at the centerline of the actuating cam mounting bolt and the lower edge at the centerline of the lower link attaching bolt (Ref. Figure 201). Check the linkage for overcenter of 0.03 to 0.50 inch. If the overcenter is within this range no further adjustments are necessary; proceed to Step (11). If the overcenter dimension is not within the range, then perform Step (10). (10) If the overcenter dimension was not 0.03 to 0.50 inch, adjust the actuating cam stop block to achieve a dimension within the range. Partially retract the landing gear until the rollers are positioned next to the lower finger of the actuating cams. Position the adjusting block to have a clearance of 0.02 to 0.06 inch between the adjusting block and the roller during the retraction cycle. Extend the landing gear. (11) Check final torque on all attaching hardware. Apply a force of approximately 25 pounds at the lower center edge of the door. The door linkage must remain overcenter and the door must remain open. Rig any door that folds under this 25 pound load.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (12) Cycle the landing gear several times. Check that the doors shut snugly, front and rear. During the landing gear extension, check the door overcenter mechanism by pushing on the center of the door toward the gear with a force of approximately 25 pounds. The doors must open and stay in the locked position. Rig any door that folds under this 25 pound load. (13) Lubricate the door hinges and the actuating mechanisms with the proper lubricant. Refer to MAIN LANDING GEAR LUBRICATION procedure (Ref. Chapter 12-20-00). CAUTION: When cycling the landing gear with the power pack, do not exceed three cycles in the first six minutes of operation; then allow a two minute cooling period and resume operation of the power pack, allowing five minutes between cycles. To prevent serious damage to the pump, never operate the power pack when the engines are not operating without supplying 18 psi of regulated dry air to pressurize the power pack reservoir in place of the engine bleed air. (14) Cycle the landing gear several times with the power pack, checking each time that the doors close snugly when the landing gear is retracted and that they stay open when the landing gear is extended. WARNING: Before removing the airplane from the jacks, make sure that the landing gear emergency extend hand pump handle is in the stowed position, the plunger on the service valve is pushed down with the hinged retainer in place, the landing gear control handle is in the DOWN position, the landing gear is down and locked and the accumulator is charged to 800 ± 50 psi. (15) Perform REMOVING GROUND POWER procedure (Ref. Chapter 24-40-00). (16) Perform the LOWERING THE AIRPLANE AFTER THREE-POINT JACKING procedure (Ref. Chapter 07-10-00).

B. Hinge Wear Limits Use the following parameters for the main gear door hinge to determine when the hinge is to be replaced. (1) Minimum hinge wall thickness is 0.048 inch at any point. (2) Maximum hinge opening where the hinge pin fits is 0.060 inch at any point.

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Figure 201 Main Landing Gear Door Rigging

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LANDING GEAR NOSE LANDING GEAR MAINTENANCE PRACTICES

32-30-13 200200

1. PROCEDURES WARNING: When an airplane has experienced abnormal landing gear procedures of any type, as a safety precaution, place the airplane on jacks prior to performing any inspection or maintenance. Ensure that all three landing gears are down and locked prior to removing the airplane from jacks. When performing maintenance on a hydraulically operated landing gear system, be aware that any movement of a hydraulic actuator cylinder can displace hydraulic fluid and cause unanticipated movement of other actuator cylinders in the landing gear retraction system. Servicing of the landing gear hydraulic accumulator can also result in unanticipated movement of an actuator. Either action can result in an unsafe, unlocked landing gear system. Therefore, place the airplane on jacks prior to performing any inspection or maintenance. Cycle the landing gear and ensure that all three landing gears are down and locked prior to removing the airplane from jacks. Any time maintenance is to be performed on the landing gear system, place the airplane on jacks. When jacking the airplane in an unsheltered area where winds in excess of 35 kts will be encountered, never jack more than one gear at a time clear of the ground. Any time the landing gear is only partially retracted during maintenance, always cycle the gear with the power pack through at least one complete cycle before removing the airplane from the jacks. For safety reasons, pull the 2-ampere control circuit breaker on the pilot's inboard subpanel and place a note on the circuit breaker panel that LANDING GEAR MAINTENANCE IS IN PROCESS during maintenance on the landing gear. The landing gear control handle must never be moved from the down-and-locked position while the airplane is on the ground. It is recommended that the area be roped off during extension or retraction of the landing gear. CAUTION: Jacking of an airplane for the purpose of landing gear operation, inspection, servicing or maintenance, should be accomplished within an enclosed building or hangar. In the interest of safety, should it become necessary to jack the airplane in the open, wind velocity in any direction and terrain variations, must be compensated for prior to jacking the airplane. The landing gear must not be cycled with the power pack if low on fluid or if the landing gear system is not properly rigged. Use the emergency extension hand pump, TK229/939 hydraulic hand pump or TK229-1/939 air-driven hydraulic pump to extend and retract the landing gear for maintenance and rigging (Ref. Chapter 91-00-00).

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A. Rigging A service valve in the left hand wing, adjacent to the power pack, is provided for use with the emergency extension hand pump to raise and lower the landing gear while the landing gear is being rigged. The service valve is normally in a mode to extend the landing gear (red knob pushed down). The service valve is accessible through a door in the wing skin just inboard of the LH nacelle. The red knob cannot be pulled up until the hinged retainer covering it is unlatched and moved out of the way. The landing gear must not be cycled electrically until the landing gear is properly rigged. With the airplane on jacks and an external power supply connected and adjusted to 28 ± 0.25 volts, the landing gear is retracted by pulling the red knob on the service valve up and then operating the emergency extension hand pump. When the red knob is pulled up, the two switches mounted on the service valve are actuated to open the circuit to the power pack motor and to complete a circuit from the 2-ampere control circuit breaker to the retract solenoid on the gear selector valve. The same action positions the service valve to route the hand pump pressure fluid to the normal retract mode plumbing to the actuators. Remove the pump handle from the securing clip and pump the handle up and down to retract the landing gear. After the landing gear is retracted, position the emergency extend hand handle in the stowed position and turn off the electrical power. To extend the landing gear push the red knob on the service valve down and extend the landing gear using the emergency extension hand pump; the electrical power does not have to be on to extend the landing gear. For safety reasons, pull the 2-ampere control circuit breaker on the pilot's inboard subpanel. CAUTION: If the red knob on the service valve is pushed down while the landing gear is retracted, the electrical power is on and the landing gear control handle is in the down position, the landing gear will extend immediately. To prevent serious damage to the pump, never operate the power pack when the engines are not operating without supplying 18 psi of regulated dry air to pressurize the power pack reservoir in place of the engine bleed air. When cycling the landing gear with the power pack, do not exceed three cycles in the first six minutes of operation with approximately a two-minute cooling period between cycles, then with a five-minute cooling interval between each cycle. Before removing the airplane from the jacks, make sure that the emergency extension hand pump handle is in the stowed position and that the red knob on the service valve is pushed down with the hinged retainer in place. Retract and extend the landing gear, using the hydraulic power pack to ensure the landing gear is in the down-and-locked position before removing the airplane from the jacks. CAUTION: Do not fully cycle the landing gear electrically until the gear is properly rigged. Use of the emergency hand pump is recommended for initial rigging. (1) Perform the THREE-POINT JACKING (PREFERRED PROCEDURE) procedure (Ref. Chapter 07-10-00). All tires must be clear of the floor. Strut limiters will reduce strut extension (Ref. Chapter 91-00-00). (2) Perform APPLYING GROUND POWER procedure (Ref. Chapter 24-40-00). (3) Disconnect the push-pull links that operate the nose gear door from the strut by removing the nuts, washers, and bolts that attach the rod ends to the strut. NOTE: Use the emergency extension hand pump for all extensions and retractions during this procedure. (4) Retract the gear slightly to take the load off the drag brace and to unlock the actuator.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (5) Remove the nut, washer, and bolt that attach the actuator clevis to the yoke on the drag brace assembly (Ref. Figure 201). NOTE: Verify continuity of the actuator switch by checking continuity between pins A and E then H and K of connector P161. (6) With the actuator fully extended and locked and the lower drag leg arm against the stop bolt, loosen the jam nut on the actuator clevis and adjust the clevis so that the clevis hole aligns with the hole in the yoke. The bolt should install freely in the yoke. Lengthen the clevis to the next keying position and tighten the jam nut. Attach the actuator clevis to the yoke with the bolt, washer, and nut. (7) To eliminate any side load on the actuator, the alignment of the yoke assembly to the drag brace assembly must be checked. Check for proper alignment as follows: NOTE: The yoke is offset to facilitate alignment and may be installed with the short lug on either side. If it is necessary to rotate the yoke for proper alignment, the yoke and clevis may be rotated as an assembly by turning the actuator piston rod. Do not change the clevis adjustment set in Step (6). (a) Remove the cotter pin, nut, washer, bushings, and stop bolt that attach the yoke to the upper drag leg. (b) With the actuator fully extended, position the upper drag leg to install with the yoke. Measure the space between the yoke and the upper drag leg on each side of the yoke. (c) Retract the actuator to its fully retracted position. Manually push the nose gear to the retracted position and locate the yoke assembly to install with the upper drag leg. Measure the space between the yoke and the upper drag leg on each side of the yoke. (d) Manually pull the drag brace to the extended position and extend the actuator. Install AN960-616 and/or AN960-616L washers, as required, on each side of the yoke to obtain proper alignment as determined in Steps (7) (b) and (7) (c). Install the stop bolt, bushing, washer, and nut. A minimum of one AN960-616 or AN960-616L washer must be installed on each side of the yoke. (8) With the landing gear fully extended, apply moderate pressure to the drag brace assembly to bottom out any end play in the actuator. Check for clearance between the lower drag leg and both stops. If a clearance of 0.005 +0.005/-0.004 inch does not exist, loosen the clevis jam nut and lengthen the clevis to the next keying position by rotating the actuator piston rod. Tighten the jam nut and safety wire. (9) With the landing gear fully retracted, measure the distance from the upper wheel well panel to the O.D. of the nose gear axle. If this measurement is not 11.09 to 11.69 inches, adjust the actuator stroke as outlined in Step (10) (Ref. Figure 202). (10) The stroke length of the actuator (9.79 ± 0.12 inches) is preset by the supplier and should not require adjustment. If the clearance measured in Step (9) is not within tolerance, remove the actuator from the airplane and adjust the actuator stroke length as follows: NOTE: To ensure proper alignment of the hydraulic plumbing to the shuttle valve located in the end cap, the end cap must be rotated in 360° rotations only. If more than two complete rotations are necessary to obtain the proper stroke, structural damage may be present and the airplane structure must be inspected.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (a) Remove the safety wire and loosen the end cap jam nut. Rotate the end cap one full turn at a time until the proper adjustment is obtained. When the end cap is rotated out, the stroke length is increased, and when rotated in, the stroke length is decreased. Each full turn of the end cap will change the stroke length by 0.06 inch. (b) After the stroke length is properly adjusted, tighten the jam nut and safety wire. (c) Perform the LANDING GEAR HYDRAULIC SYSTEM FILLING AND BLEEDING procedure (Ref. 32-30-00). (11) Lubricate the landing gear and drag brace hinge points with the proper lubricant. Refer to NOSE LANDING GEAR LUBRICATION procedure (Ref. Chapter 12-20-00). CAUTION: When cycling the landing gear with the power pack, do not exceed three cycles in the first six minutes of operation with approximately a two minute cooling period between cycles, then with a five minute cooling interval between each cycle. (12) Cycle the landing gear and check for proper operation of the in-transit and gear-down lights. If necessary, adjust the downlock, up-position, and actuator downlock switches (Ref. 32-60-00). (13) Connect the nose gear door to the strut. If necessary rig the door. Refer to NOSE GEAR DOOR RIGGING procedure (Ref. 32-30-15). WARNING: Before removing the airplane from the jacks, make sure that the landing gear emergency extend hand pump handle is in the stowed position, the plunger on the service valve is pushed down with the hinged retainer in place, the landing gear control handle is in the DOWN position, the landing gear is down and locked and the accumulator is charged to 800 ± 50 psi. (14) Perform REMOVING GROUND POWER procedure (Ref. Chapter 24-40-00). (15) Perform LOWERING THE (Ref. Chapter 07-10-00).

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AIRPLANE

AFTER

THREE-POINT

JACKING

procedure

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 201 Nose Landing Gear Rigging

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Figure 202 Nose Landing Gear Clearance Check

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

LANDING GEAR NOSE LANDING GEAR ACTUATOR MAINTENANCE PRACTICES

32-30-14 200200

1. PROCEDURES WARNING: When an airplane has experienced abnormal landing gear procedures of any type, as a safety precaution, place the airplane on jacks prior to performing any inspection or maintenance. Ensure that all three landing gears are down and locked prior to removing the airplane from jacks. When performing maintenance on a hydraulically operated landing gear system, be aware that any movement of a hydraulic actuator cylinder can displace hydraulic fluid and cause unanticipated movement of other actuator cylinders in the landing gear retraction system. Servicing of the landing gear hydraulic accumulator can also result in unanticipated movement of an actuator. Either action can result in an unsafe, unlocked landing gear system. Therefore, place the airplane on jacks prior to performing any inspection or maintenance. Cycle the landing gear and ensure that all three landing gears are down and locked prior to removing the airplane from jacks. Any time maintenance is to be performed on the landing gear system, place the airplane on jacks. When jacking the airplane in an unsheltered area where winds in excess of 35 kts will be encountered, never jack more than one gear at a time clear of the ground. Any time the landing gear is only partially retracted during maintenance, always cycle the gear with the power pack through at least one complete cycle before removing the airplane from the jacks. For safety reasons, pull the 2-ampere control circuit breaker on the pilot's inboard subpanel and place a note on the circuit breaker panel that LANDING GEAR MAINTENANCE IS IN PROCESS during maintenance on the landing gear. The landing gear control handle must never be moved from the down-and-locked position while the airplane is on the ground. It is recommended that the area be roped off during extension or retraction of the landing gear. CAUTION: Jacking of an airplane for the purpose of landing gear operation, inspection, servicing or maintenance, should be accomplished within an enclosed building or hangar. In the interest of safety, should it become necessary to jack the airplane in the open, wind velocity in any direction and terrain variations, must be compensated for prior to jacking the airplane. The landing gear must not be cycled with the power pack if low on fluid or if the landing gear system is not properly rigged. Use the emergency extension hand pump, TK229/939 hydraulic hand pump or TK229-1/939 air-driven hydraulic pump to extend and retract the landing gear for maintenance and rigging (Ref. Chapter 91-00-00).

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A. Removal (1) Perform the THREE-POINT JACKING (PREFERRED PROCEDURE) procedure (Ref. Chapter 07-10-00). All tires must be clear of the floor. Strut limiters will reduce strut extension (Ref. Chapter 91-00-00). (2) Perform REMOVING GROUND POWER procedure (Ref. Chapter 24-40-00). (3) Perform BATTERY DISCONNECTION procedure (Ref. Chapter 24-31-00). NOTE: If additional working room is needed in the wheel well, Perform the NOSE LANDING GEAR DRAG BRACE REMOVAL procedures (Ref. 32-20-00). (4) Disconnect the actuator downlock switch receptacle plug located in the LH keel of the nose wheel well. CAUTION: As the hydraulic hoses are disconnected, plug or cap all openings to prevent entry of foreign material into the hoses or actuator. (5) Working through the access panel just aft of the nose cone and above the actuator, remove the two hydraulic hoses (14) from the actuator. Identify the hoses to facilitate installation. Disconnect the hydraulic line (15) from the swivel fitting (10) in the actuator trunnion. Remove safety wire and the swivel fitting from the actuator (Ref. Figure 201). (6) Remove each actuator support bracket (1) as follows: (a) Working inside the avionics/baggage compartments, remove the five bolts (2) attaching the aft end of the bracket. (b) Remove the row of five bolts (3) attaching the bracket to the actuator support (4) and support plate (5). (c) Remove the bolt (6) securing the forward end of the bracket and remove the bracket from the airplane. (7) Remove the three bolts attaching the actuator support bearing assemblies (7) to the actuator supports and withdraw the bearing assemblies from the actuator trunnions. The shims (8) installed between the bearing assemblies and the actuator supports should be retained in the positions from which they were removed. (8) Support the actuator and remove the four remaining bolts (9) securing the support plates (5) to the actuator supports (4). The support plates and actuator will now be free to lower out of the wheel well.

B. Installation (1) Place one of the support plates (5) over each actuator trunnion. Position the actuator and support plates between the actuator supports (4) and secure with the four bolts (9) (Ref. Figure 201). (2) Install each actuator support bracket (1) as follows: (a) Position the bracket and install the bolt (6) to secure the forward end of the bracket.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (b) Install the row of five bolts (3) attaching the bracket to the actuator support (4) and support plate (5). (c) Working inside the avionics/baggage compartments, install the five bolts (2) attaching the aft end of the bracket. NOTE: Inspect the actuator support bearing assemblies (7) for corrosion or damage and ensure free rotation. (3) Install the shims (8) over the actuator trunnions in the positions from which they were removed. Install the actuator support bearing assemblies (7) on the actuator trunnions and secure to the actuator supports with the bolts. (4) Manually push the landing gear to the retracted position and check that the actuator clevis is centered with the nose landing gear assembly. If it is not centered, add or remove the shims (8) between the actuator supports and the support bearing assemblies to correct the misalignment. The allowable end play of the actuator with respect to the support bearing assemblies is 0.005 to 0.040 inch. (5) Using a new packing (11), install the swivel fitting (10) in the actuator and safety wire. Connect the hydraulic line (15) to the swivel fitting (10). (6) Working through the access panel just aft of the nose cone and above the actuator, connect the two hydraulic hoses (14) to the actuator. If the unions (12 and 16) between the actuator and hose ends were removed or loosened, install new packings (13). (7) Connect the actuator downlock switch wiring in the LH keel of the nose wheel well. (8) Install the drag brace assembly (Ref. 32-20-00, DRAG BRACE INSTALLATION). NOTE: If the shims between the actuator supports and the support bearing assemblies were moved, if the actuator stroke length or rod end length has been changed, or if the actuator being installed is not the one that was removed, rig the nose landing gear (Ref. 32-30-13, NOSE LANDING GEAR RIGGING). (9) Connect the actuator clevis to the yoke with the bolt, nut and washers. (10) Perform the HYDRAULIC SYSTEM FILLING AND BLEEDING procedure to fill and bleed the landing gear hydraulic system (Ref. 32-30-00). (11) Perform BATTERY CONNECTION procedure (Ref. Chapter 24-31-00). (12) Perform APPLYING GROUND POWER procedure (Ref. Chapter 24-40-00). CAUTION: To prevent serious damage to the pump, never operate the power pack when the engines are not running without supplying 18 psi of regulated dry air to pressurize the power pack reservoir in place of the engine bleed air. When cycling the landing gear with the power pack, do not exceed three cycles in the first six minutes of operation with a two minute (approximately) cooling period between cycles. After the first six minute of operation, a five minute cooling interval between each cycle.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (13) Cycle the landing gear with the power pack, and check for proper operation of the in-transit and gear down lights. If necessary, adjust the downlock, up position, and actuator downlock switches (Ref. 32-60-00). (14) Before removing the airplane from the jacks, make sure the emergency extension hand pump handle is in the stowed position and that the red knob on the service valve is pushed down with the hinged retainer in place. Retract and extend the landing gear using the hydraulic power pack to ensure the landing gear is in the down and locked position before removing the airplane from the jacks. (15) Perform REMOVING GROUND POWER procedure (Ref. Chapter 24-40-00). (16) Perform the LOWERING THE AIRPLANE AFTER THREE-POINT JACKING procedure (Ref. Chapter 07-10-00).

C. Shuttle Valve Functional Test This procedure may be performed with the actuator on the bench or installed in the airplane. If the procedure is to be performed with the actuator on the bench, perform the NOSE LANDING GEAR ACTUATOR REMOVAL procedure in this section then proceed to Step (5) of this procedure. (1) Perform the THREE-POINT JACKING (PREFERRED PROCEDURE) procedure (Ref. Chapter 07-10-00). All tires must be clear of the floor. Strut limiters will reduce strut extension (Ref. Chapter 91-00-00). (2) Perform REMOVING GROUND POWER procedure (Ref. Chapter 24-40-00). (3) Perform BATTERY DISCONNECTION procedure (Ref. Chapter 24-31-00). CAUTION: As the hydraulic hoses and hydraulic line are disconnected, plug or cap all openings to prevent entry of foreign material into the hoses, line or actuator. (4) Disconnect the hydraulic line (15) from the swivel fitting (10) (Ref. Figure 201). (5) Working through the access panel just aft of the nose cone and above the actuator, tag the two hydraulic hoses to facilitate installation (Ref. Chapter 6-50-00, FUSELAGE ACCESS PANELS). Remove the two hydraulic hoses from the unions (12 and 16). (6) Connect a hand pump and pressure gauge to the nose landing gear hydraulic actuator’s secondary extend 1/4 inch port (17) (Ref. Chapter 91-00-00, SPECIAL TOOLS AND EQUIPMENT). (7) The hydraulic actuator’s primary extend 3/8 inch port (18) and retract 3/8 inch port (19) must be open (Ref. Figure 201). (8) Slowly increase pressure to the secondary extend 1/4 inch port (17) while observing the pressure gauge. The shuttle valve will move to close the primary extend 3/8 inch port (18). The actuator rod will extend. (9) When the rod is fully extended, increase pressure to 650 psi. (10) With 650 psi applied to the secondary 1/4 inch port (17), maximum leakage from the primary extend port (18) shall not exceed 10 drops per minute (20 drops equal one milliliter). (11) Slowly decrease pressure from the hand pump and disconnect it from the actuator. Page 204 Nov 1/13

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (12) If there is excessive shuttle valve leakage, the actuator must be overhauled or replaced. (13) If the actuator has been removed from the airplane, perform the NOSE LANDING GEAR ACTUATOR INSTALLATION procedure in this section and then proceed to Step (16). (14) If the unions (12 and 16) were removed or loosened, install new packings (13) (Ref. Figure 201). (15) If the swivel fitting (10) was removed or loosened, install a new packing (11). (16) Connect the hydraulic hoses to the actuator with respect to identification tags. (17) Safety wire the swivel fitting (10) to the actuator. (18) Perform the HYDRAULIC SYSTEM FILLING AND BLEEDING procedure to fill and bleed the landing gear hydraulic system (Ref. 32-30-00). (19) Perform BATTERY CONNECTION procedure (Ref. Chapter 24-31-00). (20) Perform APPLYING GROUND POWER procedure (Ref. Chapter 24-40-00). CAUTION: To prevent serious damage to the pump, never operate the power pack when the engines are not running without supplying 18 psi of regulated dry air to pressurize the power pack reservoir in place of the engine bleed air. When cycling the landing gear with the power pack, do not exceed three cycles in the first six minutes of operation with approximately a two minute cooling period between cycles, then with a five minute cooling interval between each cycle. (21) Cycle the landing gear with the power pack through at least 3 complete cycles and check for proper operation of the in-transit and gear down lights. If necessary, adjust the down position, up position, and actuator down position switches (Ref. 32-60-00). WARNING: Before removing the airplane from the jacks, make sure that the landing gear emergency extend hand pump handle is in the stowed position, the plunger on the service valve is pushed down with the hinged retainer in place, the landing gear control handle is in the DOWN position, the landing gear is down and locked and the accumulator is charged to 800 ± 50 psi. (22) Perform REMOVING GROUND POWER procedure (Ref. Chapter 24-40-00). (23) Perform the LOWERING THE AIRPLANE AFTER THREE-POINT JACKING procedure (Ref. Chapter 07-10-00).

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17 18

10

A

19

DETAIL

B 10

15

13 12

11 4 1. ACTUATOR SUPPORT BRACKET 2. BOLT 3. BOLT 4. ACTUATOR SUPPORT 5. ACTUATOR SUPPORT PLATE 6. BOLT 7. ACTUATOR SUPPORT BEARING ASSEMBLY 8. SHIM 9. BOLT 10. SWIVEL FITTING 11. PACKING 12. UNION 13. PACKING 14. HYDRAULIC HOSE 15. HYDRAULIC LINE 16. UNION RESTRICTOR 17. SECONDARY EXTEND 1/4 INCH PORT 18. PRIMARY EXTEND 3/8 INCH PORT 19. RETRACT 3/8 INCH PORT

16 2 14

1 8 7

B 5 6 DETAIL

A

Figure 201 Nose Landing Gear Actuator Assembly

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9

9 3

UB32B 016401AB.AI

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

D. Nose Landing Gear Actuator Ultrasonic Inspection WARNING: Any time maintenance is to be performed on the landing gear system, place the airplane on jacks. When jacking the airplane in an unsheltered area where winds in excess of 35 kts will be encountered, never jack more than one gear at a time clear of the ground. Any time the landing gear is only partially retracted during maintenance, always cycle the gear with the power pack through at least one complete cycle before removing the airplane from the jacks. For safety reasons, pull the 2-ampere control circuit breaker on the pilot's inboard subpanel and place a note on the circuit breaker panel that LANDING GEAR MAINTENANCE IS IN PROCESS during maintenance on the landing gear. The landing gear control handle must never be moved from the down-and-locked position while the airplane is on the ground. It is recommended that the area be roped off during extension or retraction of the landing gear. CAUTION: The landing gear must not be cycled with the power pack if low on fluid or if the landing gear system is not properly rigged. Use the emergency extension hand pump, TK229/939 hydraulic hand pump or TK229-1/939 air-driven hydraulic pump to extend and retract the landing gear for maintenance and rigging (Ref. Chapter 91-00-00). This inspection procedure specifies the requirements and instructions for ultrasonic angle beam inspection of the NLG actuator assembly P/N 112-380022-23 (or -3 or -15) for cracks propagating from the inside radius of the end cap. Figure 202 shows an illustration of the actuator and the area of inspection.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

INSPECTION ZONE 1

INSPECTION ZONE 2

1.3” R

1.0”

360 DEGREES

ZONE 1

COVERAGE

ZONE 2

UC32B 110086AA.AI

Figure 202 Nose Landing Gear Actuator Assembly Inspection Area

Page 208 Nov 1/13

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (1) Preparation NOTE: Personnel shall be qualified and certified minimum level II in accordance with NAS 410. Equipment:

GE USN 60, Olympus Sonic 1200 ultrasonic instrument or equivalent instrument with A-scan display. The ultrasonic unit shall be capable of meeting performance characteristics as described in ASTM E-317 and AMS-STD-2154 for Horizontal Limit and Linearity; Vertical Limit and Linearity; Attenuator/Decade Switch accuracy: Sensitivity and Noise; Resolution - Entry Surface and Back Surface. 45 degree (steel) 5 MHz angle beam search unit shall be used, Panametrics P/N A5020 or equivalent, the search unit’s angle beam wedge or casing. If transducer and angled wedge are built as one unit, it shall be no larger than 0.4 x 0.25 inch in the “X” and “Y” dimensions. Couplant Exosen 20 Mfg. Krautkramer, Lewiston PA or equivalent. Aluminum IIW block

Clean the surfaces to be inspected using a shop rag dampened with solvent (24, 30 or 54, Table 1, Chapter 91-00-00). Ensure the inspection surface is free from grease, oil, sealer, loose or flaking paint, or any other substance that would prohibit the coupling of the search unit to the part to be inspected. (2) Equipment Setup/Standardization Periodically during the inspection and following the completion of all inspections, verify the standardization by using the calibration standard to insure the instrument remains within calibration limits. The time between standardization shall not exceed 20 minutes. If the original calibration requirements are not met, all inspections performed since the last successful calibration shall be re-inspected. The following instrument settings are for a typical A-scan presentation ultrasonic instrument, and are meant as a guide; however sensitivity requirements shall be met. (a) Connect the probe to the cable and the cable to the instrument and turn the instrument on. (b) Set the instrument to the initial settings of Table 201.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Table 201 Initial Instrument Settings Description

Settings

Gain

40 dB (Decibels)

Range

1.25 inch

Delay

0 in/μs (inch/microseconds)

Velocity

0.1230 in/μs (inch/microseconds)

Pulse

50 ns (nanoseconds)

Damp

50Ω

Mode

Single

Gate 1

Positive

Gate Position

0.475 inch

Gate Width

0.304 inch

Gate Amplitude

30% FSH (Full Screen Height)

Display

Full Wave

Frequency

2 to 25 MHz BB (Broad Band)

Reject

OFF

(c) Apply couplant. Couple the transducer to the IIW block and adjust the gain to achieve an 80% FSH signal from the 0.060 inch diameter Sensitivity Hole, then add 6 dB. (d) Adjust the horizontal sweep to position the leading edge of the signal from the 0.060 inch diameter hole at 8.0 on the horizontal baseline. (e) Adjust the gate controls to position the gate at 30% FSH at 4.0 to 6.0 on the horizontal baseline. (3) Inspection Zone 1 NOTE: Inspection of Zone 1 is applicable to inspection as installed or uninstalled on the aircraft. (a) Apply couplant. Couple the transducer to the center of the cap with the sound beam directed at the edge of the actuator cap as shown in Figure 203.

Page 210 Nov 1/13

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

TRANSDUCER

UC32B 110087AA.AI

Figure 203 Transducer Scan (b) Scan 360° of the circumference of the part in the area of interest shown in Figures 202 and 203 for Zone 1. Index no greater than 0.10 inch for 100% coverage of the inspection zone. Use a typical “Z” scan path while angling the transducer to maximize any crack response received between 5.0 to 6.0 on the horizontal baseline. NOTE: A template (Ref. Figure 204) may be used to locate a reference line that will aid in the inspection.

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Page 211 Nov 1/13

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

FROM CENTER OF CAP OUTWARD 270 DEGREES

NON INSPECTION ZONE

A TEMPLATE WITH A 1.5” DIAMETER CIRCLE AS SHOWN BELOW MAY BE USED TO LOCATE A REFERENCE LINE ON THE PART SURFACE TO REPRESENT THE APPROXIMATE LOCATION ON THE SURFACE OF THE CAP IN WHICH THE SOUND BEAM’S EXIT POINT ON THE TRANSDUCER WOULD COINCIDE WITH THE DETECTION OF A CRACK AT THE INNER RADII.

0.60”

1.5” DIA.

UC32B 110088AA.AI

Figure 204 Scan Plan and Template (c) Scan outward and maximize the reflected signal from the machined contour of the internal surface of the cap, the signal will first appear at 7.5 and peak at approximately 6.5 on the horizontal baseline (Ref. Figure 205). Scan outward towards the outer radius of the cap.

TRANSDUCER

CRACKS HAVE BEEN FOUND IN THIS INTERNAL RADIUS.

UC32B 110173AA.AI

Figure 205 Outward Scan Page 212 Nov 1/13

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (d) As the transducer is scanned across the inspection zone toward the radius of the actuator end cap, the reflected signal from a crack on the first leg of the “V” path will first appear at 6.5 on the horizontal baseline and peak at 5.5 on the horizontal baseline. For detection of a crack at the inner radii, the sound beam’s exit point of the transducer will be approximately 0.60 inch to the edge of the cap. As the transducer is scanned toward the edge of the actuator cap other reflections may be seen further out in time, these reflections are non-relevant, the area of interest will be represented between 4.0 and 6.0 on the horizontal baseline. Reference Figure 206.

RELEVANT INDICATION GATE 30% FSH NON-RELEVANT INDICATIONS

1

2

3

4

5

6

7

8

9

10

UC32B 110090AA.AI

Figure 206 Relevant Indication NOTE: Some actuators may exhibit multiple non-relevant indications adjacent to the area of interest on the horizontal baseline. Such non-relevant indications may appear at 7 to 10 on the horizontal baseline (Ref. Figure 207). (e) Confirm instrument calibration at completion of inspection. If the original calibration requirements are not met, all inspections performed since the last successful calibration shall be re-inspected.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

GATE 30% FSH NON-RELEVANT INDICATIONS

1

2

3

4

5

6

7

8

9

10

UC32B 110091AA.AI

Figure 207 Non-relevant Indications (4) Indication Evaluation/Criteria Zone 1 (a) Any indication between 5.0 to 6.0 on the horizontal baseline shall be re-scanned to determine if the indication is false, caused by excessive couplant, hydraulic fluid within the actuator, part geometry, surface condition, or a defect. NOTE: Indications caused by droplets of hydraulic fluid on the inner surface will change their location and amplitude on the horizontal baseline when the actuator is manipulated or rotated. (b) Any repeatable sustainable crack response of 30% FSH or greater within the area of interest shall be rejected. (5) Inspection Zone 2 NOTE: Inspection of Zone 2 is applicable to inspection performed on actuators not installed on the aircraft. (a) Apply couplant. Couple the transducer to the part with the sound beam directed at the edge of the actuator cap as shown in Figure 208.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

TRANSDUCER

UC32B 110092AA.AI

Figure 208 Transducer Scan on Side of End Cap (b) Mark a reference line around the circumference 1.0 inch back from the end of the cap representing the inspection zone on the surface of the part. (c) Mark a reference line around the circumference 0.60 inch back from the end of the cap on the surface of the cap. This line represents the approximate location of the transducer sound beam‘s exit point in relation to crack detection at the inner radii. (d) Scan the outer circumference of the part in the area of interest shown in Figures 202 and 209. Index no greater than 0.10 inch for 100% coverage of the inspection zone. Use a typical “Z” scan path while angling the transducer to maximize any crack response received between 4.0 to 6.0 on the horizontal baseline. NOTE: Scanning must be contained within the inspection zone. Scanning outside of the inspection zone can lead to miss interpretation of indication as cracks due to normal ID geometry.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

INSPECTION ZONE 2 TRANSDUCER 1.0"

SCANNING DIRECTION SCAN PATH SOUND BEAM DIRECTED AT THE END OF THE CAP

UC32B 110093AA.AI

Figure 209 Scan Plan (e) As the transducer is scanned across the inspection zone toward the radius of the end cap of the actuator a reflected signal from the “O” ring groove will appear at 5.5 and peak at 5.0 on the horizontal baseline. As the transducer is moved further toward the end of the cap to inspect the area of interest the reflected signal from a crack will immediately follow the “O” ring signal appearing at 5.5 and peaking at 4.8 on the horizontal baseline. For detection of a crack at the inner radii, the sound beam’s exit point of the transducer will be approximately 0.60 inch to the edge of the cap. As the transducer is scanned further toward the edge of the actuator cap other reflections may be seen further out in time, these reflections are non-relevant, the area of interest will be represented between 3.0 and 6.0 on the horizontal baseline (Ref. Figure 210). NOTE: Indications caused by droplets of hydraulic fluid on the inner surface will change their location and amplitude on the horizontal baseline when the actuator is manipulated or rotated.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

RELEVANT INDICATION GATE 30% FSH NON-RELEVANT INDICATIONS

1

2

3

4

5

6

7

8

9

10

UC32B 110094AA.AI

Figure 210 Relevant Indication on Horizontal Baseline NOTE: Some actuators may exhibit multiple non-relevant indications adjacent to the area of interest on the horizontal baseline. Such non-relevant indications may appear at 7 to 10 on the horizontal baseline (Ref. Figure 211).

GATE 30% FSH THESE INDICATIONS MAY APPEAR IMMEDIATELY BEFORE A RELEVANT CRACK INDICATION WHEN SCANNING TOWARD THE END OF THE ACTUATOR CAP.

NON-RELEVANT INDICATIONS

1

2

3

4

5

6

7

8

9

10

GATE 30% FSH

NON-RELEVANT INDICATION

NON-RELEVANT INDICATIONS

1

2

3

4

5

6

7

8

9

10

UC32B 110095AA.AI

Figure 211 Non-relevant Indication

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (f) Confirm instrument calibration at completion of inspection. If the original calibration requirements are not met, all inspections performed since the last successful calibration shall be re-inspected. (6) Indication Evaluation/Criteria (a) Any indication between 4.0 to 5.0 on the horizontal baseline shall be re-scanned to determine if the indication is false, caused by excessive couplant, part geometry, surface condition, or a defect. NOTE: Indications caused by droplets of hydraulic fluid on the inner surface will change their location and amplitude on the horizontal baseline when the actuator is manipulative or rotated. (b) Any repeatable sustainable crack response of 30% FSH or greater within the area of interest shall be rejected. (7) Reporting (a) Mark cracks and record in aircraft log book. Report location and size to BC Service Engineering.

Page 218 Nov 1/13

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

LANDING GEAR NOSE LANDING GEAR DOOR MAINTENANCE PRACTICES

32-30-15 200200

1. PROCEDURES WARNING: When an airplane has experienced abnormal landing gear procedures of any type, as a safety precaution, place the airplane on jacks prior to performing any inspection or maintenance. Ensure that all three landing gears are down and locked prior to removing the airplane from jacks. When performing maintenance on a hydraulically operated landing gear system, be aware that any movement of a hydraulic actuator cylinder can displace hydraulic fluid and cause unanticipated movement of other actuator cylinders in the landing gear retraction system. Servicing of the landing gear hydraulic accumulator can also result in unanticipated movement of an actuator. Either action can result in an unsafe, unlocked landing gear system. Therefore, place the airplane on jacks prior to performing any inspection or maintenance. Cycle the landing gear and ensure that all three landing gears are down and locked prior to removing the airplane from jacks. Any time maintenance is to be performed on the landing gear system, place the airplane on jacks. When jacking the airplane in an unsheltered area where winds in excess of 35 kts will be encountered, never jack more than one gear at a time clear of the ground. Any time the landing gear is only partially retracted during maintenance, always cycle the gear with the power pack through at least one complete cycle before removing the airplane from the jacks. For safety reasons, pull the 2-ampere control circuit breaker on the pilot's inboard subpanel and place a note on the circuit breaker panel that LANDING GEAR MAINTENANCE IS IN PROCESS during maintenance on the landing gear. The landing gear control handle must never be moved from the down-and-locked position while the airplane is on the ground. It is recommended that the area be roped off during extension or retraction of the landing gear. CAUTION: Jacking of an airplane for the purpose of landing gear operation, inspection, servicing or maintenance, should be accomplished within an enclosed building or hangar. In the interest of safety, should it become necessary to jack the airplane in the open, wind velocity in any direction and terrain variations, must be compensated for prior to jacking the airplane. The landing gear must not be cycled with the power pack if low on fluid or if the landing gear system is not properly rigged. Use the emergency extension hand pump, TK229/939 hydraulic hand pump or TK229-1/939 air-driven hydraulic pump to extend and retract the landing gear for maintenance and rigging (Ref. Chapter 91-00-00).

A. Rigging The nose gear door is hinged at the front and connected to the nose gear brace with push-pull link assembly. As the nose gear is retracted, the door is pulled closed and as the nose gear is extended, the door is pushed open.

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Page 201 Nov 1/09

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Prior to rigging the nose landing gear door, the landing gear must be properly rigged (Ref. 32-30-13, NOSE LANDING GEAR RIGGING). While rigging the door, use the emergency extension hand pump to raise and lower the landing gear (Ref. 32-30-00, LANDING GEAR HAND PUMP CYCLING). (1) Perform the THREE-POINT JACKING (PREFERRED PROCEDURE) procedure (Ref. Chapter 07-10-00). All tires must be clear of the floor. Strut limiters will reduce strut extension (Ref. Chapter 91-00-00). (2) Perform APPLYING GROUND POWER procedure (Ref. Chapter 24-40-00). (3) Disconnect the two push-pull links from the nose landing gear brace. (4) With the landing gear fully retracted, manually close the door and check that the door is centered in the wheel well opening. If further adjustment is necessary, reposition the AN960-716 and/or AN960-716L washers. A minimum of one washer is required on each side of the hinges (Ref. Figure 201). (5) With the landing gear still retracted, manually close the door and check that the contour of the door matches that of the nose. Add or remove the laminated shims as necessary to correct any mismatch between the door and nose. (6) Extend the landing gear and connect the push-pull links to the nose landing gear brace. Install the washers in the same positions as when removed. Slowly retract the landing gear and check for a snug fit of the door when the landing gear is fully retracted. Adjust the length of the push-pull links at the rod ends to get the door to close snugly. (7) Lubricate the door hinges and the push-pull links with the proper lubricant (Ref. Chapter 12-20-00, NOSE LANDING GEAR LUBRICATION). CAUTION: To prevent serious damage to the pump, never operate the power pack when the engines are not operating without supplying 18 psi of regulated dry air to pressurize the power pack reservoir. When cycling the landing gear with the power pack, do not exceed three cycles in the first six minutes of operation with a two minute (approximately) cooling period between cycles. After the first six minute of operation, a five minute cooling interval between each cycle. (8) Cycle the landing gear several times with the power pack, checking each time that the door closes snugly when the landing gear is retracted. WARNING: Before removing the airplane from the jacks, make sure that the landing gear emergency extend hand pump handle is in the stowed position, the plunger on the service valve is pushed down with the hinged retainer in place, the landing gear control handle is in the DOWN position, the landing gear is down and locked and the accumulator is charged to 800 ± 50 psi. (9) Perform REMOVING GROUND POWER procedure (Ref. Chapter 24-40-00). (10) Perform LOWERING THE (Ref. Chapter 07-10-00).

Page 202 Nov 1/09

32-30-15

AIRPLANE

AFTER

THREE-POINT

JACKING

procedure

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 201 Nose Landing Gear Door Rigging

32-30-15

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

32-30-16 200200

LANDING GEAR HYDRAULIC LANDING GEAR SERVICE VALVE ASSEMBLY MAINTENANCE PRACTICES 1. PROCEDURES WARNING: When an airplane has experienced abnormal landing gear procedures of any type, as a safety precaution, place the airplane on jacks prior to performing any inspection or maintenance. Ensure that all three landing gears are down and locked prior to removing the airplane from jacks. When performing maintenance on a hydraulically operated landing gear system, be aware that any movement of a hydraulic actuator cylinder can displace hydraulic fluid and cause unanticipated movement of other actuator cylinders in the landing gear retraction system. Servicing of the landing gear hydraulic accumulator can also result in unanticipated movement of an actuator. Either action can result in an unsafe, unlocked landing gear system. Therefore, place the airplane on jacks prior to performing any inspection or maintenance. Cycle the landing gear and ensure that all three landing gears are down and locked prior to removing the airplane from jacks. Any time maintenance is to be performed on the landing gear system, place the airplane on jacks. When jacking the airplane in an unsheltered area where winds in excess of 35 kts will be encountered, never jack more than one gear at a time clear of the ground. Any time the landing gear is only partially retracted during maintenance, always cycle the gear with the power pack through at least one complete cycle before removing the airplane from the jacks. For safety reasons, pull the 2-ampere control circuit breaker on the pilot's inboard subpanel and place a note on the circuit breaker panel that LANDING GEAR MAINTENANCE IS IN PROCESS during maintenance on the landing gear. The landing gear control handle must never be moved from the down-and-locked position while the airplane is on the ground. It is recommended that the area be roped off during extension or retraction of the landing gear. CAUTION: Jacking of an airplane for the purpose of landing gear operation, inspection, servicing or maintenance, should be accomplished within an enclosed building or hangar. In the interest of safety, should it become necessary to jack the airplane in the open, wind velocity in any direction and terrain variations, must be compensated for prior to jacking the airplane. The landing gear must not be cycled with the power pack if low on fluid or if the landing gear system is not properly rigged. Use the emergency extension hand pump, TK229/939 hydraulic hand pump or TK229-1/939 air-driven hydraulic pump to extend and retract the landing gear for maintenance and rigging (Ref. Chapter 91-00-00).

32-30-16

Page 201 Nov 1/09

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

A. Removal (1) Perform the THREE-POINT JACKING (PREFERRED PROCEDURE) procedure (Ref. Chapter 07-10-00). All tires must be clear of the floor. Strut limiters will reduce strut extension (Ref. Chapter 91-00-00). (2) Remove the service valve access panel from the left center section of the wing (Ref. Figure 201). (3) Disconnect the service valve electrical connector (15) from the airplane wiring harness (Ref. Figure 202). (4) Remove the four tube assemblies (4, 8, 10, and 11) from the service valve (1) and cap the lines. (5) Cut the safety wire securing the two bolts (7) attaching the service valve assembly to the airplane structure. Remove the bolts (7), washers (6), nuts (12), the service valve (1), the plate assembly (5) and bracket assembly (14). (6) If the service valve is to be replaced, remove check valve (9), two unions (3) and packings (2).

B. Installation (1) If installing a new service valve, install the check valve (9) and two unions (3) with new packings (2). The check valve is to be installed with the arrow pointing away from the service valve (Ref. Figure 202). (2) Place the service valve assembly (1), plate assembly (5) and bracket assembly (14) on the bracket (13) and secure with two bolts (7), washers (6) and nuts (12). (3) Remove the caps from the tube assemblies (4, 8, 10, and 11) and connect the tube assemblies to the correct ports. (4) Perform ACCUMULATOR SERVICING procedure (Ref. 32-30-01) to ensure that the accumulator is charged to 800 ± 50 psi. (5) Perform the HYDRAULIC SYSTEM FILLING AND BLEEDING procedure to fill and bleed the landing gear hydraulic system (Ref. 32-30-00). (6) Connect the service valve electrical connector (15) to the airplane wiring harness. (7) Inspect all connectors for leaks. (8) If necessary, push down the service valve plunger (17), close the plunger retainer and secure with the turnlock (16). (9) Safety wire the two mounting bolts (7). CAUTION: To prevent serious damage to the pump, never operate the power pack when the engines are not operating without supplying 18 psi of regulated dry air to pressurize the power pack reservoir. When cycling the landing gear with the power pack, do not exceed three cycles in the first six minutes of operation with a two minute (approximately) cooling period between cycles. After the first six minute of operation, a five minute cooling interval between each cycle.

Page 202 Nov 1/09

32-30-16

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (10) Retract the landing gear using the power pack. (11) Extend the landing gear using the hand pump. (12) Install the service valve access panel on the left center section of the wing (Ref. Figure 201). (13) Cycle the landing gear and ensure that all three landing gears are down and locked prior to removing the airplane from jacks. WARNING: Before removing the airplane from the jacks, make sure that the landing gear emergency extend hand pump handle is in the stowed position, the plunger on the service valve is pushed down with the hinged retainer in place, the landing gear control handle is in the DOWN position, the landing gear is down and locked and the accumulator is charged to 800 ± 50 psi. (14) Perform REMOVING GROUND POWER procedure (Ref. Chapter 24-40-00). (15) Perform LOWERING THE (Ref. Chapter 07-10-00).

AIRPLANE

AFTER

THREE-POINT

JACKING

32-30-16

procedure

Page 203 Nov 1/09

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 201 Landing Gear Component Access

Page 204 Nov 1/09

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

9. CHECK VALVE 10. EMERGENCY EXTEND TUBE 11. RETRACT TUBE 12. NUT 13. BRACKET 14. SERVICE VALVE BRACKET ASSEMBLY 15. ELECTRICAL CONNECTOR 16. TURN LOCK 17. SERVICE VALVE PLUNGER

1. SERVICE VALVE 2. PACKING 3. UNION 4. HANDPUMP PRESSURE TUBE 5. PLATE ASSEMBLY 6. WASHER 7. BOLT 8. RETURN TUBE (PNEUMATIC) 12

6

A 16

17

15

1

4

14 2

3

13 5

12

6 6

7

2

11

9 8

2 3 10

6

DETAIL

A 7 UC32B 024040AB.AI

Figure 202 Hydraulic Landing Gear Service Valve Installation

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

LANDING GEAR EMERGENCY EXTENSION HAND PUMP ASSEMBLY MAINTENANCE PRACTICES

32-30-17 200200

1. PROCEDURE WARNING: When an airplane has experienced abnormal landing gear procedures of any type, as a safety precaution, place the airplane on jacks prior to performing any inspection or maintenance. Ensure that all three landing gears are down and locked prior to removing the airplane from jacks. When performing maintenance on a hydraulically operated landing gear system, be aware that any movement of a hydraulic actuator cylinder can displace hydraulic fluid and cause unanticipated movement of other actuator cylinders in the landing gear retraction system. Servicing of the landing gear hydraulic accumulator can also result in unanticipated movement of an actuator. Either action can result in an unsafe, unlocked landing gear system. Therefore, place the airplane on jacks prior to performing any inspection or maintenance. Cycle the landing gear and ensure that all three landing gears are down and locked prior to removing the airplane from jacks. Any time maintenance is to be performed on the landing gear system, place the airplane on jacks. When jacking the airplane in an unsheltered area where winds in excess of 35 kts will be encountered, never jack more than one gear at a time clear of the ground. Any time the landing gear is only partially retracted during maintenance, always cycle the gear with the power pack through at least one complete cycle before removing the airplane from the jacks. For safety reasons, pull the 2-ampere control circuit breaker on the pilot's inboard subpanel and place a note on the circuit breaker panel that LANDING GEAR MAINTENANCE IS IN PROCESS during maintenance on the landing gear. The landing gear control handle must never be moved from the down-and-locked position while the airplane is on the ground. It is recommended that the area be roped off during extension or retraction of the landing gear. CAUTION: Jacking of an airplane for the purpose of landing gear operation, inspection, servicing or maintenance, should be accomplished within an enclosed building or hangar. In the interest of safety, should it become necessary to jack the airplane in the open, wind velocity in any direction and terrain variations, must be compensated for prior to jacking the airplane. The landing gear must not be cycled with the power pack if low on fluid or if the landing gear system is not properly rigged. Use the emergency extension hand pump, TK229/939 hydraulic hand pump or TK229-1/939 air-driven hydraulic pump to extend and retract the landing gear for maintenance and rigging (Ref. Chapter 91-00-00).

32-30-17

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

A. Removal (1) Perform the THREE-POINT JACKING (PREFERRED PROCEDURE) procedure (Ref. Chapter 07-10-00). All tires must be clear of the floor. Strut limiters will reduce strut extension (Ref. Chapter 91-00-00). (2) Remove the center aisle carpet forward of the main spar cover. (3) Remove floorboard 4 adjacent to the emergency extension hand pump (Ref. Chapter 6-50-00). (4) Remove the three screws (12) from the upper hydraulic fluid drain cover assembly (15) and loosen the two screws on the side of the pump (Ref. Figure 201). Slide the cover up and out. (5) Remove the cotter pin (3), washer (4), and pin (11) connecting the pivot arm (1) to the pump clevis (8). (6) Disconnect and cap the hand pump pressure tube assembly (5) and the hand pump suction tube assembly (13) from the hand pump (10). (7) Remove the two nuts (14) from the hand pump (10) and remove the hand pump (10), two washers (4) and bolts (2). (8) Remove the reducers (6) and packings (7) from the hand pump (10).

B. Installation (1) Lubricate new packings with hydraulic fluid (39, Table 1, 91-00-00) and install the packings (7) on the reducers (6) (Ref. Figure 201). (2) Install the reducers (6) in the ports on the hand pump (10). (3) Place the hand pump (10) on the airplane structure and secure with two bolts (2), washers (4), and nuts (14). Ensure that the washers (4) are installed between the hand pump (10) and the structure. (4) Connect the hand pump clevis (8) to the pump pivot arm (1) with the clevis pin (11), washer (4), and cotter pin (3). (5) Connect the hand pump pressure tube assembly (5) and the hand pump suction tube assembly (13) to the hand pump reducers (6). (6) Operate the pump handle (16) to ensure there is no binding or misalignment in the connecting linkage. Inspect the pump (10) and its plumbing for fluid leaks. (7) Check the hand pump handle stow position by slowly lowering the handle (16) to the level of its stow clip. A pressure release in the pump should be felt just before the handle reaches the stow position. If pressure releases when the pump handle (16) is in any other position, adjust the linkage as follows: (a) Remove the cotter pin (3), washer (4), and clevis pin (11) connecting the clevis (8) to the handle pivot arm (1). (b) Loosen the jam nut (9) on the clevises.

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32-30-17

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL CAUTION: The length between the center lines of the attaching holes of each clevis must not exceed 1.90 inches after adjustment. (c) Lengthen or shorten the clevises as required by rotating clevis (8). (d) Ensure that the tang of the key washer is in the key slot and tighten the jam nut (9). (e) Connect the clevis (8) to the handle pivot arm (1) with the clevis pin (11), washer (4), and cotter pin (3). (8) Stow the handle under the clip. (9) Ensure that the accumulator is charged to 800 ± 50 psi with bottled dry nitrogen. Refer to LANDING GEAR ACCUMULATOR SERVICING procedure (Ref. 32-30-01). (10) Replace any hydraulic fluid lost and bleed the system. Refer to BLEEDING THE EMERGENCY EXTENSION HAND PUMP procedure below. (11) Install the cover (15) and secure with three screws (12) and tighten the two screws on the side of the pump. (12) Install floorboard 4 (Ref. Chapter 6-50-00) and the center aisle carpet. WARNING: Before removing the airplane from the jacks, make sure that the landing gear emergency extend hand pump handle is in the stowed position, the plunger on the service valve is pushed down with the hinged retainer in place, the landing gear control handle is in the DOWN position, the landing gear is down and locked and the accumulator is charged to 800 ± 50 psi. (13) Perform the LOWERING THE AIRPLANE AFTER THREE-POINT JACKING procedure (Ref. Chapter 07-10-00).

C. Bleeding WARNING: Any time maintenance is to be performed on the landing gear system, place the airplane on jacks. When jacking the airplane in an unsheltered area where winds in excess of 35 kts will be encountered, never jack more than one gear at a time clear of the ground. Any time the landing gear is only partially retracted during maintenance, always cycle the gear with the power pack through at least one complete cycle before removing the airplane from the jacks. For safety reasons, pull the 2-ampere control circuit breaker on the pilot's inboard subpanel and place a note on the circuit breaker panel that LANDING GEAR MAINTENANCE IS IN PROCESS during maintenance on the landing gear. The landing gear control handle must never be moved from the down-and-locked position while the airplane is on the ground. It is recommended that the area be roped off during extension or retraction of the landing gear.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL CAUTION: The landing gear must not be cycled with the power pack if low on fluid or if the landing gear system is not properly rigged. Use the emergency extension hand pump, TK229/939 hydraulic hand pump or TK229-1/939 air-driven hydraulic pump (Ref. Chapter 91-00-00) to extend and retract the landing gear for maintenance and rigging. Bleed trapped air from the emergency extension hand pump after the hand pump has been replaced or when any of the hydraulic system lines are opened at the hand pump. (1) Perform the THREE-POINT JACKING (PREFERRED PROCEDURE) procedure (Ref. Chapter 07-10-00). All tires must be clear of the floor. (2) Perform the APPLYING GROUND POWER procedure (Ref. Chapter 24-40-00). (3) Remove the upper and lower wing panel assembly as shown in Figure 201, 32-30-00. (4) Ensure the accumulator is charged to 800 ± 50 psi per LANDING GEAR HYDRAULIC ACCUMULATOR SERVICING, 32-30-01. NOTE: The regulated power supply should incorporate a shutoff valve to control the flow of pressurized air to the power pack. (5) Locate the cross-fitting forward of the power pack, remove the cap from the fitting, and connect a regulated supply of 18 to 20 psi dry air to the fitting. Shut off the air supply. (6) Pull the 2-ampere landing gear control circuit breaker. (7) Loosen the hand pump pressure line at the hand pump port of the service valve. (8) Unstow the handle of the emergency extension hand pump and pump the handle until fluid appears at the fitting and all indications of trapped air are eliminated. Tighten the pressure line fitting of the hand pump. (9) Reset the landing gear control circuit breaker and restore the air supply. CAUTION: To prevent serious damage to the pump, never operate the power pack without supplying 18 to 20 psi of regulated dry air to the power pack reservoir during ground operation of the power pack. When cycling the landing gear with the power pack, allow a one-minute cooling period between cycles and a five-minute cooling period every five cycles. (10) Extend and retract the landing gear with the power pack through ten complete cycles. After every third or fourth cycle, shut off the air supply and add fluid to the fill reservoir as required to replace any fluid lost from the displacement of trapped air. (11) Shut off the air supply. Disconnect the air supply from the cross fitting and replace the cap on the fitting. (12) Install the upper lower and lower wing panel assembly. (13) Perform the LOWERING THE AIRPLANE AFTER THREE-POINT JACKING procedure (Ref Chapter 07-10-00). (14) Perform the REMOVING GROUND POWER procedure (Ref. Chapter 24-40-00).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

1. PIVOT ARM 2. BOLT 3. COTTOR PIN 4. WASHER 5. HAND PUMP PRESSURE TUBE 6. REDUCER 7. PACKING 8. CLEVIS 9. JAM NUT 10. HAND PUMP 11. PIN 12. SCREW 13. HAND PUMP SUCTION TUBE 14. NUT 15. COVER 16. HANDLE

3

A

14 4

1

2 16 4

3

11

15 8 14

6

9

4

5

7 3 11

4

12

2

10

7

6 13 DETAIL

A

UC32B 024041AB.AI

Figure 201 Emergency Extension Hand Pump Installation

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

LANDING GEAR WHEELS AND BRAKES DESCRIPTION AND OPERATION

32-40-00 00

1. GENERAL WARNING: When an airplane has experienced abnormal landing gear procedures of any type, as a safety precaution, place the airplane on jacks prior to performing any inspection or maintenance. Ensure that all three landing gears are down and locked prior to removing the aircraft from jacks. CAUTION: Jacking of an airplane for the purpose of landing gear operation, inspection, servicing or maintenance, should be accomplished within an enclosed building or hangar. In the interest of safety, should it become necessary to jack the airplane in the open, wind velocity in any direction and terrain variations, must be compensated for prior to jacking the airplane.

A. Main Gear Wheel Assemblies Two 6.50 X 10 wheels are installed on each main landing gear (Ref. TIRES for size and type). Each wheel consists of an inner and outer wheel half held together with seven bolts and nuts with one washer under each bolt head and nut to prevent galling and stress concentration. A packing mounted in a groove on the inner wheel half seals the wheel against air leakage when the wheel halves are joined together. Seven keyway liners are riveted to each inboard wheel half and project into the disc drive slots of the flange to protect the slots from being battered by the brake rotating disc tangs. The wheels rotate on tapered roller bearings seated in bearing cups that are shrink-fitted into each wheel half. Grease seals are installed over each wheel bearing. The wheel assemblies are retained on each axle with two washers, a nut and a cotter pin.

B. Nose Gear Wheel Assembly One 6.50 X 8 wheel is installed on the nose landing gear (Ref. TIRES for size and type). The wheel consists of an inner and outer wheel half held together with eight bolts and nuts with one washer under each bolt head and nut to prevent galling and stress concentration. A packing mounted in a groove on the inner wheel half seals the wheel against air leakage when the wheel halves are joined together. The wheel rotates on tapered roller bearings seated in bearing cups that are shrink-fitted into each wheel half. Grease seals are installed over each wheel bearing. The wheel is retained on the axle with two washers, a nut and a cotter pin.

C. Tires The nose landing gear wheel is equipped with a 19.50 x 6.75-8, 10 ply rated, tube type or tubeless tire. Each main landing gear wheel is equipped with a 22 x 6.75-10, 8 ply rated, tubeless rim inflated tire.

D. Brake Assembly The airplane is equipped with four hydraulically operated brake assemblies (Ref. Figure 1). Each main landing gear incorporates two multi-disc, metallic lined brake assemblies bolted together, one on each side of the strut. The two are hydraulically interconnected by an inlet swivel fitting. Each brake assembly contains two inlet ports; one port is plugged and the other port accepts the inlet fitting. The brake assemblies are interchangeable by changing the locations of the inlet plug and fitting. A bleeder screw is located adjacent to each inlet port.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Each brake assembly contains two rotating discs, which are keyed to rotate with the wheel, a stationary disc and a pressure plate and back plate which are attached to the brake housing. Braking action occurs when hydraulic pressure is applied to the five small pistons in the brake housing, forcing the disc stack together and creating friction between the rotating discs and the stationary components of the brake.

E. Hydraulic Brake System The dual hydraulic brakes are operated by depressing the toe portion of either the pilot's or copilot's rudder pedals. The depression of either set of pedals compresses the piston rod in the master cylinder attached to each pedal. The hydraulic pressure resulting from the movement of the pistons in the master cylinders is transmitted through flexible hoses and fixed aluminum tubing to the disc brake assembly on each main landing gear wheel. Braking action occurs when hydraulic pressure at the brake pistons forces the disc stack of each brake assembly together creating friction between the rotating discs and the stationary components of the brake assembly. Hydraulic fluid is supplied to the master cylinders from a reservoir accessible through the nose avionics compartment door.

F. Parking Brake Valves Dual parking brake valves are plumbed in the lines between the master cylinders of the copilot's rudder pedals and the wheel brakes. The two lever type valves are located just aft of the flight compartment under the center aisle floorboard. A push-pull cable from the valve control levers runs to the pedestal, terminating with a knob. After the brake pedals have been depressed to build up pressure in the brake lines, both valves can be closed simultaneously by pulling out the parking brake knob on the pedestal. This closes the valves to retain the pressure that was previously pumped into the brake lines. The parking brake is released when the brake pedals are depressed briefly to equalize the pressure on both sides of the valves and the parking brake knob is pushed in.

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32-40-00

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 1 Brake System

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

LANDING GEAR WHEELS AND BRAKES TROUBLESHOOTING

100100

1. PROCEDURES WARNING: When an airplane has experienced abnormal landing gear procedures of any type, as a safety precaution, place the airplane on jacks prior to performing any inspection or maintenance. Ensure that all three landing gears are down and locked prior to removing the airplane from jacks. CAUTION: Jacking of an airplane for the purpose of landing gear operation, inspection, servicing or maintenance, should be accomplished within an enclosed building or hangar. In the interest of safety, should it become necessary to jack the airplane in the open, wind velocity in any direction and terrain variations, must be compensated for prior to jacking the airplane. Table 101 Troubleshooting Brake System PROBLEM

PROBABLE CAUSE

CORRECTIVE ACTION

1. Solid pedal and no brakes.

a. Brake lining worn beyond allowable limit.

a. Replace brake lining.

2. Spongy brake.

a. Air in system.

a. Bleed the brake system.

3. Unable to hold pressure.

a. Leak in brake system.

a. Visually check entire system for evidence of leaks. b. Check master cylinder seals, replace if scored.

4. Parking brake will not hold.

5. Brakes grab.

a. Air in system.

a. Bleed the brake system.

b. Parking brake valve fails to operate properly.

b. Replace parking brake valve.

c. Control cable out of adjustment.

c. Adjust cable.

a. Stones or foreign matter locking brake disc.

a. Remove foreign matter from brake discs.

b. Warped or bent disc.

b. Replace disc.

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LANDING GEAR WHEELS AND BRAKES MAINTENANCE PRACTICES

200200

1. PROCEDURES WARNING: When performing maintenance on a hydraulically operated landing gear system, be aware that any movement of a hydraulic actuator cylinder can displace hydraulic fluid and cause unanticipated movement of other actuator cylinders in the landing gear retraction system. Servicing of the landing gear hydraulic accumulator can also result in unanticipated movement of an actuator. Either action can result in an unsafe, unlocked landing gear system. Therefore, place the airplane on jacks prior to performing any inspection or maintenance. Cycle the landing gear and ensure that all three landing gears are down and locked prior to removing the airplane from jacks.

A. Tires Inflate the nose gear tire to 60 psi and the main gear tires to 95 psi loaded or 91 psi if the airplane is on jacks. CAUTION: Tires that have picked up a fuel or oil film should be washed down as soon as possible with a detergent solution to prevent contamination of the rubber. Maintaining the proper tire inflation will help to avoid damage from landing shock and contact with sharp stones and ruts, and will minimize tread wear. When inflating the tires, inspect them for cuts, cracks, breaks and tread wear. The pressure of a serviceable tire that is fully inflated should not drop more than 4 percent over a 24 hour period. NOTE: While Hawker Beechcraft Corporation cannot recommend the use of recapped tires, tires retreaded by an FAA approved repair station with a specialized service limited rating for TSO-C62c may be used.

B. Sealing Minor Leaks in Rim-Inflated Tubeless Tires Tire sealer (74, Table 1, Chapter 91-00-00), is recommended as an effective means of controlling the gradual loss of tire inflation pressure when the leakage rate does not exceed 5 percent over a 24 hour period. A M400 injector, for applying the sealant, is available from Hawker Beechcraft Corporation authorized outlets. Procedures for use of tire sealant are as follows: (1) Perform the THREE-POINT JACKING procedure (Ref. Chapter 07-10-00). Reduce the tire pressure to 35 psig or less. CAUTION: Keep out of line with the injector handle and remove the injector from the valve stem before releasing the handle. (2) Inject the tire sealant into the tire using one of the following methods: (a) If using the M400 injector and gallon container: 1 Screw the adapter (furnished with the injector) into the container. 2 Place the M400 ejector squarely on the adapter in a vertical position. 3 Prime the injector with several slow strokes of the handle.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL 4 Pull the handle full back to fill the injector with tire sealant. 5 Place the tip of the injector squarely on the valve stem and inject the tire sealant by pushing the handle completely in. The valve core does not need to be removed. One injection will be sufficient for each tire. (b) If using the pint container and the attached pump: 1 Install the pump on the pint container and tighten firmly. 2 Screw the adapter firmly onto the tire valve stem. 3 Operate the pump until two ounces, indicated on the container, has been pumped into the tire. 4 Remove the adapter from the valve stem. (3) Inflate the tire to the proper pressure, as specified under TIRES, to distribute the tire sealant within the tire. (4) Perform the LOWERING THE AIRPLANE AFTER THREE-POINT JACKING procedure (Ref. Chapter 07-10-00).

2. BRAKE SYSTEM BLEEDING Brake system bleeding will be required whenever the system is opened at any point between the master cylinder and the wheel brake assembly, whenever the brakes become spongy in service, or whenever the parking brakes will no longer hold. In the latter instance, the system should be further checked for leakage. Use only hydraulic fluid (39, Table 1, Chapter 91-00-00) in the brake system, and ensure that no dirt or foreign matter is allowed to get into the brake system. Dirt can get under seals and cause leaks or clog the compensating ports in the master cylinders and cause the brakes to lock. Hawker Beechcraft Corporation recommends the use of pressure pot brake bleeding. If the pressure pot bleeding method is not available, electric bleeding is recommended. Use the gravity method only if the other two methods are not available. If the gravity system is used, pressure bleed the brakes at the earliest possible time. Using any method, the parking brake lever and toe brake pedals must both be fully released to open the compensating ports in the brake master cylinders. If the brakes feel soft or spongy after the bleeding operation, air may be trapped in the cylinders. Remove the brake and lay it on its side. Add brake fluid as needed through the bleed port and tap the brake lightly with a rubber hammer to dislodge any air bubbles. When air bubbles no longer appear at the port, install the brake and repeat the bleeding procedure.

A. Pressure Bleeding Pressure bleeding is the most efficient method of bleeding the brake system. This procedure involves attaching a pressure pot to the brake assembly bleeder ports and back bleeding the system to the fluid reservoir. Procedures for utilizing the preferred pressure pot, the electric bleeder and the gravity bleed method are outlined below.

B. Pressure Pot Bleeding Method NOTE: The line hookup for pressure pot bleeding is the same as with the electric bleeder, except the electric bleeder is replaced with a pressure pot (Ref. Figure 201).

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32-40-00

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (1) Remove the sheet metal shield from around the brake fluid reservoir. (2) Disconnect the pressure equalization line from the reservoir and attach an extension line approximately three feet in length. (3) Place the end of the extension line in a clean receptacle to collect the brake fluid overflow. (4) Cut the safety wire and remove the screws from the bleeder ports of each brake assembly. Install a bleeder hose adapter into each brake bleeder port. Fabricate a bleeder hose assembly for each set of brakes; connect the bleeder hose assemblies between the bleeder hose adapters and the pressure lines of the pressure pot (Ref. Figure 201). (5) Apply a constant pressure of approximately 15 pounds to the pressure pot. Open the pressure pot control valve. (6) Bleed the system until the draining fluid is free of air bubbles. (7) Close the pressure pot valve. Remove the bleeder hose adapters and hose assemblies from each landing gear. Install the screws into the bleeder ports of each brake assembly and safety wire. (8) Remove the extension line from the pressure equalization port on the reservoir. (9) Connect the pressure equalization line to the reservoir. (10) Install the shield around the brake fluid reservoir. (11) Remove the cap from the hydraulic fluid reservoir and add hydraulic fluid (39, Table 1, Chapter 91-00-00) as required to obtain a full reading. (12) Check the operation of the brakes. There should be no soft or spongy feeling at the brake pedals and the pedal pressure should be equal on both brakes.

C. Electric Bleeder Method (1) Remove the sheet metal shield from around the brake fluid reservoir. (2) Disconnect the pressure equalization line from the reservoir and attach the electric bleeder fluid return line to the reservoir (Ref. Figure 201). (3) Cut the safety wire and remove the screws from the bleeder ports of each brake assembly. Install a bleeder hose adapter into each brake bleeder port. Fabricate a bleeder hose assembly for each set of brakes; connect the bleeder hose assemblies between the bleeder hose adapters and the pressure lines of the electric pressure bleeder. (4) Activate the bleeder and set the relief valve to approximately 15 pounds; this may be ascertained by observing the pressure gage prior to opening the electric bleeder control valve. (5) Open the electric bleeder control valve and observe the returning fluid through the inline sight glass. Pumping the pilot's and copilot's pedals during the bleeding process may help to dislodge any air bubbles trapped in the master cylinders. (6) When the returning fluid shows no further evidence of air bubbles, close the electric bleeder control valve.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (7) Disconnect the fluid infusion lines from the bleeder hose assemblies and remove the bleeder hose assemblies and adapters from the brake assemblies. Install the screws into the bleeder ports of each brake assembly and safety wire. (8) Disconnect the fluid return line from the brake fluid reservoir and reconnect the pressure equalization line. (9) Install the shield around the brake fluid reservoir. (10) Check the brake reservoir fluid level and add hydraulic fluid (39, Table 1, Chapter 91-00-00) as required to obtain a full reading. (11) Check the operation of the brakes. When the brake pedals are pressed there should be no spongy feeling and the pedal pressure should be equal on both brakes.

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32-40-00

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 201 Pressure Bleeding the Brake System

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

This Page Intentionally Left Blank

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

D. Gravity Bleeding Method This method of bleeding is done from the master cylinder down to the brake assembly. The brake fluid reservoir must be kept full during the bleeding operation. Since the pilot's and copilot's master cylinders are plumbed in series, the entire system may be bled by operating the pilot's brake pedals in the following manner: (1) Cut the safety wire and open the bleeder port screws of both brake assemblies on one landing gear. (2) Press the pilot's corresponding brake pedal slowly and smoothly to eliminate air trapped in the system. (3) Hold the brake pedal in the pressed position and close the bleeder port screws at the brake assemblies. (4) Release the brake pedal. (5) Repeat Steps (1), (2), (3) and (4) until no more air bubbles appear in the drained fluid. (6) Open the bleeder port screws of both brake assemblies on the other landing gear and repeat Steps (1), (2), (3) and (4), pressing the other brake pedal until no more air bubbles appear in the drained fluid. (7) Tighten the bleeder port screws at all four brake assemblies and safety wire. (8) Check the brake reservoir fluid level and add hydraulic fluid (39, Table 1, Chapter 91-00-00) as required to obtain a full reading. (9) Check the brakes for proper operation. When the brake pedals are pressed there should be no spongy feeling and the pedal pressure should be equal for both brakes.

3. BRAKE MASTER CYLINDER A. Removal (1) Bleed the hydraulic fluid from the brake system. (2) Remove the floorboards from around the master cylinder. (3) Disconnect and plug the brake lines at the cylinder. (4) Remove the master cylinder attaching clevis pins, washers, and cotter pins at the upper and lower end of the master cylinder. Remove the cylinder.

B. Installation (1) Place the brake master cylinder in position on the rudder pedal and install the attaching clevis pins, washers and cotter pins. (2) Connect the brake lines to the master cylinder. (3) Install the floorboards. (4) Perform the BRAKE SYSTEM BLEEDING procedure in this section.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

4. PARKING BRAKE VALVE A. Removal (1) Bleed the hydraulic fluid from the brake system. (2) Remove the first center aisle floorboard panel aft of the flight compartment. (3) Remove cover (1) from bowl assembly (2) (Ref. Figure 202). (4) Remove two screws (4) connecting control cable (5) to link (6). (5) Remove the bolts securing phenolic blocks (7) to bracket (8). (6) Disconnect the hydraulic lines from the valves. (7) Remove the bolts securing the valves to bracket (9). (8) Remove the valves and bracket (8) from the airplane. (9) Cut the safety wire and remove screws (10) attaching the valve control levers to link (6).

B. Installation (1) If elbows (11) or unions (12) were removed from the valves, install them using new rings (13) and packings (14) (Ref. Figures 202 and 203). (2) Connect valve control levers to link (6) with screws (10) and safety wire. (3) Position the valves and bracket (8) in the airplane and install the bolts securing them to bracket (9). (4) Attach the phenolic blocks (7) and control cable (5) to bracket (8). (5) Insert the end of the control cable into link (6) and secure with screws (4). (6) Connect hydraulic lines to the valves. (7) Perform the BRAKE SYSTEM BLEEDING procedure in this section. (8) Adjust the parking brake control cable. Refer to PARKING BRAKE CONTROL CABLE ADJUSTMENT. (9) Install cover (1) on bowl assembly (2). (10) Install center aisle floorboard.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 202 Parking Brake Valves

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

FITTING

COAT THE RING, PACKING, AND THE MALE THREADS OF THE FITTING WITH MIL-H-5606 HYDRAULIC FLUID AND ASSEMBLE AS SHOWN. WORK THE RING INTO THE COUNTERBORE OF THE NUT AND TURN THE NUT DOWN UNTIL THE PACKING IS PUSHED FIRMLY AGAINST THE LOWER THREADED PORTION OF THE FITTING.

NUT

RING PACKING STEP 1

INSTALL THE FITTING, TURNING THE NUT WITH THE FITTING, UNTIL THE PACKING CONTACTS THE SURFACE WITH THE FITTING IN THIS POSITION. PUT A WRENCH ON THE NUT TO PREVENT IT FROM TURNING, AND TURN THE FITTING IN 1 1/2 TURNS. POSITION THE FITTING BY TURNING IT IN NOT MORE THAN ONE ADDITIONAL TURN.

FITTING

NUT

RING PACKING

STEP 2

WHILE HOLDING THE FITTING TO PREVENT MOVEMENT, TURN THE NUT DOWN AGAINST THE SURFACE. SLIGHT EXTRUSION OF THE RING IS PERMISSIBLE.

FITTING

NUT RING PACKING

STEP 3

UC32B 062159AA.AI

Figure 203 Hydraulic Fitting Installation

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

5. NOSE WHEEL A. Removal (1) Perform the NOSE JACKING procedure (Ref. Chapter 07-10-00). Install strut limiters (Ref. SPECIAL TOOLS, Chapter 91-00-00) to reduce strut extension. (2) Remove cotter pin (1), axle nut (2), and washers (3) and (4) (Ref. Figure 204). CAUTION: Use care when removing all internal and external nose wheel components to avoid damaging the axle surface and/or nicking the axle threads. (3) Remove outer grease seal (5) and bearing cone (6) and remove the wheel and tire from axle. (4) Remove the inner bearing cone (7) and grease seal (8) from the axle.

B. Installation WARNING: The following procedures for wheel and bearing installation must be followed exactly. Failure to do so could result in the loss of wheels and/or property damage. CAUTION: Do not mix greases of different types or manufacturers. If the grease is changed, make certain that all the affected components are thoroughly cleaned before relubrication. NOTE: Wheel bearing lubricants authorized for the Model 1900/1900C Airliner are listed in Table 1 under “Wheel Bearing Grease” (87, Table 1, Chapter 91-00-00). (1) Pack grease (87, Table 1, Chapter 91-00-00) into bearing cones (6 and 7) and apply grease on ends of the rollers. Do not over lubricate. Spread a thin coat of bearing grease on surface of the bearing cups (9) (Ref. Figure 204). (2) Check the axle and nut for burrs or rough threads. (3) Carefully place inner grease seal (8) on the axle against axle shoulder with the lips of the seal facing the bearing. Place inner bearing cone (7) on axle against the seal, place the wheel and tire on the axle, and seat the bearing cone in the cup. Install outer bearing cone (6) on the axle and seat the cone in the outer bearing cup. (4) Install the outer grease seal (5) on the axle with the lips of seal facing the bearing. NOTE: To ensure cotter key (1) is held securely (not loosely) a maximum of two washers (4) may be used. (5) Apply a light coat of grease to threads of the axle and the nut and install the axle nut (2) and washers (3) and (4) on the axle. Use the same grease as used on wheel bearings. (6) Tighten axle nut 250 to 300 inch-pounds torque while rotating wheel to ensure proper seating of the bearings. (7) Back off axle nut to zero torque, then torque nut 125 to 145 inch-pounds while rotating wheel. Check to see that there is no side motion of the wheel. (8) Install cotter pin (1). If holes do not align, tighten nut to the next available keying position.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (9) Perform the LOWERING (Ref. Chapter 07-10-00).

THE

AIRPLANE

AFTER

NOSE

JACKING

procedure

Figure 204 Nose Wheel

6. MAIN WHEEL AND BRAKE A. Removal (1) Perform the THREE-POINT JACKING (PREFERRED PROCEDURE) procedure or SINGLE-POINT JACKING (FOR WHEEL, TIRE AND BRAKE MAINTENANCE ONLY) procedures (Ref. Chapter 07-10-00). Install strut limiters reduce strut extension (Ref. SPECIAL TOOLS, Chapter 91-00-00). (2) Remove screws securing hubcap (1) or (17) to the wheel and remove hubcap (Ref. Figure 205). (3) Remove the wheel retaining nut (3 or 21), washer (5) and key washer (4 or 22) as follows (Ref. Figure 205): (a) On airplanes with antiskid brakes installed, cut the safety wire and loosen the lock screws (18) and remove the lock ring (20) from the axle. Remove the wheel retaining nut (21) and key washer (22). (b) On airplanes without antiskid brakes installed, remove the cotter pin (2) securing the wheel retaining nut (3). Remove the wheel retaining nut (3) and key washer (4).

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32-40-00

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL CAUTION: Use care when removing all internal and external wheel components to avoid damaging the axle surface and/or nicking the axle threads. (4) Disconnect brake hydraulic line (11) from inlet swivel fitting (12) and cap line and fitting. NOTE: On airplanes with brake deice installed, remove the brake assemblies and deice manifold (Ref. 32-42-00). (5) Remove bolts (16), washers, and nuts securing the two brake assemblies to one another. (6) Remove both brake assemblies from the strut, being careful not to drop or damage inlet swivel fitting (12).

B. Installation WARNING: The following wheel and wheel bearing installation procedures must be followed exactly. Failure to do so could result in wheel loss and/or property damage. CAUTION: DO NOT MIX greases of different types or manufacturers. If the grease is changed, make certain that all the affected components are thoroughly cleaned before lubrication. NOTE: On airplanes with brake deice installed, install the brake assemblies and deice manifold then proceed with Step (7) (Ref. 32-42-00). Wheel bearing lubricants authorized for the Model 1900/1900C Airliner are listed in Table 1 under “Wheel Bearing Grease” (87, Table 1, Chapter 91-00-00). (1) Position the inboard brake assembly on the shock strut. Apply a thin coat of grease to the inside diameter of the brake housing and on the axle at assembly. (2) Lubricate packings (13) with hydraulic fluid (39, Table 1, Chapter 91-00-00), and install the packings in the grooves in the inlet swivel fitting (12). Lightly lubricate the surfaces next to the grooves (Ref. Figure 205). (3) Install the inlet swivel fitting on inlet adapter (14) of the inboard brake assembly, being careful not to displace or damage packing (13). (4) Position the outboard brake assembly on the shock strut. Apply a thin coat of grease (same as that used on the wheel bearings) to inside diameter of the brake housing and on the axle at assembly. Carefully mate inlet adapter (14) on the outboard brake with inlet swivel fitting (12), being careful not to displace or damage packing (13). (5) Ascertain that both brake assemblies are correctly seated on the strut and axle and that the inlet swivel fitting (12) swivels freely. (6) Secure the brake assemblies together with attaching bolts (16), washers, and nuts; install one washer under each bolt head and one washer under each nut. At the time of installation lubricate the bolt and nut threads and the bearing surfaces of the bolts, washers, and nuts with antiseize thread compound (76, Table 1, Chapter 91-00-00). Torque the self-locking nuts to 120 inch-pounds. (7) Connect the hydraulic brake fluid line (11) to the inlet swivel fitting (12). (8) Pack grease into bearing cones (7) and (9) and smear grease on the ends of the rollers. Do not over lubricate. Spread a thin coat of bearing grease on the surface of the bearing cups (8) and the lips of the grease seals (6) and (10).

32-40-00

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (9) Check the axle and nut for burrs or rough threads. (10) Apply grease to the axle threads and all bearing surfaces of the washers (4) and (5) or (22) and nut (3) or (21). (11) Install inner grease seal (10) and bearing cone (9) on the axle. The lips of the grease seal must be toward the wheel. (12) Place wheel and tire on the axle and install outer bearing cone (7) and grease seal (6). CAUTION: Make sure that the grease seals are properly seated in the wheel against the bearing cups and that the seals are not damaged. Visually inspect that the main landing gear brake rotor tangs for both disks are positioned in the main wheel assembly slots. (13) Install washers (4) and (5), or washer (22), on the axle. Place the axle nut (3) or (21) on the axle. While normally rotating the wheel, torque the axle nut 225 inch-pounds. (14) Back off the axle nut to 0 to 25 inch-pounds. While normally rotating the wheel retorque to 110 inch-pounds. (15) Secure the wheel retaining nut as follows (Ref. Figure 205): (a) On airplanes without antiskid brakes installed, install a new cotter pin (2) through the axle and nut. If necessary to install the cotter pin, tighten the nut to the next castellation, but do not exceed 160 inch-pounds of torque or 30° of arc. Additional 101-810207-5 washers may be used as required to allow alignment of the nut slot and the cotter pin hole. After the cotter pin is installed, bend the cotter pin along the curve of the axle to prevent contact between the cotter pin and the hubcap. (b) On airplanes with antiskid brakes installed, place the lock ring (20) against the nut (21). Align the holes in the lock ring (20) and axle. If the holes do not align, tighten the nut to allow installation of the lock ring. Use thin washers as required so as not to exceed 160 inch-pounds of torque or 30° of arc on the nut. Additional 101-810207-5 washers may be used as required to allow alignment of the lock ring screws and the lock ring screw holes. Tighten the screws (18) in the lock ring (20) and safety wire. (16) Secure hubcap (1) or (17) to the wheel with the attaching screws. (17) Repeat Steps (8) thru (16) to install the opposite wheel and tire. (18) Perform the BRAKE SYSTEM BLEEDING in this section. Bleed antiskid brakes (Ref. 32-41-00). WARNING: Before removing the airplane from the jacks, make sure that the landing gear emergency extend hand pump handle is in the stowed position, the plunger on the service valve is pushed down with the hinged retainer in place, the landing gear control handle is in the DOWN position, the landing gear is down and locked and the accumulator is charged to 800 ± 50 psi. (19) Perform LOWERING THE AIRPLANE AFTER THREE-POINT JACKING or LOWERING THE AIRPLANE AFTER SINGLE-POINT JACKING procedures (Ref. Chapter 07-10-00).

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7. MAIN WHEEL A. Removal (Using a Ramp) WARNING: This procedure is an alternate method for removing the main landing gear wheel. Instead of placing the airplane on jacks, a ramp may be used. It may only be used for removal and installation of a main landing gear wheel, and no other maintenance may be performed on the airplane while the airplane is on the ramp. (1) Install nose and main landing gear downlocks (Ref. SPECIAL TOOLS, Chapter 91-00-00). (2) Place airplane on the ramp as follows (Ref. SPECIAL TOOLS, Chapter 91-00-00): (a) Position ramp directly in front of wheel adjacent to the wheel being removed. WARNING: Due to the weight of the airplane being supported on one wheel, the supporting wheel must be centered on the ramp to avoid any unanticipated movements of the airplane that may cause damage to the airplane and/or injury to personnel. (b) Carefully tow airplane forward and allow desired wheel to roll onto the ramp. Refer to Chapter 09-10-00 for towing instructions. NOTE: The ramp may move while attempting to roll airplane onto the ramp. A sandbag may be used to chock ramp or attach a rubber mat to bottom of the ramp using best shop practice. (3) Set parking brake (Ref. Chapter 10-10-00). (4) Chock the opposite side main landing gear wheels and nose landing gear wheel. NOTE: In ice or snow conditions, ice-grip wheel chocks are preferred. Sandbags may be used if ice-grip chocks are not available, or if the airplane is parked on a steel mat. (5) Remove the screws securing hubcap (1) or (17) to the wheel and remove hubcap (Ref. Figure 205). (6) Remove cotter pin (2); on airplanes with antiskid brakes, cut the safety wire and back off screws (18) far enough to allow lock ring (20) to be removed from nut (21). Remove nut (3) or (21), and washers (4) and (5) or (22) securing each wheel and tire to the axle. When removing the wheels and tires from the axle, be careful not to drop or damage grease seals and bearing cones.

B. Installation (Using a Ramp) CAUTION: This procedure can only be used when the MAIN WHEEL REMOVAL (USING A RAMP) procedure has been performed. (1) Pack grease into bearing cones (7) and (9) and smear grease on ends of the rollers. Do not over lubricate. Spread a thin coat of bearing grease on the surface of the bearing cups (8) and lips of grease seals (6) and (10) (Ref. Figure 205). CAUTION: DO NOT MIX greases of different types or manufacturers. If grease is changed, make certain that all affected components are thoroughly cleaned before lubrication.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (2) Check the axle and nut for burrs or rough threads. (3) Apply grease to the axle threads and all bearing surfaces of washers (4) and (5) or (22) and nut (3) or (21). (4) Install inner grease seal (10) and bearing cone (9) on the axle. The lips of the grease seal must be toward the wheel. (5) Place wheel and tire on the axle and install outer bearing cone (7) and grease seal (6). CAUTION: Make sure that the grease seals are properly seated in the wheel against bearing cups and that the seals are not damaged. Visually inspect that the main landing gear brake rotor tangs for both disks are positioned in the main wheel assembly slots. (6) Install washers (4) and (5), or washer (22), on the axle. Place axle nut (3) or (21) on the axle. While normally rotating the wheel, torque the axle nut to 225 inch-pounds. (7) Back off axle nut 0 to 25 inch-pounds and retorque to 110 inch-pounds while manually rotating wheel. (8) Secure the wheel retaining nut as follows (Ref. Figure 205): (a) On airplanes without antiskid brakes installed, install a new cotter pin (2) through the axle and nut. If necessary to install the cotter pin, tighten the nut to the next castellation, but do not exceed 160 inch-pounds of torque or 30° of arc. Additional 101-810207-5 washers may be used as required to allow alignment of the nut slot and the cotter pin hole. After the cotter pin is installed, bend the cotter pin along the curve of the axle to prevent contact between the cotter pin and the hubcap. (b) On airplanes with antiskid brakes installed, place the lock ring (20) against the nut (21). Align the holes in the lock ring (20) and axle. If the holes do not align, tighten the nut to allow installation of the lock ring. Use thin washers as required so as not to exceed 160 inch-pounds of torque or 30° of arc on the nut. Additional 101-810207-5 washers may be used as required to allow alignment of the lock ring screws and the lock ring screw holes. Tighten the screws (18) in the lock ring (20) and safety wire. (9) Secure hubcap (1) or (17) to the wheel with the attaching screws. (10) Remove all chocks from the wheels. (11) Release the parking brake (Ref. Chapter 10-10-00). (12) Carefully tow the airplane off the ramp. Refer to Chapter 09-10-00 for towing instructions. (13) Remove the nose and main landing gear downlocks.

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A 22

11

18 19 20

15 12 13 14 15

17

21 19 18

B

DETAIL AIRPLANES WITH ANTISKID BRAKES

1. HUBCAP 2. COTTER PIN 3. NUT 4. KEY WASHER 5. WASHER 6. OUTER GREASE SEAL 7. OUTER BEARING CONE 8. BEARING CUP 9. INNER BEARING CONE 10. INNER GREASE SEAL 11. BRAKE HYDRAULIC LINE 12. INLET SWIVEL FITTING 13. PACKING 14. INLET ADAPTER 15. PACKING 16. BOLT 17. HUBCAP 18. SCREW 19. WASHER 20. LOCK RING 21. NUT 22. KEY WASHER

10 9

16 8 7 DETAIL

A

2 3

5

1

4

6

B UC32B 062147AA.AI

Figure 205 Main Wheels and Brakes

32-40-00

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8. BRAKE MASTER CYLINDER LINKAGE A. Adjustment (1) Loosen the locknut on the piston shaft. (2) Remove clevis pin, washer, and cotter pin attaching the master cylinder clevis to the rudder pedal. Slide clevis free of the lug on the rudder pedal. (3) The length of the cylinder piston can be adjusted by turning clevis on or off the piston shaft. Adjust clevis so that the distance between the centerline of clevis hole and the centerline of master cylinder bottom attach hole is 9 inches. The length of the shaft should be adjusted so that the piston does not bottom when the brake is applied. (4) Move master cylinder clevis into position to mate with the lug on the rudder pedal and install clevis pin, washer, and cotter pin. (5) Tighten clevis locknut. (6) Check upper and lower attachment points of the master cylinder for possible interference with attachment lugs on the rudder pedal.

9. PARKING BRAKE CONTROL CABLE A. Adjustment NOTE: Overhaul of the parking brake valves is not recommended. If leakage is encountered, replace the valve. To ensure proper operation of the parking brake valves, it is important that full travel of the valves be maintained. Adjust the control cable as follows: (1) Remove first center aisle floorboard panel aft of the flight compartment. (2) Remove cover from the bowl assembly (Ref. Figure 206). (3) Loosen bolts securing the phenolic blocks to the bracket and position outer housing of the cable in the blocks until pushing in on the control knob fully opens the valves. Tighten bolts. (4) Pull control knob out and check that the valves close fully. (5) If necessary, loosen bolts and reposition the outer housing of the cable in the blocks until the valves open and close fully. (6) Install the cover on the bowl assembly. (7) Install the center aisle floorboard.

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Figure 206 Parking Brake and Control Cable Adjustment

10. BRAKE A. Wear Limits To check the total disc wear, measure the distance between the back of the pressure plate and the brake housing with normal brake pressure applied from the master cylinder (Ref. Figure 207). Replace the brake assembly when it has reached the maximum wear limit. To determine the remaining flight cycle for a brake assembly (Ref. Table 201). Table 201 Brake Wear Measurement vs. Remaining Flight Cycles P/N 5006749-4 WEAR MEASUREMENT

P/N 5006749-5 WEAR MEASUREMENT

REMAINING FLIGHT CYCLES

0.340

0.400

0

0.339

0.399

4

0.338

0.398

8

0.337

0.397

12

0.336

0.396

16

0.335

0.395

20

0.334

0.394

24

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Table 201 Brake Wear Measurement vs. Remaining Flight Cycles (Continued) P/N 5006749-4 WEAR MEASUREMENT

P/N 5006749-5 WEAR MEASUREMENT

REMAINING FLIGHT CYCLES

0.333

0.393

28

0.332

0.392

32

0.331

0.391

36

0.330

0.390

40

0.329

0.389

44

0.328

0.388

48

0.327

0.387

52

0.326

0.386

56

0.325

0.385

60

0.324

0.384

64

0.323

0.383

68

0.322

0.382

72

0.321

0.381

76

0.320

0.380

80

NOTE: Table 201 is based on average wear rates. Wear rates may vary between operators and brakes depending on operating conditions. This Table should only be used as a guide. Brakes must be removed from service at full wear. A flight cycle is defined as: Engine start-up and increase to full or partial power (as required during a normal flight), one landing gear retraction and extension and a complete shutdown.

B. Fluid Reservoir Pressure Equalization Filter Cleaning The brake fluid reservoir is plumbed so that cabin air is directed into the top of the reservoir through the pressure equalization line. A filter is located in the pressure equalization line just forward of the forward pressure bulkhead. This filter is to be removed and cleaned as follows: (1) Disconnect the line to the top of the reservoir from the filter, and remove the filter from the tee at the forward pressure bulkhead. (2) Clean the filter with cleaning solvent (2, Table 1, Chapter 91-00-00), and blow dry with compressed air. (3) Install the filter on the tee, and connect the line from the reservoir to the filter.

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C. Fluid Reservoir Pressure Equalization Orifice Cleaning Cabin air contaminants (fuzz, lint, nicotine, tar, etc.) can find their way into the orifice installed in the pressure equalization line and eventually reduce the orifice so that excessive air is diverted into the brake fluid reservoir, causing excessive pressure to be applied to the master cylinders. Clean the orifice as follows: (1) Disconnect the line from the brake fluid reservoir to the pressure equalization line filter and remove the filter. (2) Where the filter was removed, inject a small amount of cleaning solvent (2, Table 1, Chapter 91-00-00) into that line with low air pressure. Check the solvent flowing from the bottom drain for cleanliness. Repeat as necessary until clean solvent flows from the bottom drain. (3) After clean solvent flows from the bottom drain, blow out the line thoroughly with low air pressure. (4) Install the filter. Connect the line from the brake fluid reservoir to the filter.

Figure 207 Brake Wear Limits

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LANDING GEAR ANTISKID BRAKES DESCRIPTION AND OPERATION

32-41-00 00

1. GENERAL WARNING: When an airplane has experienced abnormal landing gear procedures of any type, as a safety precaution, place the airplane on jacks prior to performing any inspection or maintenance. Ensure that all three landing gears are down and locked prior to removing the aircraft from jacks. CAUTION: Jacking of an airplane for the purpose of landing gear operation, inspection, servicing or maintenance, should be accomplished within an enclosed building or hangar. In the interest of safety, should it become necessary to jack the airplane in the open, wind velocity in any direction and terrain variations, must be compensated for prior to jacking the airplane. In addition to the components of the standard brake system, the antiskid system incorporates a number of electrical and hydraulic components: an antiskid box assembly mounted in the nose gear wheel well contains a pump and motor assembly, filter, check valve, power brake valve, and high and low pressure switches (Ref. Figure 1). Additional equipment includes a control switch mounted on the pedestal, a pump motor relay and two-position three-way solenoid valve mounted above the antiskid box assembly, an accumulator mounted in the nose wheel well, a skid control unit mounted under the copilot's floorboard and wheel speed transducers mounted in each main gear axle. The antiskid control circuit is protected by a 5-ampere circuit breaker located on the RH circuit breaker panel. The control switch, placarded ANTISKID-OFF, energizes the motor relay and the two-position solenoid valve when it is placed in the ANTISKID position. Power for the pump motor is supplied through the motor relay and the 20-ampere limiter located in the LH nacelle electrical distribution panel. The motor relay is grounded through the high pressure switch and the nose gear downlock switch. The actuation points of the high pressure switch are set to maintain the system pressure between 950 to 1375 psi. Fluid under pressure from the pump is directed through the solenoid valve to the pressure inlet port of the power brake valve. An accumulator (precharged to 600 ±100 psi) plumbed in the line from the pump to the power brake valve aids in maintaining the system pressure. Any time the system pressure drops below 615 to 685 psi, the low pressure switch will ground a relay located under the second floorboard panel forward of the main spar, energizing the time delay relay under the first floorboard panel aft of the main spar, which will illuminate a yellow ANTISKID FAIL annunciator in the CAUTION/ADVISORY panel. A maintenance switch mounted on the brake fluid reservoir provides power to open the two-position solenoid valve when the reservoir cap is removed. The dual hydraulic brakes are operated by depressing the toe portion of either the pilot's or copilot's rudder pedals. The depression of either set of pedals compresses the piston rod in the master cylinder attached to each pedal. The hydraulic pressure resulting from the movement of the pistons in the master cylinders is transmitted through flexible hoses and fixed aluminum tubing to the power brake valve. Braking action occurs when hydraulic pressure (generated by the master cylinders or metered pressure from the pump) at the brake pistons forces the disc stack of each brake assembly together creating friction between the rotating discs and the stationary components of the brake assembly.

A. Manual Mode When the antiskid system is in the manual mode (control switch on pedestal in the OFF position), there will be no fluid pressure from the pump at the pressure inlet port of the power brake valve (Ref. Figure 2). This allows the pressure generated at the master cylinders to force the shuttle valves in the power brake valve to open.

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B. Antiskid Mode With the brake system in the antiskid mode, the pump pressure present at the inlet port of the power brake valve will position the internal shuttle valves to isolate the master cylinder pressure from the brake assemblies and to open a passage for metered pump pressure to the brake assemblies (Ref. Figure 3). Master Cylinder pressure in the power brake valve acts on the control valve piston to close the return passage and to open the passage for pump pressure into the control valve. The resulting metered pressure is applied to the control valve spool to balance the pressure delivered to the brake assemblies with the master cylinder pressure. As the master cylinder pressure is increased or decreased, the control valve spool moves to allow fluid to escape through the return passage decreasing the metered pressure or to close the return port allowing the metered pressure to increase. The wheel speed transducers monitor the wheel RPM and relay the information to the skid control unit to be processed. The skid control unit sends electrical signals to the servo valve mounted on the power brake valve based on the information provided by the wheel transducers. The servo valve responds to the input from the skid control unit by varying its control pressure. The control pressure is created in the cavity of the servo valve between the system pressure nozzle and the system return nozzle. Signals from the skid control unit control the opening and closing of the two nozzles. When no electrical signal from the skid control unit is present, the pressure nozzle is open and the return nozzle is closed; therefore, the control pressure of the servo valve is the same as the pump pressure and has no effect on the operation of the control valve. When a skid condition is imminent, the appropriate signal is sent to the servo valve and the servo valve control pressure is reduced as the pressure nozzle is closed. When the control pressure is reduced, pump pressure at the opposite end of the control valve spools moves the spools to open the return passages allowing fluid in the control valves to bleed off, resulting in a reduced pressure to all four brake assemblies. As the wheel RPM begins to increase, the servo valve control pressure increases, resulting in an increased brake pressure. If the wheel transducers indicate that the wheels are still in a skid condition, the brake pressure will be decreased again. This cycle will continue until the tendency for the wheels to skid has ended. Differential braking between the LH and RH master cylinders will be maintained while the antiskid function is operating.

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Figure 1 (Sheet 1 of 2) Antiskid Brake System

32-41-00

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Figure 1 (Sheet 2 of 2) Antiskid Brake System

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UE32B 990886AA.AI

Figure 2 (Sheet 1 of 2) Antiskid Brake System Schematic (Manual Mode)

32-41-00

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UE32B 990885AA.AI

Figure 2 (Sheet 2 of 2) Antiskid Brake System Schematic (Manual Mode)

32-41-00

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UE32B 990884AA.AI

Figure 3 (Sheet 1 of 2) Antiskid Brake System Schematic (Antiskid Mode)

32-41-00

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

MAINTENANCE SWITCH

CABIN PRESSURE FILTER OVERBOARD LINE

RESERVOIR

ORIFICE PUMP MOTOR RELIEF VALVE

PUMP

20A

28VDC

LIMITER TB1

FILTER CHECK VALVE

CAUTION/ADVISORY ANNUNCIATOR ANTISKID FAIL

ON

MOTOR RELAY

5A 28VDC ANTISKID CIRCUIT BREAKER

OFF CONTROL SWITCH

ACCUMULATOR PILOT'S MASTER CYLINDERS

COPILOT'S MASTER CYLINDERS

SOLENOID VALVE TB2

CONTROL VALVES

RELAY PANEL ASSY NO. 3 A120

TIME DELAY RELAY

SHUTTLE VALVE

LH MAIN GEAR

SHUTTLE VALVE

ANTISKID BOX ASSEMBLY

PARKING BRAKE VALVE

PARKING BRAKE VALVE

CODE

INBOARD TRANSDUCER

SUPPLY FLUID

SERVO VALVE

MASTER CYLINDER PRESSURE PUMP PRESSURE OUTBOARD TRANSDUCER

INBOARD TRANSDUCER

POWER BRAKE VALVE

METERED PRESSURE RETURN FLUID SERVO VALVE CONTROL PRESSURE

RH MAIN GEAR

LOW PRESSURE SWITCH

HIGH PRESSURE SWITCH NOSE GEAR DOWN-POSITION SWITCH

A B N M

R OUTBD TRANSDUCER R OUTBD TRANS RETURN SYSTEM GROUND CHASSIS GROUND

E F K L G H C D J C

R INBD TRANSDUCER R INBD TRANS RETURN VALVE VALVE RETURN L INBD TRANSDUCER L INBD TRANS RETURN L INBD TRANSDUCER L INBD TRANS RETURN FAULT IND DRIVER +28 VOLTS DC

OUTBOARD TRANSDUCER

UE32B 990887AA.AI

SKID CONTROL UNIT

Figure 3 (Sheet 2 of 2) Antiskid Brake System Schematic (Antiskid Mode)

32-41-00

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LANDING GEAR ANTISKID BRAKES MAINTENANCE PRACTICES

32-41-00 200200

1. PROCEDURES WARNING: When performing maintenance on a hydraulically operated landing gear system, be aware that any movement of a hydraulic actuator cylinder can displace hydraulic fluid and cause unanticipated movement of other actuator cylinders in the landing gear retraction system. Servicing of the landing gear hydraulic accumulator can also result in unanticipated movement of an actuator. Either action can result in an unsafe, unlocked landing gear system. Therefore, place the airplane on jacks prior to performing any inspection or maintenance. Cycle the landing gear and ensure that all three landing gears are down and locked prior to removing the airplane from jacks. NOTE: Refer to 32-40-00 for the following maintenance practices: removal and installation of the master cylinders, parking brake valves, nose wheel, and main wheel and brakes; adjustment of the parking brake cable and the master cylinder linkage; brake wear limits; and cleaning of the pressure equalization filter and orifice.

A. Servicing Brake system servicing is limited primarily to maintaining the hydraulic fluid level in the reservoir mounted in the upper RH corner of the aft bulkhead of the nose avionics compartment and to maintaining an accumulator precharge of 600 ± 100 psi. To check the fluid level, turn the master switch ON, turn the antiskid switch OFF, ensure that the parking brake is off, remove the reservoir cap and depress the brake pedals 15 to 20 times to deplete the accumulator. If no fluid is visible in the reservoir sight glass, add a sufficient quantity of hydraulic fluid (39, Table 1, 91-00-00) to raise the fluid level to the lower edge of the filler neck.

B. Bleeding Brake system bleeding will be required whenever the system is opened at any point between the master cylinders and the wheel brake assemblies, whenever the brakes become spongy in service, or whenever the parking brakes will no longer hold. In the latter instance, the system should be further checked for leakage. Use only hydraulic fluid (39, Table 1, 91-00-00) in the brake system and ensure that no dirt or foreign matter is allowed to get into the brake system. Dirt can get under seals and cause leaks or clog the compensating ports in the master cylinders and cause the brakes to lock. Hawker Beechcraft Corporation recommends the use of pressure pot bleeding. If the pressure pot bleeding method is not available, electric bleeding is recommended. Use the gravity method only if the other two methods are not available. If the gravity method is used, pressure bleed the brakes at the earliest possible time. Using any method, the parking brake control and toe brake pedals must both be fully released to open the compensating ports in the brake master cylinders. If the brakes feel soft or spongy after the bleeding operation, air may be trapped in the cylinders. Remove the brake and lay it on its side. Add brake fluid as needed through the bleed port and tap the brake lightly with a rubber hammer to dislodge any air bubbles. When air bubbles no longer appear at the port, install the brake and repeat the bleeding procedure.

32-41-00

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C. Gravity Bleeding Method This method of bleeding is done from the master cylinder down to the brake assembly. The brake fluid reservoir must be kept full during the bleeding operation. Since the pilot's and copilot's master cylinders are plumbed in series, the manual portion of the system may be bled by operating the pilot's brake pedals only. Bleed the brakes in the following manner: (1) Cut the safety wire and open the bleeder port screws of both brake assemblies on one landing gear. (2) Press the pilot's corresponding brake pedal slowly and smoothly to eliminate air trapped in the system. (3) Hold the brake pedal in the pressed position and close the bleeder port screws at the brake assemblies. (4) Release the brake pedal. (5) Repeat Steps (1), (2), (3), and (4) until no more air bubbles appear in the drained fluid. (6) Open the bleeder port screws of the other brake assembly and repeat Steps (1), (2), (3), and (4) while pressing the other brake pedal until no more air bubbles appear in the drained fluid. (7) Check the brake reservoir fluid level and add hydraulic fluid (39, Table 1, 91-00-00) as required to obtain a full reading. (8) Remove the screws from the brake bleeder ports of each brake assembly. Install a bleeder hose adapter into each brake bleeder port. Fabricate a bleeder hose assembly for each set of brakes; connect the bleeder hose assemblies to the bleeder hose adapters (Ref. 32-40-00). (9) Connect 25 feet of tygon tubing with an in-line filter to the bleeder hose assembly on the LH landing gear; remove the brake fluid reservoir cap and place the other end of the tygon tubing in the reservoir. Cap the bleeder hose assembly on the RH landing gear. (10) Connect an external power supply to the airplane and adjust the voltage to 28 ± 0.25 volts (Ref. Chapter 24-40-00). (11) Place the antiskid control switch in the ANTISKID position and allow the system to normalize. (12) Apply light pressure to the pilot's left brake pedal and hold. Allow the antiskid system pump to slowly purge all the remaining air from the left side of the brake system. When the brake fluid returning to the reservoir is free of air, place the antiskid control switch in the OFF position. (13) Disconnect the tygon tubing from the LH brake assemblies and connect it to the bleeder hose assembly on the RH brake assemblies; cap the bleeder hose assembly on the LH landing gear. (14) Turn the antiskid control switch to the ANTISKID position and allow the system to normalize. (15) Apply light pressure to the pilot's right brake pedal and hold. Allow the antiskid system pump to slowly purge all the remaining air from the right side of the brake system. When the brake fluid returning to the reservoir is free of air, place the antiskid control switch in the OFF position. (16) Disconnect the tygon tubing from the RH brake assemblies.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (17) Remove the bleeder hose assemblies and adapters from the brake assemblies. Install the screws into the bleeder ports of each brake assembly and safety wire. (18) Check the brake reservoir fluid level and add hydraulic fluid (39, Table 1, 91-00-00) as required to obtain a full reading. (19) With the antiskid control switch in the OFF position check the brakes for proper operation. When the brake pedals are pressed, there should be no spongy feeling and the pedal pressure should be equal for both brakes.

D. Pressure Bleeding Pressure bleeding is the most efficient method of bleeding the brake system. This procedure involves attaching a pressure pot or an electric bleeder to the brake fluid reservoir and bleeding the system to the brake assemblies. Procedures for utilizing the pressure pot and the electric bleeder are outlined below.

E. Pressure Pot Bleeding Method (1) Remove the sheet metal shield from around the brake fluid reservoir. (2) Disconnect the pressure equalization line from the reservoir and attach the pressure line of the pressure pot to the reservoir. (3) Cut the safety wire and remove the screws from the bleeder ports of each brake assembly. Install a bleeder hose adapter into each brake bleeder port. Fabricate a bleeder hose assembly for each set of brakes; connect the bleeder hose assemblies to the bleeder hose adapters (Ref. 32-40-00). (4) Connect an extension line to each bleeder hose assembly and place the ends of the extension lines in a clean receptacle to collect the brake fluid overflow. (5) Apply a constant pressure of approximately 15 pounds to the pressure pot. Open the pressure pot control valve. (6) Bleed the system until the draining fluid is free of air bubbles. (7) Close the pressure pot valve. (8) Remove the pressure pot pressure line from the reservoir. Connect the pressure equalization line to the reservoir, and install the shield around the reservoir. (9) Remove the extension lines from the bleeder hose assemblies. (10) Remove the cap from the hydraulic fluid reservoir and add hydraulic fluid (39, Table 1, 91-00-00) as required to obtain a full reading. (11) Connect 25 feet of tygon tubing with an in-line filter to the bleeder hose assembly of the LH landing gear; remove the brake fluid reservoir cap and place the other end of the tygon tubing in the reservoir. Cap the bleeder hose assembly on the RH landing gear. (12) Connect an external power supply to the airplane and adjust the voltage to 28 ± 0.25 volts (Ref. Chapter 24-40-00). (13) Place the antiskid control switch in the ANTISKID position and allow the system to normalize.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (14) Apply light pressure to the pilot's left brake pedal and hold. Allow the antiskid system pump to slowly purge all the remaining air from the left side of the brake system. When the brake fluid returning to the reservoir is free of air, place the antiskid control switch in the OFF position. (15) Disconnect the tygon tubing from the LH brake assemblies and connect it to the bleeder hose assembly on the RH brake assemblies; cap the bleeder hose assembly on the LH landing gear. (16) Turn the antiskid control switch to the ANTISKID position and allow the system to normalize. (17) Apply light pressure to the pilot's right brake pedal and hold. Allow the antiskid system pump to slowly purge all the remaining air from the right side of the brake system. When the brake fluid returning to the reservoir is free of air, place the antiskid control switch in the OFF position. (18) Disconnect the tygon tubing from the RH brake assemblies. (19) Remove the bleeder hose assemblies and adapters from the brake assemblies. Install the screws into the bleeder ports of each brake assembly and safety wire. (20) Check the brake reservoir fluid level and add hydraulic fluid (39, Table 1, 91-00-00) as required to obtain a full reading. (21) With the antiskid control switch in the OFF position, check the brakes for proper operation. When the brake pedals are pressed, there should be no spongy feeling and the pedal pressure should be equal for both brakes.

F. Electric Bleeder Method (1) Remove the sheet metal shield from around the brake fluid reservoir. (2) Disconnect the pressure equalization line from the reservoir and attach the electric bleeder fluid infusion line. (3) Cut the safety wire and remove the screws from the bleeder ports of each brake assembly. Install a bleeder hose adapter into each brake bleeder port. Fabricate a bleeder hose assembly for each set of brakes; connect the bleeder hose assemblies to the bleeder hose adapters (Ref. 32-40-00). (4) Connect the electric bleeder fluid return line to the bleeder hose assemblies on each brake assembly. (5) Activate the bleeder and set the relief valve to approximately 15 pounds; this may be ascertained by observing the pressure gage prior to opening the brake bleeder valves. (6) Open the electric bleeder control valve and observe the returning fluid through the in-line sight glass. Pumping the pilot's and copilot's pedals during the bleeding process may help to dislodge any air bubbles trapped in the master cylinders. (7) When the returning fluid shows no further evidence of air bubbles, close the electric bleeder control valve. (8) Disconnect the fluid infusion line and the fluid return line from the airplane. (9) Connect 25 feet of tygon tubing with an in-line filter to the bleeder hose assembly of the LH landing gear; remove the brake fluid reservoir cap and place the other end of the tygon tubing in the reservoir. Cap the bleeder hose assembly on the RH landing gear.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (10) Connect an external power supply to the airplane and adjust the voltage to 28 ± 0.25 volts (Ref. Chapter 24-40-00). (11) Place the antiskid control switch in the ANTISKID position and allow the system to normalize. (12) Apply light pressure to the pilot's left brake pedal and hold. Allow the antiskid system pump to slowly purge all the remaining air from the left side of the brake system. When the brake fluid returning to the reservoir is free of air, place the antiskid control switch in the OFF position. (13) Disconnect the tygon tubing from the LH brake assemblies and connect it to the bleeder hose assembly on the RH brake assemblies; cap the bleeder hose assembly on the LH landing gear. (14) Turn the antiskid control switch to the ANTISKID position and allow the system to normalize. (15) Apply light pressure to the pilot's right brake pedal and hold. Allow the antiskid system pump to slowly purge all the remaining air from the right side of the brake system. When the brake fluid returning to the reservoir is free of air, place the antiskid control switch in the OFF position. (16) Disconnect the tygon tubing from the RH brake assemblies. (17) Remove the bleeder hose assemblies and adapters from the brake assemblies. Install the screws into the bleeder ports of each brake assembly and safety wire. (18) Check the brake reservoir fluid level and add hydraulic fluid (39, Table 1, 91-00-00) as required to obtain a full reading. (19) With the antiskid control switch in the OFF position, check the brakes for proper operation. When the brake pedals are pressed, there should be no spongy feeling and the pedal pressure should be equal for both brakes.

2. ANTISKID SYSTEM A. Functional Check This check is to be used whenever major components of the antiskid system have been replaced, when faulty system operation is suspected, or at intervals specified in Chapter 05-10-00. The procedure will verify proper operation of the complete system including components, plumbing, control and associated wiring. The mechanical coupling of the hubcap and drive clip to the wheel speed transducer is not covered in this check. This must be visually checked during installation of the hubcaps. NOTE: The brake hydraulic system must be thoroughly bled before performing this procedure. The following equipment is required to perform the functional test: •

A hand held drill with a range of 1000 to 2000 rpm.

•

A transducer drive fixture fabricated locally (Ref. Figure 201 or 202).

•

Pressure gages with a range of 0 to 1000 psi.

Two technicians are required to perform the test. (1) Release the parking brake and chock both the left and right wheels. (2) Remove all the main gear hubcaps (Ref. 32-40-00).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (3) Install pressure gages in the brake bleed port. (4) Connect the airplane to an external 28 vdc power source (Ref. Chapter 24-40-00). (5) Close the antiskid circuit breaker and place the antiskid control switch in the ON position. (6) Apply full brake pressure and hold through the next five Steps. All brake pressures should be approximately 600 psi. (7) Momentarily place the antiskid switch to the OFF position. The brake pressure should not change more than approximately 50 psi. Return the antiskid switch to the ON position. (8) Spin the left outboard wheel speed transducer at a constant 1000 to 2000 rpm using the hand held drill and transducer drive fixture. Abruptly stop rotation of the transducer and note an immediate drop in pressure at all brakes followed by gradual increase of pressure (smooth return to original pressure) within 10 to 20 seconds (time is not critical). (9) Spin the left inboard wheel speed transducer at a constant 1000 to 2000 rpm with the hand held drill. Abruptly stop rotation of the transducer and note an immediate drop in pressure at all brakes followed by gradual increase of pressure (smooth return to original pressure) within 10 to 20 seconds (time is not critical). (10) Spin the right outboard wheel speed transducer at a constant 1000 to 2000 rpm with the hand held drill. Abruptly stop rotation of the transducer and note an immediate drop in pressure at all brakes followed by gradual increase of pressure (smooth return to original pressure) within about 10 to 20 seconds (time is not critical). (11) Spin the right inboard wheel speed transducer at a constant 1000 to 2000 rpm with the hand held drill. Abruptly stop rotation of the transducer and note an immediate drop in pressure at all brakes followed by a gradual increase of pressure (smooth return to original pressure) within about 10 to 20 seconds (time is not critical). (12) The functional check is now complete. Remove the pressure gages and install all hubcaps so the transducer drive clips properly engage the wheel speed transducer couplings (Ref. 32-40-00).

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0.50 0.25

0.64

0.06

0.32

0.32

0.50

DRILL 0.189 HOLE FOR 10-32 SCREW 0.26

GRIND TO CONICAL POINT AS SHOWN

135°

1.40

1.5 X 10-32 SCREW 0.50

REMOVE HEAD

1. 2. 3. 4.

ALL DIMENTIONS ARE IN INCHES. TWO STANDARD NUTS ARE USED TO RETAIN SCREW TO CHANNEL. BEND ENDS OF CHANNEL UNIT A SPACE OF 0.26 INCH IS OBTAINED. CONICAL TIP ENGAGES RETAINING SCREW IN TRANSDUCER DRIVE COUPLING MAINTAINING CONCENTRICITY DURING ROTATION. 5. FABRICATE FROM 0.05 ALUMINUM CHANNEL OR 0.060 SHEET ALUMINUM. UC32B 091926AA.AI

Figure 201 Transducer Drive Fixture

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NOTES 1. MAKE TOOL BY CUTTING AND FILING ON DOTTED LINES PER BEST SHOP PRACTICE. 2. MAINTAIN DIMENSIONS SHOWN. 3. ROUND ALL SHARP EDGE. 0.25

0.60 0.56

MAKE FROM MS21252-ERS CLEVIS

1.125

0.25

UC32B 091927AA.AI

Figure 202 Transducer Drive Fixture

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3. ANTISKID SYSTEM ACCUMULATOR A. Removal (1) Turn the master switch ON, turn the antiskid switch OFF, ensure that the parking brake is off, remove the reservoir cap and depress the brake pedals 15 to 20 times to deplete the accumulator. (2) Turn the master switch to OFF and disconnect the battery. (3) Remove 12 screws and two bolts from the antiskid accumulator cover and remove the cover assembly. (4) Slowly open the schrader valve to deplete the air side of the accumulator. (5) Disconnect the lines to the accumulator and cap all hydraulic lines to prevent contamination. (6) Remove the two screws from the adel clamps holding the accumulator and remove the accumulator.

B. Installation (1) Position accumulator into the two adel mount clamps and secure with two screws. (2) Remove the protective caps from the hydraulic lines and install the lines to the accumulator. (3) Service the accumulator with nitrogen to 600 ± 100 psi. (4) Perform the antiskid Functional Check (Ref. Pressure Check). (5) After completion of test, ensure that there are no leaks at the accumulator. (6) Install the antiskid accumulator cover with 12 screws and two bolts.

C. Pressure Check NOTE: Do not check accumulator pressure immediately after operating the antiskid system. Pressure buildup in the system may inhibit accurate accumulator pre-charge readings. (1) Turn the master switch ON. (2) Turn the antiskid switch OFF. (3) Ensure that the parking brake is off. (4) Remove the reservoir cap. (5) Press the brake pedal 15 to 20 times to deplete the accumulator charge. (6) Visually check the accumulator pressure indicator gage. (7) Install the reservoir cap. NOTE: Accumulator precharge pressure should be 600 ± 100 psi.

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4. ANTISKID BRAKE SYSTEM FILTER A. Removal WARNING: Whenever the brakes are to be released, make sure the airplane is on level ground and the wheels are chocked. (1) Remove the cover from the antiskid brake box assembly located at top of the nose gear wheel well (Ref. Figure 203). (2) Clean the outside of the filter housing with solvent (2, Table 1, 91-00-00). (3) Unscrew the filter housing.

B. Installation (1) Clean the inside of the filter housing with solvent (2, Table 1, 91-00-00) (Ref. Figure 203). (2) Lube the packing of the new filter element with hydraulic fluid (39, Table 1, 91-00-00). (3) Install the new filter element in the filter housing. (4) Lube the packing of the filter housing with hydraulic fluid (39, Table 1, 91-00-00). (5) Install a new packing on the filter housing. (6) Install the filter housing. (7) Turn aircraft master switch and antiskid switch to on. (8) Pump brake pedals 15 to 20 times. (9) Turn aircraft master switch and antiskid switch to off. (10) Verify there are no leaks at the filter housing. (11) Install the cover on the antiskid brake box assembly.

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TWO-POSITION SOLENOID VALVE HIGH PRESSURE SWITCH FILTER LOW PRESSURE SWITCH CHECK VALVE

POWER BRAKE VALVE

PUMP AND MOTOR ASSEMBLY

C9100657.AI

Figure 203 Antiskid Brake System Filter

5. HYDRAULIC FITTINGS A. Installation To reduce the possibility of leaks in the hydraulic system, care should be taken when installing the hydraulic fittings. Anytime a fitting is loosened or removed, discard the packing and install the fitting with a new packing (Ref. Figure 204). To prevent damage to the packing, coat the packing and the threads of the fitting with hydraulic fluid (39, Table 1, 91-00-00) before installing the packing on the fitting. Some fittings in the brake system employ both an packing and a ring. When this type of fitting is loosened or removed, both the packing and the ring must be replaced when the fitting is installed.

6. SOLENOID VALVE A. Removal The solenoid valve is located on top of the antiskid box assembly, which must be removed to gain access. (1) Turn the master switch ON, turn the antiskid switch OFF, ensure that the parking brake is off, remove the reservoir cap and depress the brake pedals 15 to 20 times to deplete the accumulator. (2) Remove electrical power from the airplane (Ref. Chapter 24-40-00). (3) Working in the nose gear wheel well, remove the screws securing the cover to the antiskid box (Ref. Figure 1, Description and Operation section). (4) Remove the hydraulic line/accumulator cover. (5) Slowly open the schrader valve on the accumulator to deplete the air.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (6) Drain the hydraulic fluid from the system by loosening the tee caps from the two lower tubes under the pump motor. (7) Disconnect the tubes below the antiskid box as necessary to remove the box. Cap the ends of the open lines as they are disconnected. (8) Disconnect the electrical connector to the box. (9) Remove the bolts securing the box assembly to the airplane and remove the box assembly. (10) Disconnect the electrical connector from the solenoid valve. (11) Disconnect the hydraulic lines from the solenoid valve. CAUTION: As the hydraulic lines are disconnected from the valve, plug or cap the openings to prevent entry of foreign material into the lines or valve. (12) Remove the two bolts securing the solenoid valve to the antiskid box assembly.

B. Installation The solenoid valve is located on top of the antiskid box assembly, which must be removed to gain access. (1) Install the two bolts securing the solenoid valve to the antiskid box assembly (Ref. Figure 1, Description and Operation section). (2) Connect the hydraulic lines to the solenoid valve. (3) Connect the electrical connector to the solenoid valve. (4) Secure the box assembly to the airplane with the attaching bolts. (5) Remove the caps and connect the tubes previously removed below the box assembly. (6) Connect the electrical connector to the box. (7) Tighten any loose tubes as required. (8) Refill the brake fluid reservoir and bleed the system as instructed in this section. (9) Service the accumulator (Ref. PRESSURE CHECK). (10) Connect the battery and restore electrical power to the airplane. (11) Perform the Antiskid Functional Check as instructed in the section. (12) After completion of test, ensure that there is no leaks at all hydraulic lines that were disturbed. (13) Install the hydraulic line/accumulator cover. (14) Install the cover on the antiskid box assembly and secure with the attaching screws.

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7. POWER BRAKE VALVE A. Removal The solenoid valve is located on top of the antiskid box assembly, which must be removed to gain access. (1) Turn the master switch ON, turn the antiskid switch OFF, ensure that the parking brake is off, remove the reservoir cap and depress the brake pedals 15 to 20 times to deplete the accumulator. (2) Remove the hydraulic line/accumulator access cover. (3) Slowly open the schrader valve on the accumulator to deplete the air. (4) Working in the nose gear wheel well, remove the lower cover from the antiskid box assembly (Ref. Figure 1, Description and Operation section). (5) Disconnect the electrical wiring from the valve from TB2. (6) Disconnect the hydraulic lines from the valve. CAUTION: As the hydraulic lines are disconnected from the valve, plug or cap the openings to prevent entry of foreign material into the lines or valve. (7) Remove the three bolts securing the valve to the antiskid box assembly.

B. Installation (1) Position the valve in the antiskid box assembly and secure the valve to the box with the three screws (Ref. Figure 1, Description and Operation section). (2) Connect the hydraulic lines to the valve. (3) Connect the electrical wiring to the valve to TB2 (Ref. WDM 32-42-02). (4) Service the accumulator (Ref. PRESSURE CHECK). (5) Install the hydraulic line/accumulator access cover. (6) Connect the battery and restore electrical power to the airplane. (7) Refill the brake fluid reservoir and bleed the system as instructed in this section. NOTE: If power steering is installed, bleed the power steering system (Ref. 32-52-00). (8) Perform Antiskid Functional Check (Ref. 32-41-00). (9) Install the lower cover removed from the antiskid box assembly.

8. PUMP AND MOTOR A. Removal WARNING: Whenever the brakes are to be released, make sure the airplane is on level ground with the wheels chocked.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (1) Turn the master switch ON, turn the antiskid switch OFF, ensure that the parking brake is off, remove the reservoir cap and depress the brake pedals 15 to 20 times to deplete the accumulator. (2) Remove the hydraulic line/accumulator access cover. (3) Slowly open the schrader valve on the accumulator to deplete the air. (4) Remove electrical power from the airplane (Ref. Chapter 24-40-00). (5) Working in the nose gear wheel well, remove the lower cover from the antiskid box assembly (Ref. Figure 1, Description and Operation section). NOTE: Airplanes prior to UC-48 Without Kit No. 114-8017 Installed, it is necessary to remove the entire antiskid box assembly in order to detach the pump motor power wire from the relay located on top of the box. On airplanes UA-1, UB-1 and UC-1 thru UC-47, Kit No. 114-8017 must be installed on the airplane the first time the pump and motor are replaced. (6) Drain the hydraulic fluid from the system by loosening the tee caps from the two lower tubes under the pump motor. (7) Disconnect the electrical wiring from the motor. (8) Disconnect the hydraulic lines from the pump. CAUTION: As the hydraulic lines are disconnected from the pump, plug or cap the openings to prevent entry of foreign material into the lines or valve. (9) Remove the bolts securing the pump and motor to the antiskid box assembly.

B. Installation (1) Position the pump and motor in the antiskid box assembly and secure the pump and motor to the box with the bolts (Ref. Figure 1, Description and Operation section). (2) Remove caps and connect the hydraulic lines to the pump. (3) Connect the electrical wiring to the motor. (4) Service the accumulator (Ref. PRESSURE CHECK). (5) Install the hydraulic line/accumulator access cover. (6) Connect the battery and restore electrical power to the airplane. (7) Refill the brake fluid reservoir and bleed the system as instructed in this section. NOTE: If power steering is installed, bleed the power steering system (Ref. 32-52-00). (8) Select the master switch to ON, select the antiskid switch to ON. Pump the brake pedals for 15 to 20 times. Select the antiskid and master switches to OFF. (9) Verify that there are no leaks. (10) Install the lower cover removed from the antiskid box assembly.

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Figure 204 Hydraulic Fitting Installation

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9. ANTISKID BRAKE SYSTEM WHEEL-SPEED TRANSDUCER A. Removal WARNING: Whenever the brakes are to be released, make sure the airplane is on level ground with the wheels chocked. NOTE: This procedure is typical for the right and left components. (1) Remove electrical power to the airplane and disconnect the battery. (2) Remove three screws (8) securing the hubcap (7) to the wheel and remove the hubcap (7) (Ref. Figure 205). CAUTION: Use care when removing all internal and external wheel components to avoid damaging the axle surface and/or nicking the axle threads. (3) Remove two lock screws (1) and two washers (2) from the landing gear axle. (4) Slide the wheel-speed transducer (5) out far enough from axle to access mating electrical connector (3). (5) Disconnect electrical connector (3) from wheel-speed transducer (5) and remove wheel-speed transducer.

B. Installation (1) Connect electrical connector (3) to wheel-speed transducer (5) (Ref. Figure 205). (2) Carefully pull wiring between the two wheels out of the gear socket while positioning wheel speed transducer (5) into position in the axle. CAUTION: Do not allow the axle nut and lock screws (1) to contact the wheel-speed transducer tang (6) and inhibit transducer rotation. (3) Align the holes in the rim of the wheel-speed transducer (5) with mounting holes in axle. (4) Apply coat of Thread lock (161, Table 1, Chapter 91-00-00) onto threads of lock screws (1) and install two lock screws (1) and two washers (2), into axle and wheel-speed transducer (5). (5) Align hubcap spring drive clip (9) with wheel-speed transducer tang (6) and install hubcap (7). (6) Install three screws (8) through hubcap (7) and secure to wheel, safety wire screws.

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1. LOCK SCREW 2. WASHER 3. ELECTRICAL CONNECTOR 4. LOCK RING 5. WHEEL-SPEED TRANSDUCER 6. WHEEL-SPEED TRANSDUCER TANG 7. HUB CAP 8. SCREW 9. HUBCAP SPRING DRIVE CLIP

1 3

2

4 5 6

2

7

1 8 7

8 5

9

6

A

8

3

1 2

AXLE (REF) DETAIL

A

UE32B 022293AA.AI

Figure 205 Antiskid Brake System Wheel-Speed Transducer

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LANDING GEAR BRAKE DEICE SYSTEM DESCRIPTION AND OPERATION

32-42-00 00

1. GENERAL WARNING: When an airplane has experienced abnormal landing gear procedures of any type, as a safety precaution, place the airplane on jacks prior to performing any inspection or maintenance. Ensure that all three landing gears are down and locked prior to removing the aircraft from jacks. CAUTION: Jacking of an airplane for the purpose of landing gear operation, inspection, servicing or maintenance, should be accomplished within an enclosed building or hangar. In the interest of safety, should it become necessary to jack the airplane in the open, wind velocity in any direction and terrain variations, must be compensated for prior to jacking the airplane. The heated air for brake deicing is supplied by bleeding air from the compressor of each engine. The brake deicer is plumbed into the bleed air system that provides air for surface deice and instrument vacuum operation. The engine bleed air line is routed near the center of each nacelle to a solenoid-operated firewall shutoff valve in each main gear wheel well. From the shutoff valve, bleed air is routed through a hose secured to the aft side of the landing gear strut and down to a distributor manifold attached to each brake assembly. The bleed air is directed to the brake for each wheel through orifices around the circumference of each ring of the distributor manifold (Ref Figures 1 and 2). The brake deice system is controlled by a toggle switch (placarded BRAKE DEICE/OFF) mounted on the pilot's inboard subpanel. When this switch is in the BRAKE DEICE position, power from the airplane electrical system is supplied through a 5-ampere circuit breaker in the RH circuit breaker panel to the brake deice control module assembly, located under the center aisle floorboard nine feet forward of the main spar. Current is then supplied from this point to open the solenoid shutoff valves on the aft side of the firewall in the main gear wheel wells, allowing the hot bleed air to enter the distributor manifold for diffusion through the orifices to deice the brakes. A switch, which is part of the solenoid shutoff valve, provides a signal to light the green L BK DEICE ON and R BK DEICE ON annunciators in the CAUTION/ADVISORY panel when the shutoff valve is open. If the airplane is flown without the brake deice switch having been turned to OFF, a circuit is completed through the brake deice control module relay to a timing circuit in the brake deice control module. The timing circuit will close the shutoff valves after 10 minutes of operation, shutting off the flow of bleed air to the brakes so that adjacent components will incur no damage through overheating. If it is desired to turn the brake deice system back on, the landing gear must be lowered and the switch cycled OFF, then ON, before the system will operate. Additional protection against damage from overheating of the brake deice system is provided by the brake deice overtemp warning system. In this system low pressure plumbing lines are routed from the engine bleed air distributor manifold beneath the cabin floorboards into each wheel well, where the lines are terminated in soft plastic tubing plugged at the end to contain pressure. The soft plastic tubing is routed in close proximity to the main engine bleed air lines which provide hot air to deice the brakes. If overheating of the brake deice system should occur, the soft plastic lines will melt, thus relieving the pressure in the warning system lines. As pressure is relieved, a pressure switch, which is tapped off of the soft plastic line in each wheel well, will activate yellow L BK DI OVHT and/or R BK DI OVHT annunciators in the CAUTION/ADVISORY panel, thereby warning the operator that the brake deice system should be turned OFF.

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Figure 1 Brake Deice System

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UE32B 990888AA.AI

Figure 2 Brake Deice System (System On)

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LANDING GEAR BRAKE DEICE SYSTEM MAINTENANCE PRACTICES

200200

1. PROCEDURES WARNING: When performing maintenance on a hydraulically operated landing gear system, be aware that any movement of a hydraulic actuator cylinder can displace hydraulic fluid and cause unanticipated movement of other actuator cylinders in the landing gear retraction system. Servicing of the landing gear hydraulic accumulator can also result in unanticipated movement of an actuator. Either action can result in an unsafe, unlocked landing gear system. Therefore, place the airplane on jacks prior to performing any inspection or maintenance. Cycle the landing gear and ensure that all three landing gears are down and locked prior to removing the airplane from jacks.

2. DISTRIBUTOR MANIFOLD NOTE: Refer to 32-40-00 for the following maintenance practices: removal and installation of the master cylinders, parking brake valves, nose wheel, and main wheel and brakes; adjustment of the parking brake cable and the master cylinder linkage; brake wear limits; and cleaning of the pressure equalization filter and orifice.

A. Removal (1) Remove the main landing gear wheels (Ref. 32-40-00). (2) Disconnect brake hydraulic line (1) from the inlet swivel fitting (2) and cap the line and fitting (Ref. Figure 201). (3) Disconnect the bleed air deice hose (6) from the journal on the distributor manifold (9). (4) Remove the eight bolts (11), washers, and nuts securing the distributor manifold to the outboard manifold support (10). (5) Remove the bolts (13), washers, and nuts securing the two brake assemblies to one another. (6) Remove both brake assemblies from the landing gear strut while being careful not to drop or damage the inlet swivel fitting (2). (7) The manifold supports (8) and (10) can be removed from the brake assemblies by removing the five bolts (14) and self-locking nuts (15).

B. Installation NOTE: Ascertain that the orifices in the distributor manifold are clean and unobstructed prior to installation of the manifold. (1) Attach the manifold supports (8) and (10) to the brake assemblies with the five bolts (14) and self-locking nuts (15). At the time of installation, lubricate the bolt and nut threads and the bearing surfaces of the bolts and nuts with antiseize thread compound (76, Table 1, Chapter 91-00-00). If P/N 5006749 basic, -2 or -3 brakes are installed, torque the self-locking nuts to 120 inch-pounds. If P/N 5006749-4, -5 or -5R brakes are installed, torque the self-locking nuts to 300 inch-pounds (Ref Figure 201).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL NOTE: If the inlet adapters (5) are replaced, use new packings (3). (2) Position the inboard brake assembly (7) on the shock strut. Apply a thin coat of grease (same as that used on the wheel bearings) to the inside diameter of the brake housing and on the axle at assembly. (3) Lubricate the new packings (4) with hydraulic fluid (39, Table 1, Chapter 91-00-00) and install the packings in the grooves in the inlet swivel fitting (2). Lightly lubricate the surfaces next to the grooves. (4) Install the inlet swivel fitting on the inlet adapter (5) of the inboard brake assembly, being careful not to displace or damage the packing (4) (Ref. Figure 201). (5) Position the outboard brake assembly (12) on the shock strut. Apply a thin coat of grease (same as that used on the wheel bearings) to the inside diameter of the brake housing and on the axle at assembly. Carefully mate the inlet adapter (5) on the outboard brake with the inlet swivel fitting (2), being careful not to displace or damage the packing (4). (6) Ascertain that both brake assemblies are correctly seated on the strut and axle and that the inlet swivel fitting (2) swivels freely. (7) Secure the brake assemblies together with the attaching bolts (13), washers, and nuts; install one washer under each bolt head and one washer under each nut. At the time of installation, lubricate the bolt and nut threads and the bearing surfaces of the bolts, washers, and nuts with antiseize thread compound (76, Table 1, Chapter 91-00-00). Torque the self-locking nuts to 120 inch-pounds. (8) Install the eight bolts (11), washers and nuts securing the distributor manifold to the outboard manifold support. (9) Connect the hydraulic brake fluid line (1) to the inlet swivel fitting (2). (10) Connect the bleed air deice hose (6) to the distributor manifold. (11) Install the main landing gear wheels (Ref. 32-40-00).

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Figure 201 Brake Deice Distributor Manifold

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3. SHUTOFF VALVE A. Removal (1) Working inside the main gear wheel well, disconnect the electrical wiring to the shutoff valve solenoid and to the annunciator light switch. Tag and identify the wires. (2) Remove the tube from the inlet side of the shutoff valve and the bleed air duct forward of the firewall. (3) Remove the tube from the outlet side of the shutoff valve. (4) Remove the nut that secures the shutoff valve to the firewall and remove the shutoff valve from the firewall. NOTE: When removing the bleed air tubes and the shutoff valve, do not break the torque paint on the inlet and outlet fittings installed in the shutoff valve.

B. Installation (1) Install the shutoff valve through the firewall and secure the valve to the firewall with the nut. (2) Install the tube between the inlet side of the shutoff valve and the bleed air duct forward of the firewall. (3) Connect the tube to the outlet side of the shutoff valve. (4) Connect the electrical wiring to the shutoff valve solenoid and to the annunciator light switch. NOTE: When installing the bleed air tubes and the shutoff valve, do not break the torque paint on the inlet and outlet fittings installed in the shutoff valve.

C. Pressure Test and Functional Test Any time maintenance or replacement of components in the system is performed, the system should be pressure and/or functionally tested as follows: (1) Pressure test all plumbing connections, unions, lines, etc. from the solenoid-operated shutoff valve to the manifold on each main landing gear. (2) Disconnect the deice plumbing at the wheel deice manifold and cap the line. (3) Disconnect the line from the engine to the solenoid-operated shutoff valve at the valve. (4) Connect an air supply regulated at 80 psi to the shutoff valve. (5) Turn the battery master switch, bus tie switch, and brake deice switch ON. CAUTION: Leave the weight of the airplane on the wheels when the battery master switch is turned ON. (6) Apply a leak check solution (mild detergent and water) to the connections and fittings. (7) No leaks are allowed, except a minute leak in the valve.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (8) Turn the brake deice switch OFF and ON to determine if the solenoid is functioning properly. (9) Turn the brake deice, bus tie, and battery master switches OFF. (10) Let the pressure drop to zero, remove the air supply and reconnect the engine bleed air line to the solenoid-operated shutoff valve. (11) Remove the cap from the line at the brake deice manifold and reconnect the line to the manifold.

D. Overtemperature Warning Functional Test (1) Gain access to the bleed air manifold under the center aisle floorboard, forward of the main spar. (2) Disconnect one of the brake deice warning lines from the manifold and connect an air supply regulated to 30 ± 2 psi to the warning line. (3) Turn off the air, thereby trapping the 30 psi in the line. There should be no loss of pressure in 2 minutes. (4) Turn ON the battery master, bus tie, brake deice switches. CAUTION: Leave the weight of the airplane on the wheels when the battery master switch is turned ON. (5) The green L BK DI ON or R BK DI ON annunciator in the CAUTION/ADVISORY panel should illuminate. (6) Allow the pressure to bleed off. The yellow L BK OVHT or R BK OVHT annunciator in the CAUTION/ADVISORY panel should illuminate when the pressure drops to 1.5 ± 0.5 psi. (7) Turn OFF the battery master, bus tie, and brake deice switches. (8) Disconnect the air supply and connect the warning line to the manifold. (9) Repeat Steps (1) through (8) on the other warning line. (10) Install the access covers and any other items removed for this check.

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LANDING GEAR MECHANICAL STEERING DESCRIPTION AND OPERATION

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1. GENERAL WARNING: When an airplane has experienced abnormal landing gear procedures of any type, as a safety precaution, place the airplane on jacks prior to performing any inspection or maintenance. Ensure that all three landing gears are down and locked prior to removing the aircraft from jacks. CAUTION: Jacking of an airplane for the purpose of landing gear operation, inspection, servicing or maintenance, should be accomplished within an enclosed building or hangar. In the interest of safety, should it become necessary to jack the airplane in the open, wind velocity in any direction and terrain variations, must be compensated for prior to jacking the airplane. The nose wheel steering mechanism, which is mounted on the LH side of the nose wheel well, provides direct linkage to the rudder pedals, allowing the nose wheel to be turned 15° to the left of center and 15° to the right. A spring mechanism in the steering linkage dampens the transmission of excessive shock loads to the rudder pedals and permits tow bar or braked steering angles greater than that provided by the rudder pedal system. A straightener bracket on the airplane structure and the straightener guide on the steering bellcrank automatically centers the nose gear wheel when the landing gear is extended or retracted. To remove the steering action from the rudder pedals when the airplane is in flight, a steering disconnect actuator, disconnect link and disconnect cam are incorporated into the steering mechanism. The aft end of the disconnect link is connected to the disconnect cam with a bearing that travels in a slot in the cam. The actuator, which is attached to the disconnect link, moves the bearing up and down in the slot as the actuator is extended and retracted. Switches in the actuator limit the travel of the actuator. Power for the electric disconnect actuator is drawn from a 5-ampere NOSE GEAR STEER circuit breaker located on the RH circuit breaker panel. When the weight is removed from the LH main gear strut, the safety switch is actuated to complete a circuit to the extend circuit of the actuator. As the actuator is extended, the aft end of the steering disconnect link follows the slot in the disconnect cam, rendering the steering mechanism unaffected by the rudder pedals. When weight is placed on the LH landing gear strut, the safety switch is actuated to complete a circuit to the actuator retract circuit. As the actuator is retracted the aft end of the steering disconnect link follows the slot in the disconnect cam, making the nose wheel steering operable through the rudder pedals.

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LANDING GEAR MECHANICAL STEERING MAINTENANCE PRACTICES

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1. PROCEDURES WARNING: When performing maintenance on a hydraulically operated landing gear system, be aware that any movement of a hydraulic actuator cylinder can displace hydraulic fluid and cause unanticipated movement of other actuator cylinders in the landing gear retraction system. Servicing of the landing gear hydraulic accumulator can also result in unanticipated movement of an actuator. Either action can result in an unsafe, unlocked landing gear system. Therefore, place the airplane on jacks prior to performing any inspection or maintenance. Cycle the landing gear and ensure that all three landing gears are down and locked prior to removing the aircraft from jacks.

2. MECHANICAL STEERING MECHANISM A. Removal (1) Aft steering link (20) (Ref. Figure 201): (a) Remove the nut, washers and bolt attaching the forward end of the aft link to the steering disconnect cam (16). (b) Remove the eight studs securing the boot and ring to the aft wheel well panel. (c) Remove the nut, washers and bolt attaching the aft end of the aft link to the steering torque arm (21). (2) Steering disconnect link (14): (a) Remove the nut, washers, spacers and bolt attaching the disconnect actuator (11) to the disconnect link. (b) Remove the nut, washer, bearing and bolt attaching the aft end of the disconnect link to the disconnect cam (16). (c) Remove the nut, washer and bolt attaching the forward end of the disconnect link to the idler assembly (10). (3) Forward steering link (9 or 9A): (a) Remove the nut, washer and bolt attaching the aft end of the forward link to the idler assembly (10). (b) Remove the nut, washer and bolt attaching the forward end of the forward link to the steering bellcrank (5). (4) Steering disconnect cam (16): (a) Remove the four bolts securing the cam support assembly (18) to the LH wheel well keel. To gain access to the nuts, remove the nose baggage compartment floor.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (b) Remove the nut, washers, grip bushing (17) and bolt attaching the disconnect cam to the cam support assembly (18). (5) Idler assembly (10): (a) Remove the nut, washers, grip bushing and bolt attaching the idler assembly to the LH wheel well keel; the nut is accessible through the nose baggage compartment floor. (6) Shimmy damper (8): (a) Remove the nut, washer and bolt attaching the shimmy damper piston rod end to the steering bellcrank (5). (b) Remove the nut, washer, grip bushing and bolt securing the shimmy damper to the RH wheel well keel. (7) Steering bellcrank (5): (a) Remove the nut, washers, grip bushing and bolt attaching the steering bellcrank to the airplane structure. (b) The straightener guide (4) can be removed from the bellcrank after removal of the two attaching bolts.

B. Installation (1) Steering bellcrank (5) (Ref. Figure 201): (a) Secure the straightener guide (4) to the bellcrank with the two bolts, nuts, and washers. (b) Attach the bellcrank to the airplane structure with the bolt, grip bushing (7), washers, and nut. Install 100951-X-032-XH washers (6) up to a maximum of three to provide 0.09-inch clearance between the straightener guide and the straightener bracket (3) when the nose wheel is rotated as far as possible to the left and to the right. (2) Shimmy damper (8): (a) Secure the shimmy damper to the RH wheel well keel with the nut, washer, bushing and bolt. (b) Attach the shimmy damper piston rod end to the steering bellcrank with the bolt, washer, and nut. (3) Idler assembly (10): (a) Secure the idler assembly to the LH wheel well keel with the bolt, grip bushing, washers and nut. Remove the baggage compartment floor to install the nut. (4) Steering disconnect cam (16): (a) Attach the cam to the cam support assembly (18) with the bolt, grip bushing (17), washers and nut. Install AN960-516L and/or AN960-516 washers (19) under the nut and washer as required to obtain 0.01 to 0.04-inch protrusion of the grip bushing (17) through the washers (19).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (b) Secure the cam support assembly (18) to the LH wheel well keel with the four bolts, washers, and nuts. Remove the nose baggage compartment floor to install the nuts on the bolts. (5) Forward steering link (9 or 9A): (a) Install the bolt, washer and nut attaching the forward end of the forward link to the steering bellcrank (5). (b) Install the bolt, washer and nut attaching the aft end of the forward link to the lower hole in the idler assembly (10). NOTE: The aft end of the forward link is swaged. (6) Steering disconnect link (14): (a) Connect the forward end of the disconnect link to the upper hole in the idler assembly (10) with the bolt, washer, and nut. (b) Connect the aft end of the disconnect link to the disconnect cam (16) with the bearing, bolt, washer and nut. (c) Attach the disconnect actuator (11) to the disconnect link with the bolt, spacers, washers and nut. To center the actuator clevis to the disconnect link, install AN960-416 or AN960-416L washers (13) as required between the spacers (12) and the actuator clevis. A minimum of one washer should be placed on each side of the clevis. The maximum allowed end play in this joint is 0.03-inch. (7) Aft steering link (20): (a) Install the bolt, washers, and nut attaching the aft end of the aft link to the steering torque arm (21). (b) Install the eight studs securing the boot and ring to the aft wheel well panel. (c) Attach the forward end of the aft link to the disconnect cam assembly (16) with the bolt, washers and nut.

3. DISCONNECT ACTUATOR A. Removal (1) Disconnect the actuator wiring at the receptacle plug located in the LH wheel well keel. (2) Remove the nut, washers, spacers, and bolt attaching the disconnect actuator (3) to the disconnect link (6) (Ref. Figure 202). (3) Remove the bolt, nut, and washers that attach the actuator to the actuator support bracket (1) and remove the actuator.

B. Installation NOTE: Verify that the actuator support bracket (1) is installed with the actuator attach hole toward the aft end of the bracket (Ref. Figure 202).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (1) Position the actuator and install the bolt, washers, and nut that attach the actuator to the actuator support bracket. Center the actuator with the actuator support bracket by installing AN960-10 or AN960-10L washers (2), as required, between the support bracket (1) and the actuator. A minimum of one washer should be placed on each side of the actuator support bracket. The maximum allowed end play in this joint is 0.03-inch. (2) Attach the disconnect actuator to the disconnect link (6) with the bolt, spacers, washers and nut. To center the actuator clevis to the disconnect link, install AN960-416 or AN960-416L washers (5) as required between the spacers (4) and the actuator clevis. A minimum of one washer should be placed on each side of the clevis. The maximum allowed end play in this joint is 0.03-inch. (3) Connect the actuator wiring at the receptacle plug located in the LH wheel well keel.

4. FORWARD STEERING LINK BOOT A. Removal (1) Remove the forward steering link. Refer to Step (3) of the MECHANICAL STEERING MECHANISM REMOVAL, in this section. (2) Remove two clamps (25) holding boot (26) onto the forward steering link (9A) and slide the boot off one end of the steering link (Ref. Figure 201).

B. Installation (1) Ensure spring (27) is liberally lubricated with grease (61, Table 1, Chapter 91-00-00) and slide a boot (26) over the steering link (9A) and spring (27) (Ref. Figure 201). (2) Adjust the boot ends (26) to seal on the collars (28) and install the two clamps (25). (3) Install the forward steering link. Refer to Step (5) of the MECHANICAL STEERING MECHANISM INSTALLATION, in this section. (4) Perform MECHANICAL STEERING NOSE GEAR CENTERING ADJUSTMENT, in this section.

C. (With Boot) Inspection Inspect and lubricate the forward steering link at the interval specified in Chapter 05-00-00. (1) Perform FORWARD STEERING LINK BOOT REMOVAL, in this section. (2) Inspect boot (26) for cracks, cuts, or excessive wear and replace if damaged (Ref. Figure 201). (3) Clean the spring (27) and forward steering link (9A) with solvent (2, Table 1, Chapter 91-00-00), visually inspect and replace worn or damaged parts. (4) If complete disassembly is required, remove spring pins (29), remove collars (28), remove spring (27), loosen jam nuts (30) and unscrew rod ends (31). Assemble in reverse order. (5) Perform FORWARD STEERING LINK BOOT INSTALLATION, in this section.

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5. MECHANICAL STEERING NOSE GEAR A. Centering Adjustment (1) With the airplane on jacks, connect an external power supply to the airplane and adjust the voltage to 28 ± 0.25 volts. (2) Check the three steering links for the following dimensions: aft steering link, 12.54-inches; steering disconnect link (4), 18.00-inches; forward steering link (3), 29.95-inches (Ref. Figure 203). Adjust the link lengths, if necessary, by rotating each rod end in or out as required to obtain the correct length between the centers of the holes in the rod ends of the respective links. (3) With the rudder pedals in the neutral position and the steering disconnect actuator in the extended position, slowly retract the landing gear and observe the straightener roller (2) on the landing gear as it makes contact with the straightener bracket (1) on the airplane structure. During the retraction of the gear, the straightener roller should travel into the center slot of the straightener bracket without contacting the ramp portion of the bracket. (4) Adjust the length of the forward steering link (3) as required to center the roller in the straightener bracket. If the roller cannot be completely centered as stated in this Step, proceed to Step (5) If complete centering is accomplished by adjusting the forward steering link, proceed to Step (6). (5) With the rudder pedals in the neutral position and the steering disconnect actuator in the retracted position, slowly retract the landing gear and observe the straightener roller (2) on the landing gear as it makes contact with the straightener bracket (1) on the airplane structure. During the retraction of gear, the straightener roller should travel into the center slot of the straightener bracket without contacting the ramp portion of the bracket. Adjust the length of the aft steering link (5) as required to center the roller in the straightener bracket. (6) Disconnect the external power supply and remove the airplane from the jacks.

B. Stop Removal (1) Remove cotter pins (1), nuts (2), washers (3) and bolts (4) securing the steering stop (6) to the nose gear brace (refer to Figure 4). (2) Remove the steering stop (6) and shims (5). Perform MECHANICAL STEERING NOSE GEAR STOP INSTALLATION, in this section.

C. Stop Installation (1) Position the steering stop (6) and shims (5) between the nose gear brace lugs (Ref. Figure 204). NOTE: Install shims (5) as required on each side to center the steering stop and maintain a total clearance of 0.000 to 0.002-inch between stop and brace. (2) Attach the steering stop (6) to the nose gear brace with bolts (4), washers (3), nuts (2) and cotter pins (1). Tighten nut and bolt assembly snug (finger tight) using a minimum of one washer to allow installation of cotter pin.

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Figure 201 Nose Gear Steering Mechanism

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Figure 202 Steering Disconnect Actuator

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Figure 203 Mechanical Steering Nose Gear Centering Adjustment

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Figure 204 Mechanical Steering Nose Gear Stop

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LANDING GEAR POWER STEERING (UA-1 AND AFTER; UB-1 AND AFTER) DESCRIPTION AND OPERATION

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1. GENERAL WARNING: When an airplane has experienced abnormal landing gear procedures of any type, as a safety precaution, place the airplane on jacks prior to performing any inspection or maintenance. Ensure that all three landing gears are down and locked prior to removing the aircraft from jacks. CAUTION: Jacking of an airplane for the purpose of landing gear operation, inspection, servicing or maintenance, should be accomplished within an enclosed building or hangar. In the interest of safety, should it become necessary to jack the airplane in the open, wind velocity in any direction and terrain variations, must be compensated for prior to jacking the airplane. The power steering system provides two steering modes, a ± 10° taxi mode and a ± 63° parking mode, both of which are selected by the control switch located on the pedestal and placarded OFF - TAXI - PARK. Rotation of the nose gear shock absorber for steering purposes is performed by a rotary actuator mounted on top of the nose gear assembly. The actuator consists of a two-position solenoid arming valve, a servo valve, a pressure-actuated selector valve, and two pistons. A pump and motor assembly located in the LH wheel well drives the actuator hydraulically. The power steering pump uses the primary reservoir of the landing gear power pack as a fluid supply. The nose landing gear normal extend line is utilized as the pressure line to the power steering actuator. A two-position solenoid valve mounted just outboard of the power pack is plumbed into the nose gear extend line to block the entry of the power steering system pressure fluid into the power pack. A tee located in the nose gear extend line in the nose wheel well directs the fluid to the pressure port of the actuator. Until it is energized, the actuator arming valve prevents the entry of the pressure fluid into the actuator. Return fluid from the power steering actuator enters the emergency extension hand pump suction line, returning the fluid to the power pack (Ref. Figures 1 and 2). CAUTION: Do not operate the power steering pump without having the landing gear power pack reservoir pressurized to 18 psi with regulated dry air. Command potentiometers, mounted on the pilot's rudder pedal assembly, transmit the pilot's steering input to the power steering signal amplifier located under the copilot's seat. Nose gear follow-up potentiometers, located above the nose gear assembly, monitor the nose gear steering action and relay this information to the power steering signal amplifier. The signal amplifier then transmits electrical signals to the actuator servo valve where the appropriate orifices and valves are closed or opened to achieve the correct steering direction as commanded by the rudder pedal steering input. The power steering control circuit is protected by a five ampere circuit breaker located on the RH circuit breaker panel. Power for the pump motor is supplied through the power steering motor relay and 35-ampere limiter, both of which are located in the LH nacelle electrical power distribution panel. The motor relay, two position solenoid valve, and actuator arming valve are energized simultaneously by current from the control relay. The control relay is energized by the power steering control switch and is grounded through the locking relay and the power steering activation switch located on the LH power lever. The locking relay will retain the ground circuit through the steering disconnect switches as long as either power lever is retarded below an N1 speed of 89%. Once the power levers are advanced beyond an N1 speed of 91% the control relay will be grounded only while the steering activation switch is actuated. The LH main gear downlock switch and safety switch will open the control circuit, nullifying the steering system, when the airplane leaves the ground. A nose gear position switch mounted on the nose gear straightener link (on airplane prior to S/N UB-68) will prevent retraction of the landing gear if the nose wheel is not centered. This switch will not be present on airplane that have complied with Optional Service Bulletin 2191.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL When the power steering control switch is in the OFF position or in the TAXI or PARK position without the power steering activation switch being actuated, the steering system is inoperative and the steering actuator will be in the caster mode functioning as a shimmy damper. When the power steering control switch is in the PARK position, the steering system provides a ± 63° steering mode when the steering activation switch is actuated, on the condition that at least one power lever is retarded below an N1 speed of 89%. Once both power levers are advanced beyond an N1 speed of 91%, the control switch solenoid will no longer hold the switch in the PARK position and it will move to the TAXI position, with the only steering available being ±10° steering while the power steering activation switch is actuated. The ± 63° steering mode can be reactivated by retarding one or both power levers below an N1 speed of 89% and placing the control switch in the PARK position. When the control switch is in the TAXI position, the steering system provides a ± 10° steering mode when the steering activation switch is actuated. A yellow MAN STEER FAIL annunciator in the CAUTION/ADVISORY panel illuminates if the actuator selector valve is in the power steering mode with: (a) The airplane on the ground and the control switch OFF. (b)

The airplane on the ground with the control switch in the TAXI or PARK position without the steering activation switch actuated.

(c) The airplane in flight and the control switch in the OFF or TAXI position with landing gear extended. (d) The airplane in flight and the control switch in the TAXI position with landing gear retracted.

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Figure 1 Power Steering System (UA-1 and After; UB-1 and After)

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UE32B 990406AA.AI

Figure 2 Power Steering System Schematic (UA-1 and After; UB-1 and After)

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LANDING GEAR POWER STEERING (UA-1 AND AFTER; UB-1 AND AFTER) TROUBLESHOOTING 1. PROCEDURES WARNING: When an airplane has experienced abnormal landing gear procedures of any type, as a safety precaution, place the airplane on jacks prior to performing any inspection or maintenance. Ensure that all three landing gears are down and locked prior to removing the aircraft from jacks. CAUTION: Jacking of an airplane for the purpose of landing gear operation, inspection, servicing or maintenance, should be accomplished within an enclosed building or hangar. In the interest of safety, should it become necessary to jack the airplane in the open, wind velocity in any direction and terrain variations, must be compensated for prior to jacking the airplane. The 1900 Airliner power steering system utilizes a number of primary, secondary and tertiary controls which are interfaced to provide an efficient and highly dependable system. A fault detection system provides for monitoring the response of the nose wheel to inputs from the flight crew when the power steering is in use. Fault sensing is accomplished by the power steering signal amplifier through which all command and monitoring information is processed. The various amplifier inputs and outputs have very predictable characteristics in the operating system. It is possible to access these inputs and outputs by using a special breakout unit. The predictable nature of the amplifier signals makes it possible to use these values to troubleshoot faults in the system. The breakout unit provides parallel access to the amplifier signals. Refer to Charts in Figures 102, 103, 104, 105 and 106.

A. Troubleshooting Notes (UA-1 and After; UB-1 and After) Prior to beginning any troubleshooting of the power steering system, the airplane must be placed in the following state: (1) Nose wheel jacked clear of the ground. (2) Nose wheel and rudder pedals in the neutral position. (3) Breakout unit connected between the amplifier and the airplane harness. (4) 28-vdc regulated APU connected to airplane and turned on. (5) Battery switch turned on. (6) External power switch turned on. (7) Main gear on the ground or squat switches bypassed. Refer to Wiring Diagram Manual for correct switch location. (8) References to the following notes are found on the troubleshooting Charts in this section: 1. Remove the low pressure switch that is located on the LH side of the Nose Gear Wheel Well and jumper the leads from the switch. Install in the line a pressure gage capable of measuring 3000 psi. Engage the power steering system and monitor the system pressure until the pressure meets or exceeds 2000 psig.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL 2. Disconnect the extend line from the landing gear power pack and engage the power steering system; there should be no fluid returning from the extend line if the two-position valve is operating properly. 3. System calibration is to be found in this section. Refer to POWER STEERING CALIBRATION. 4. Check the servo solenoid coils with an ohmmeter or continuity checker for open or shorted condition. 5. Continuity between pins A and B of the selector valve receptacle (J3) will confirm that the arming valve is open. 6. Diodes may be checked by using an ohmmeter or a continuity checker; continuity should be indicated in one direction only. In the event one of these diodes is shorted, damage may have resulted to the amplifier. 7. The power steering can only be engaged with the switch in the TAXI or PARK position, then momentarily pressing the activation switch on the power lever. 8. Restrictor fittings installed in the valve body inlet and return ports are color coded. The inlet restrictor is brown and the return restrictor is black. Failure to install the restrictors in the proper ports will render the landing gear inoperative (UB-1 thru UB-22).

B. Signal Amplifier (UA-1 and After; UB-1 and After) Command and monitoring logic and functional testing of the fault detection circuits are all functions of the power steering signal amplifier. Inputs from the command and nose gear follow-up potentiometers are compared by way of a Wheatstone bridge circuit. When the amplifier senses an imbalance between the sides of the bridge circuit, the nose gear actuator is driven one direction or the other in an attempt to rebalance the bridge. Similarly, the two monitor potentiometers are compared to one another by the amplifier. When an imbalance is sensed between the monitor potentiometer circuit and the control potentiometer circuit, the amplifier will interrupt ground to the controlling relays at pin 1 of the amplifier, and the system will be deactivated. On earlier model amplifiers the imbalance is sensed with 3° difference between rudder pedal position and nosewheel position. On airplane with the amplifier, P/N 50700-5, a 7° difference will deactivate the system. Inputs from the test switch bypass the command potentiometer and drive the nose gear actuator to the right or left to create a deliberate imbalance between the monitor and control circuits. This imbalance is then sensed by the amplifier, which activates the fault detection circuitry. Pin 16 inputs the test signal while pins 26 and 31 provide the signal outputs. The amplifier provides an output from pin 28 which, when applied back into the amplifier at pin 29, will reset the nose gear maximum travel circuitry to a greater maximum travel so the airplane can be more easily maneuvered for parking. Operating current to the amplifier is input at pin 20 and grounding is effected at pin 22. Pins 37 and 19 are discrete outputs which operate the servo valve in the actuator. Two adjustable potentiometers, accessible through the case of the amplifier, can be used for making final adjustments for centering of the nose gear. One potentiometer is labeled CENTER and the other is labeled MONITOR. Any time an adjustment is made to the centering potentiometer, an equal adjustment must be made to the monitoring potentiometer.

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C. System Control (UA-1 and After; UB-1 and After) When the Power steering control switch is in the OFF position, current from the 5-ampere power steering control circuit breaker arms the MAN STEER FAIL annunciator circuit through the RH gear downlock switch. When the control switch is placed in the TAXI position, current to the MAN STEER FAIL annunciator circuit is shunted around the RH gear downlock switch through the normally closed contacts of the control relay. Control current is then fed through the LH gear safety switch to power the ground latching relay and the control relay; however, neither relay is energized in this state since their ground circuits are not complete. Current through the LH gear downlock switch is supplied to the fault indicator relay, closing it. In the event the fault indicator relay does not close as a result of the amplifier removing ground from the relay, the current will be fed through the normally closed contacts of the fault indicator relay to the fault indicator time delay module and PWR STEER FAIL annunciator. The time delay module must complete a two second delay before energizing the annunciator light. This same current is also fed to the amplifier as operating power. The system is armed in this state, but not operating. When the power steering activation switch is actuated, the ground circuit to the ground latching relay is completed and the relay closes, latching to ground and simultaneously supplying ground to the control relay which closes in response. When the control relay closes, ground is applied to the arming valve and the servo valve circuits. At the same time, current which had been made available to the MAN STEER FAIL annunciator circuit is now switched by the control relay to energize the two-position solenoid valve, the pump motor power relay, the arming valve and the low pressure switch circuit and, if the low pressure switch is closed, the fault indicator time delay module and the PWR STEER FAIL annunciator after the appropriate time delay. Placing the control switch in the PARK position applies the park signal to the amplifier which activates the park circuitry after a four second time delay. In this state, should the power levers be advanced enough to open the park disconnect switches, the park signal circuit will be interrupted and the control switch will return to the TAXI position. When the park disconnect switches open, the steering disconnect switches are opened as well, thereby removing ground from the ground latching relay, and opening it, which will open the grounding circuit to the control relay, opening it. The fault indicator relay will continue to be energized. When the control relay opens, the ground signals will be removed from the arming valve and the servo valve. Power will be removed from the two-position solenoid valve, the pump motor relay, the arming valve and the low pressure switch circuit. Any time the amplifier senses a fault and system operation is interrupted, the nose gear always returns to the caster mode and taxi is accomplished by differential braking.

D. Breakout Unit (UA-1 and After; UB-1 and After) A breakout unit, used to access amplifier inputs and outputs, can be built in the shop from readily available materials (Ref. Figure 101). Each jack on the face of the unit corresponds to a pin of the connector on the power steering amplifier unit. Each wire in the length of cable which has the plug attached to it is soldered to the correspondingly numbered jack and then attached to the receptacle on the side of the unit. When the unit is completed and the unit's plug is connected to the amplifier and the airplane harness is connected to the breakout unit, the breakout unit is in parallel with the airplane harness and, therefore, in parallel with any signals carried through the breakout unit. A digital voltmeter is recommended for use when troubleshooting the power steering system. Only a digital voltmeter should be used when adjusting the potentiometers. The negative lead from the DVM should be connected to jack number 22, since this is the only pin of the amplifier which has a continuous ground.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL The troubleshooting Charts which follow are to be used as a guide for the logical flow of activity during the troubleshooting procedure. All possible faults or pilot complaints and observations were considered when building these troubleshooting Charts, and it is anticipated that all faults will be discovered during troubleshooting if the Charts are followed to conclusion.

E. System Calibration (UA-1 and After; UB-1 and After) Should troubleshooting reveal that an out-of-calibration condition has caused the fault in a system, the command and monitoring potentiometers must be adjusted. The following procedure is the only procedure recommended for calibration of the command and monitoring potentiometers. (1) Jack the nose gear off the ground. (2) Connect a regulated 28-vdc APU to the airplane. (3) Connect the breakout unit between the amplifier and the airplane harness. (4) The nose gear and the rudder pedals must be in the neutral position. (5) Turn on the battery and master switches and the external power switch, then place the power steering switch in TAXI. (6) Measure the voltage between amplifier pins 33 and 25 and between pins 34 and 23; these voltages must be equal to 24 ± 0.2 vdc. (7) Divide this voltage by 2; this voltage ± 0.1 vdc is the target voltage for adjusting the potentiometers. (8) Connect a digital voltmeter between test points 14 and 34 of the breakout unit. (9) Loosen the three clamping screws at the base of the potentiometer assembly. (10) Rotate the command potentiometer assembly until the voltmeter reading is equal to the target voltage ± 0.1 vdc. (11) Tighten the three clamping screws. (12) Connect the voltmeter between test points 6 and 36 and read this voltage. (13) If adjustment is necessary, loosen the screw on the band clamp which holds the two potentiometers together. (14) Rotate the monitor potentiometer until the voltmeter reading is equal to the target reading ± 0.1 vdc. (15) Tighten the band clamp screw securing the potentiometers. (16) Recheck the adjusted voltages to ensure that the potentiometers did not move during the tightening procedure. (17) Connect the voltmeter between test points 17 and 33 of the breakout unit. (18) Working in the nose gear wheel well, adjust this potentiometer assembly until the target voltage is established and tighten the clamping screws.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (19) Connect the voltmeter between test points 4 and 35. (20) If adjustment is necessary, loosen the band clamp screw and adjust the monitor potentiometer to the target voltage. (21) Tighten the band clamp screw and recheck the voltages. (22) Should the nose gear not be centered, the center adjusting potentiometer on the amplifier may be adjusted to effect proper centering of the nose gear; however, be certain to note the number and direction of turns, as an equal adjustment must be made to the monitor adjusting potentiometer on the amplifier for proper balancing of the monitoring circuitry. (23) Operate the nose gear in the park mode and check for proper nose gear travel limits; the nose gear should travel smoothly and freely between the red travel limit marks on the nose gear strut. NOTE: Should the rudder pedals be moved slowly, some incremental movement of the nose gear may be noticed; however, this is a normal characteristic of the system.

Figure 101 Power Steering Breakout Unit (UA-1 and After; UB-1 and After)

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Figure 102 Troubleshooting - Power Steering Not Operating Properly (7) (UA-1 and After; UB-1 and After)

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Figure 103 Troubleshooting - PWR STEER FAIL Annunciation (7) (UA-1 and After; UB-1 and After)

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Figure 104 Troubleshooting - Power Steering Control (7) (UA-1 and After; UB-1 and After)

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Figure 105 Troubleshooting - Power Steering Valves (7) (UA-1 and After; UB-1 and After)

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Figure 106 Troubleshooting - MAN STEER FAIL Annunciation (7) (UA-1 and After; UB-1 and After)

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200200

LANDING GEAR POWER STEERING (UA-1 AND AFTER; UB-1 AND AFTER) MAINTENANCE PRACTICES 1. PROCEDURES WARNING: When performing maintenance on a hydraulically operated landing gear system, be aware that any movement of a hydraulic actuator cylinder can displace hydraulic fluid and cause unanticipated movement of other actuator cylinders in the landing gear retraction system. Servicing of the landing gear hydraulic accumulator can also result in unanticipated movement of an actuator. Either action can result in an unsafe, unlocked landing gear system. Therefore, place the airplane on jacks prior to performing any inspection or maintenance. Cycle the landing gear and ensure that all three landing gears are down and locked prior to removing the aircraft from jacks.

A. System Servicing (UA-1 and After; UB-1 and After) Since the power steering system is a continuous flow system using the landing gear power pack reservoir for the fluid source, servicing the system consists of maintaining the correct fluid level in the landing gear hydraulic system as instructed in 32-30-00.

B. System Filter Replacement (UA-1 and After; UB-1 and After) Replace the system filter every 100 hours for the first 300 hours of service, then every 600 hours. The filter is located on the LH keel of the nose gear wheel well.

C. Hydraulic Fittings Installation (UA-1 and After; UB-1 and After) Anytime a fitting is loosened or removed, discard the packing and install the fitting with a new packing. To prevent damage to the packing, coat the packing and the threads of the fitting with hydraulic fluid (39, Table 1, Chapter 91-00-00) before installing the packing on the fitting. Some fittings in the steering system employ both an packing and a ring. When this type of fitting is loosened or removed, both the packing and the ring must be replaced when the fitting is installed (Ref. Figure 201).

2. POWER STEERING ACTUATOR A. Maintenance Checks (UA-1 and After; UB-1 and After) (1) Turn power steering system OFF. (2) Disconnect plug P490 from the low pressure switch located in the nose wheel well. (3) Apply 28 vdc at P490 pin B to engage the pump and arm the actuator. (4) Observe that the MAN STEER FAIL annunciator is illuminated. If not, and the continuity of the system can be verified, replace the actuator. (5) Remove the 28 vdc applied to P490 pin B. (6) Disconnect the arming valve plug P491. (7) “Cold soak” the arming valve using best shop practices.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (8) Repeat Steps (1) through (3). (9) If the MAN STEER FAIL annunciator illuminates within 10 seconds, the arming valve is leaking and the actuator needs to be replaced.

B. Removal (UA-1 and After; UB-1 and After) (1) Remove the nose landing gear from the airplane (Ref. 32-20-00). (2) Loosen the nuts on the eccentric bolts (6) and (9) and turn the bolts to move the torque stop adjustment blocks (7) and (8) away from the torque stops (11) and (12) (Ref. Figure 202). (3) Disconnect the hydraulic lines from the actuator. Remove the brackets (15) and (16) securing the hydraulic lines to the mounting plate (5). CAUTION: As the hydraulic lines are disconnected from the actuator, plug or cap the openings to prevent entry of foreign material into the lines or actuator. (4) Remove the six bolts (1) attaching the straightener arm (2) and actuator to the nose gear assembly. (5) Remove the three bolts (3) securing the actuator to the mounting plate.

C. Installation (UA-1 and After; UB-1 and After) (1) Install the actuator to the mounting plate (5) with the three bolts (3) (Ref. Figure 202). (2) Install the actuator and mounting plate on the nose gear. When installing the mounting plate, position the slots in the plate over the torque stops (11 and 12). (3) Attach the actuator and straightener arm (2) to the nose gear with the six bolts (1). When installing the actuator and straightener arm (2) to the nose gear, align the straightener roller and index mark on the actuator with the nose gear wheel. NOTE: Use care when tightening the eccentric bolts to avoid altering the adjustments. (4) With the nut loosened turn the eccentric bolt (6) until the torque stop adjustment block (7) holds the mounting plate (5) solidly against the RH torque stop (11). Tighten the nut on the eccentric bolt. (5) After the RH torque stop adjustment block (7) is adjusted, adjust one of the LH blocks (8) to where it has minimal contact up to 0.001 inch clearance from the stop (11), then adjust the other block to have minimal contact up to 0.001 inch clearance. Tighten the nuts on the eccentric bolts (6) and (9). Torque to 225 to 250 inch-pounds. Check that the shock strut rotates in the nose gear brace without binding. (6) Attach the brackets (15) and (16) securing the hydraulic lines to the mounting plate (5) and connect the hydraulic lines to the actuator. (7) Install the nose landing gear in the airplane (Ref. 32-20-00). (8) Check the fluid level of the landing gear hydraulic system. (Ref. 32-30-00).

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Figure 201 Hydraulic Fitting Installation (UA-1 and After; UB-1 and After)

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D. Repairing (UA-1 and After; UB-1 and After) Should a fault in the actuator assembly develop, certain components of the assembly may be procured from Hawker Beechcraft Parts & Distribution (HBP&D). The following instructions allow for the disassembly and assembly of the actuator for the purpose of replacing the improperly operating component. The improperly operating component should be returned to Hawker Beechcraft Parts & Distribution (HBP&D) for reconditioning or overhaul. Disassembly of any component of the actuator assembly, beyond the level detailed in these instructions, is not recommended by Hawker Beechcraft Corporation (Ref. Figure 203). (1) Release the air pressure from the hydraulic system reservoir by pressing the manual bleed valve. (2) Retain all attaching hardware removed during disassembly of the actuator. (3) Perform the POWER STEERING ACTUATOR REMOVAL procedure.

3. SERVO VALVE AND MANIFOLD (UA-1 AND AFTER; UB-1 AND AFTER) A. Removal (1) Disconnect the two hydraulic lines from the manifold. (2) Remove the servo valve (1) from the actuator by removing the four internally wrenching retaining screws (2) (Ref. Figure 203). (3) Discard the four packing (3) between the selector valve and the restrictor manifold (4). (4) Remove the restrictor manifold from the manifold assembly (5) and discard the four packings located between the restrictor manifold and the manifold assembly. (5) Remove the manifold assembly from the actuator assembly (6) by removing the four internally wrenching attaching screws (7) and discard the two packings located between the manifold assembly and the actuator assembly.

B. Installation (1) Install two new packings (16) in the seal seats on the actuator assembly (6) (Ref. Figure 203). (2) Attach the manifold assembly (5) to the actuator assembly using the four internally wrenching screws (7). (3) Install four new packings (17) in the seal seats on the restrictor manifold (4). (4) Install the restrictor manifold on the manifold assembly (5). (5) Install four new packings (3) in the seal seats on the servo valve (1). (6) Install the servo valve on the restrictor manifold and secure the servo valve and restrictor manifold to the manifold assembly with the four internally wrenching screws (2). (7) Connect the two hydraulic lines to the pressure and return ports.

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4. ARMING SOLENOID (UA-1 AND AFTER; UB-1 AND AFTER) A. Removal (1) Remove the safety wire from the arming solenoid (Ref. Figure 203). (2) Remove the arming solenoid (8) from the swivel adapter (9) located in the manifold assembly. CAUTION: The solenoid piston located inside the solenoid attach fitting is matched to the solenoid and is not interchangeable with the piston of any other solenoid. The piston may fall out during removal of the solenoid. To avoid losing the arming valve (10) located inside the manifold assembly swivel adapter, place a piece of tape over the swivel adapter opening.

B. Installation (1) Ensure that the arming valve (10) is in place within the manifold swivel adapter (9) (Ref. Figure 203). (2) Install the arming solenoid (8) in the manifold assembly, taking care that the solenoid piston is in place within the attach fitting of the solenoid. (3) Rotate the swivel adapter until the solenoid wires and connector are oriented in the same direction as the wires and connector on the servo valve, then safety wire the solenoid to the restrictor manifold.

5. SWITCH (UA-1 AND AFTER; UB-1 AND AFTER) A. Removal (1) Remove the switch cover (11) from the manifold end plate (12) by removing the three internally wrenching retaining screws (13) (Ref. Figure 203). (2) Remove the two retaining bolts (14), nuts and washers which attach the switch assembly (15) to the switch mounting and remove the switch assembly. NOTE: Note the position of the switch actuator lever; the dimple in the arm of the switch actuator lever must ride on the switch button when the actuator lever is pressed.

B. Installation (1) Position the switch actuator lever on the switch assembly (15) so that the dimple in the actuator lever arm is contacting the switch button when the actuator lever is pressed (Ref. Figure 203). (2) Install the switch assembly so that the roller on the switch actuator arm is against the manifold spool (18) and secure the switch to the end plate (12) with the two bolts (14), nuts and washers. (3) Attach a continuity checker or ohmmeter to pins A and B of the switch assembly connector. (4) Adjust the switch toward the spool until the switch actuates (indicated by continuity through the switch contacts), then back the switch off until the switch just deactuates (indicated by a loss of continuity). (5) Tighten the two retaining nuts securing the switch in this position.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (6) Install the switch cover retaining screws (13), leaving the switch cover (11) off. CAUTION: Do not apply pressure to the actuator without installing the switch cover retaining screws as damage to the actuator may occur. (7) Pressurize the hydraulic fluid reservoir with shop air regulated to 18 psig before proceeding to the next Step. (8) Energize the arming solenoid and apply hydraulic pump pressure to the system, by placing the power steering control switch in TAXI or ARM and pressing the activate button. (9) Notice that the spool moves out slightly, causing the switch to actuate when the activate button is pressed. NOTE: When the switch is properly adjusted, continuity through the switch contacts will be indicated by the continuity checker or ohmmeter. Should the switch fail to actuate during this check, readjust the switch as before and repeat the continuity check. (10) After the switch is properly adjusted, install the switch cover.

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Figure 202 Power Steering Actuator Installation (UA-1 and After; UB-1 and After)

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Figure 203 Actuator Repairing (UA-1 and After; UB-1 and After)

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6. PUMP AND MOTOR (UA-1 AND AFTER; UB-1 AND AFTER) A. Removal (1) Disconnect the electrical wiring from the pump motor (Ref. Figure 204). (2) Remove the pressure and suction lines from the pump. (3) Remove the seep drain line (16) from the bottom of the pump and motor assembly. CAUTION: As the hydraulic lines are disconnected from the pump, plug or cap the openings to prevent entry of foreign material into the lines or pump. (4) While supporting the pump and motor assembly, remove the four bolts (19) and washers securing the assembly to the mounting bracket (17).

B. Installation (1) Position the pump and motor on the mounting bracket (17) and install the four bolts (19) and washers (Ref. Figure 204). (2) Connect the suction and pressure lines to the pump. If the unions (11) were removed from the pump, install new packings (12). (3) Connect the seep drain line (16) to the pump and motor. If the elbow (15) was removed, install a new ring (14) and packing (13).

7. SOLENOID VALVE (2-POSITION) (UA-1 AND AFTER; UB-1 AND AFTER) A. Removal (1) Disconnect the electrical wiring from the solenoid valve. (2) Disconnect the hydraulic lines from the tee (1) mounted in the valve (Ref. Figure 204). (3) Disconnect the gear-up line (9) from the valve. CAUTION: As the hydraulic lines are disconnected from the valve, plug or cap all openings to prevent entry of foreign materials into the valve or hydraulic lines. (4) Remove the two bolts (8) and washers securing the valve to the mounting bracket (4).

B. Installation (1) Position the valve on the mounting bracket (4) and install the two bolts (8) and washers (Ref. Figure 204). (2) If the tee (1) was removed from the valve, install the tee with a new ring (2) and packing (3) (Ref. Figure 207). (3) Connect the hydraulic lines to the tee. (4) Connect the gear-up line (9) to the valve. Install a new packing (6) if the union (7) was removed.

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8. SIGNAL AMPLIFIER (UA-1 AND AFTER; UB-1 AND AFTER) A. Removal (1) Remove the copilot's seat (Ref. Chapter 25-10-00, SEAT REMOVAL). (2) Remove the access panel from between the copilot's seat tracks (Ref. Chapter 06-50-00, FLOOR ACCESS PANELS). (3) Disconnect the electrical wiring from the amplifier. (4) Remove the four screws and washers securing the amplifier to the mounting panel.

B. Installation (1) Position the amplifier on the mounting panel and install the four screws and washers securing the amplifier. (2) Connect the electrical wiring to the amplifier. (3) Install the access panel between the copilot's seat tracks (Ref. Chapter 06-50-00, FLOOR ACCESS PANELS). (4) Install the copilot's seat (Ref. Chapter 25-10-00, SEAT REMOVAL).

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Figure 204 Power Steering Pump and Motor and Solenoid Valve (UA-1 and After; UB-1 and After)

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9. DISCONNECT AND PARK DISCONNECT SWITCH (UA-1 AND AFTER; UB-1 AND AFTER) A. Adjustment A park disconnect and a steering disconnect, cam-operated switch for each power lever is mounted on the pedestal structure. The switches are adjusted to open when the power levers are moved beyond an N1 speed of 89% to 91%. The airplane must be in flight when determining the 89% to 91% N1 speed position of each power lever, as the ram air effect will alter the N1 position indicated during ground operation. (1) With the airplane in flight, advance the power levers until an N1 speed of 89% to 91% is attained on each engine. Set both power levers to the same N1 speed. (2) Mark the position of the power levers on the pedestal with tape and land the airplane. (3) Locate the electrical plug on the LH forward side of the pedestal. Disconnect the plug and connect a continuity checker to pins 15B and 16B of the receptacle; connect another continuity checker to pins 17B and 18B of the receptacle. (4) Place the power steering control switch in the PARK position. (5) With the RH power lever forward of the mark on the pedestal and the LH power lever aft of the mark, there should be continuity through the LH set of switches as indicated by the continuity checkers. Move the LH power lever forward through the mark on the pedestal, both continuity checkers should show a loss of continuity at the mark on the pedestal. Now move the RH power lever aft of the mark and the LH power lever forward of the mark, and repeat the above check moving the RH power lever forward through the mark. NOTE: The power steering control switch must be in the PARK position while checking the switches for proper adjustment. If the switches do not function as described in Step (5), adjust them as follows: (a) Remove the instrument panel from the pedestal. (b) Remove the knobs from the pedestal levers. (c) Remove the screws securing the electroluminescent panel to the pedestal. (d) Transfer the tape marks on the electroluminescent panel to the pedestal structure under the panel and remove the electroluminescent panel from the pedestal. (e) With the RH power lever forward of its mark on the pedestal, align the LH power lever with its mark on the pedestal. Place the power steering control switch in the PARK position. (f) Loosen, but do not remove, the screws (1) securing the LH switch mounting bracket (2) to the pedestal and move the mounting bracket forward or aft until the switches just open with the LH power lever aligned with the mark on the pedestal. When the switch mounting bracket is in the correct position, tighten the screws (1) (Ref. Figure 205). (g) With the LH power lever forward of its mark on the pedestal, align the RH power lever with its mark on the pedestal. Place the power steering control switch in the PARK position.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (h) Loosen, but do not remove, the screws (3) securing the RH switch mounting bracket (4) to the pedestal and move the mounting bracket forward or aft until the switches just open with the RH power lever aligned with the mark on the pedestal. When the switch mounting bracket is in the correct position, tighten the screws (3). (i) Check the switch adjustment. Refer to Step (5). (j) Install the electroluminescent panel to the pedestal with the attaching screws. (k) Install the knobs on the pedestal levers. (l) Install the instrument panel on the pedestal. (6) Disconnect the continuity checkers from the electrical receptacle on the pedestal and connect the airplane wiring harness to the receptacle.

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Figure 205 Steering Disconnect and Park Disconnect Switch Adjustment (UA-1 and After; UB-1 and After)

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10. NOSE WHEEL CENTERING MECHANISM (UA-1 AND AFTER; UB-1 AND AFTER) A. Removal (1) Straightener Link. (a) Remove the nut, washer and bolt attaching the aft end of the link to the bracket on the left wheel well keel (Ref. Figures 206 and 207). (b) Remove the nut, washer and bolt attaching the forward end of the link to the steering bellcrank. (2) Steering Bellcrank (a) Disconnect the nose gear follow-up potentiometer at the steering bellcrank. (b) Remove the nut, washers, grip bushing and bolt attaching the steering bellcrank to the airplane structure. (c) The straightener guide can be removed from the bellcrank with removal of the two attaching bolts.

B. Installation (1) Steering bellcrank. (a) Secure the straightener guide to the bellcrank with the two bolts, nuts and washers (Ref. Figures 206 and 207). (b) Attach the bellcrank to the airplane structure with the bolt, grip bushing, washers and nut. Install up to a maximum of three washers (P/N 100951-X-032-XH) to provide 0.09-inch clearance between the straightener guide and the straightener bracket when the nose wheel is rotated as far as possible to the left and to the right. Lubricate the bellcrank with grease (61, Table 1, Chapter 91-00-00). (c) Attach the nose gear follow-up potentiometer at the steering bellcrank. (2) Straightener Link (a) Install the bolt, washer and nut attaching the aft end of the straightener link to the steering bellcrank. (b) Install the bolt, washer and nut attaching the aft end of the straightener link to the bracket on the left wheel well keel. Following installation, check and adjust nose wheel centering. The nose gear position switch, if installed, and the nose gear follow-up potentiometer.

C. Adjustment (1) Place the airplane on jacks (Ref. Chapter 07-10-00). (2) Check the straightener link for a nominal length of 29.95 inches. Adjust the link length by rotating each rod end in or out as required to obtain the correct length (Ref. Figure 207).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (3) Slowly retract the landing gear and observe the straightener roller on the nose gear as it makes contact with the straightener bracket on the airplane structure. During the retraction of the nose gear, the straightener roller should travel into the center slot of the straightener bracket without contacting the ramp portion of the bracket. Adjust the length of the straightener link as required to center the roller in the straightener bracket during retraction of the nose gear. (4) After the straightener link is properly adjusted, adjust the nose gear position switch and the nose gear follow-up potentiometer. (5) Remove the airplane from jacks (Ref. Chapter 07-10-00).

D. Nose Gear Position Switch Adjustment With the nose gear straightener link properly adjusted, adjust the nose gear position switch mounted on the straightener link. Loosen the lock nut and screw the actuating nut in or out as required. When the straightener link is in the static position, the switch is actuated using 0.06 ± 0.03 inch of the available overtravel. After adjustment of the switch, a minimum of two threads must be exposed on both ends of the actuating rod. If necessary, loosen the switch mounting nuts and reposition the switch to meet the above requirement (Ref. Figure 207).

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Figure 206 Nose Wheel Centering Mechanism Removal and Installation (UA-1 and After; UB-1 and After)

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Figure 207 Nose Wheel Centering and Position Switch Adjustment (UA-1 and After; UB-1 and After) Page 218 May 1/10

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LANDING GEAR POWER STEERING (UC-1 AND AFTER) DESCRIPTION AND OPERATION

32-52-00 00

1. GENERAL WARNING: When an airplane has experienced abnormal landing gear procedures of any type, as a safety precaution, place the airplane on jacks prior to performing any inspection or maintenance. Ensure that all three landing gears are down and locked prior to removing the airplane from jacks. When performing maintenance on a hydraulically operated landing gear system, be aware that any movement of a hydraulic actuator cylinder can displace hydraulic fluid and cause unanticipated movement of other actuator cylinders in the landing gear retraction system. Servicing of the landing gear hydraulic accumulator can also result in unanticipated movement of an actuator. Either action can result in an unsafe, unlocked landing gear system. Therefore, place the airplane on jacks prior to performing any inspection or maintenance. Cycle the landing gear and ensure that all three landing gears are down and locked prior to removing the airplane from jacks. Any time maintenance is to be performed on the landing gear system, place the airplane on jacks. When jacking the airplane in an unsheltered area where winds in excess of 35 kts will be encountered, never jack more than one gear at a time clear of the ground. Any time the landing gear is only partially retracted during maintenance, always cycle the gear with the power pack through at least one complete cycle before removing the airplane from the jacks. For safety reasons, pull the 2-ampere control circuit breaker on the pilot's inboard subpanel and place a note on the circuit breaker panel that LANDING GEAR MAINTENANCE IS IN PROCESS during maintenance on the landing gear. The landing gear control handle must never be moved from the down-and-locked position while the airplane is on the ground. It is recommended that the area be roped off during extension or retraction of the landing gear. CAUTION: Jacking of an airplane for the purpose of landing gear operation, inspection, servicing or maintenance, should be accomplished within an enclosed building or hangar. In the interest of safety, should it become necessary to jack the airplane in the open, wind velocity in any direction and terrain variations, must be compensated for prior to jacking the airplane. The landing gear must not be cycled with the power pack if low on fluid or if the landing gear system is not properly rigged. Use the emergency extension hand pump, TK229/939 hydraulic hand pump or TK229-1/939 air-driven hydraulic pump to extend and retract the landing gear for maintenance and rigging (Ref. Chapter 91-00-00). The power steering system has three modes: TAXI, PARK and caster. The TAXI mode provides ± 15° steering and PARK mode provides ± 55° steering. Both are selected by the TAXI/PARK switch on the pedestal. The caster mode is in effect when the system is OFF. Power lever disconnect switches and an activate switch control when the two steering modes are operable.

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Three annunciators report system status: •

PWR STEER ENGA - system is activated and ready.

•

PWR STEER FAIL - system electrical failure or low hydraulic pressure.

•

MAN STEER FAIL - system actuator has not returned to caster mode.

A 3000 psig hydraulic pump and motor assembly located in the nose wheel well supplies hydraulic pressure and receives hydraulic fluid from the hydraulic reservoir mounted on the RH pressure bulkhead. Power for the pump motor is supplied through the power steering motor relay and 30 ampere circuit breaker, which is located in the LH nacelle on the electrical power distribution panel. A pressure switch and filter are located in the nose wheel well. The pressure switch detects low pressure and signals the power steering relay PCB. The filter protects the hydraulic fluid from contamination. Two command potentiometers, mounted on the pilot's rudder pedal assembly, transmit the pilot's steering input to the power steering amplifier, located under the copilots seat. The amplifier and power steering relay PCB control the actuator. The amplifier compares potentiometer inputs and directs actuator movements. The power steering relay PCB contains relay logic to activate the system with respect to power lever positions, landing gear positions and signals the annunciator fault detect PCB. The power steering actuator twists the nose gear strut right or left with respect to rudder pedal movement. The actuator consists of two pistons, a selector valve, servo valve, arming valve, response restrictors, and a feedback potentiometer. The selector valve allows hydraulic fluid to move the pistons or isolates the pistons. The servo valve directs hydraulic fluid pressure to each piston as required for nose wheel movement. The actuator arming valve prevents the entry of the pressure fluid into the actuator until it is energized. The response restrictors dampen the actuator movement. The feedback potentiometer provides voltage feedback to the power steering amplifier which is recognized as nose wheel strut movement (Ref. Figure 103, POWER STEERING - TROUBLESHOOTING).

A. Power Steering Amplifier (UC-1 and After) The power steering amplifier contains fault diagnostic circuits to aid in isolating system faults. The fault diagnostic codes are shown in the POWER STEERING - TROUBLESHOOTING section of this chapter (Ref. Figure 102). Command, feedback and logic circuits control the actuator and report status to the relay PCB. Inputs from the command and feedback circuits are compared and as a result the actuator valves are positioned to move the nose gear right or left. When an imbalance of approximately 10° TAXI mode is sensed between the feedback potentiometer circuit and the command potentiometer circuit, the amplifier will interrupt the relay PCB logic, and the system will be deactivated. An adjustable potentiometer, placarded center on the power steering amplifier, is used for centering the nose gear.

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Figure 1 Power Steering System (UC-1 and After)

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LANDING GEAR POWER STEERING (UC-1 AND AFTER) TROUBLESHOOTING

100100

1. PROCEDURES WARNING: When an airplane has experienced abnormal landing gear procedures of any type, as a safety precaution, place the airplane on jacks prior to performing any inspection or maintenance. Ensure that all three landing gears are down and locked prior to removing the airplane from jacks. When performing maintenance on a hydraulically operated landing gear system, be aware that any movement of a hydraulic actuator cylinder can displace hydraulic fluid and cause unanticipated movement of other actuator cylinders in the landing gear retraction system. Servicing of the landing gear hydraulic accumulator can also result in unanticipated movement of an actuator. Either action can result in an unsafe, unlocked landing gear system. Therefore, place the airplane on jacks prior to performing any inspection or maintenance. Cycle the landing gear and ensure that all three landing gears are down and locked prior to removing the airplane from jacks. Any time maintenance is to be performed on the landing gear system, place the airplane on jacks. When jacking the airplane in an unsheltered area where winds in excess of 35 kts will be encountered, never jack more than one gear at a time clear of the ground. Any time the landing gear is only partially retracted during maintenance, always cycle the gear with the power pack through at least one complete cycle before removing the airplane from the jacks. For safety reasons, pull the 2-ampere control circuit breaker on the pilot's inboard subpanel and place a note on the circuit breaker panel that LANDING GEAR MAINTENANCE IS IN PROCESS during maintenance on the landing gear. The landing gear control handle must never be moved from the down-and-locked position while the airplane is on the ground. It is recommended that the area be roped off during extension or retraction of the landing gear. CAUTION: Jacking of an airplane for the purpose of landing gear operation, inspection, servicing or maintenance, should be accomplished within an enclosed building or hangar. In the interest of safety, should it become necessary to jack the airplane in the open, wind velocity in any direction and terrain variations, must be compensated for prior to jacking the airplane. The landing gear must not be cycled with the power pack if low on fluid or if the landing gear system is not properly rigged. Use the emergency extension hand pump, TK229/939 hydraulic hand pump or TK229-1/939 air-driven hydraulic pump to extend and retract the landing gear for maintenance and rigging (Ref. Chapter 91-00-00 Figure 1, Sheet 8).

A. Troubleshooting Notes (UC-1 and After) Refer to Chart 1 in Figure 103, TROUBLESHOOTING FLOWCHART. (1) Circuit breakers are to be CLOSED when checking voltages. (2) Circuit breakers are to be OPEN and tagged when checking continuity.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (3) Once a malfunction is isolated to an individual printed circuit board, printed wiring board, amplifier, or relay, replace that component, then continue. (4) Check voltage or continuity across switches by opening and closing switches as required. (5) If power steering electrical system is down, check relay and limiter in LH Nacelle. (6) Check solder joints for cold or open connections. (7) Check connectors and plugs for proper grounding and bent or broken pins. Check ground and shielded wires for proper ground.

B. Airplane Configuration (UC-1 and After) Prior to troubleshooting the power steering system, place the airplane in the following configuration. (1) Landing gear - DOWN and LOCKED. (2) Perform NOSE JACKING procedure (Ref. Chapter 07-10-00). (3) Nose wheel - CENTERED. (4) Rudder pedals in NEUTRAL with rig pin installed. (5) 28-vdc ground power is applied. (6) Engine power levers - less than 90%. (7) Main landing gear on the ground or squat switches bypassed (Ref. Model 1900C Airliner Electrical Wiring Diagram Manual (UC-1 and After) P/N 114-590021-61).

C. System Configuration (UC-1 and After) Configure the circuit breakers and switches as follows: (1) CB182 NOSE GEAR STEER circuit breaker - CLOSED. (2) CB138 ANN IND circuit breaker - CLOSED. (3) CB10 Power steering pump motor circuit breaker (Left Nacelle) - CLOSED. (4) Battery switch - ON. (5) GEN TIES switch - MAN CLOSE. (6) Power steering POWER switch - ON.

D. System Checks (UC-1 and After) To assist in determining the source of a system fault the power steering amplifier is equipped with fault diagnostics. In the event there is a system fault the amplifier will display a fault code. On earlier amplifiers, P/N 2-7657-4/4A, the 5 LED’s on the face represent a digital code; on UC-46 and After, and on spares replacements, the amplifiers, P/N 2-7657-5, have a single digit alphanumeric display on the top to represent the same diagnostic code (Ref. Figure 101).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL During the first two seconds after power-up the amplifier conducts a self-test. On the -4 amplifiers the green LED will illuminate followed by LED’s 1 through 4 at 0.5 second intervals. On -5 amplifiers the alpha-numeric display will illuminate. NOTE: On -5 amplifiers, disregard a 7 or 8 readout at power-up. (1) Turn the power steering POWER switch ON. (2) Check the amplifier for a fault code (Ref. Figure 101). (3) Press the power steering enable switch on the pedestal power lever and observe the following: (a) PWR STEER FAIL annunciator extinguishes. (b) PWR STEER ENGA annunciator illuminates. (c) Power steering pump begins running. (4) Remove the rig pin from the rudder pedal bellcrank and move the pedals full-left and full-right. (5) Center the rudder pedals and observe the position of the nose wheel. If the nose wheel is not centered, perform POWER STEERING OFFSET ADJUSTMENT. (6) Toggle the power steering mode switch in and out of PARK and TAXI and note any movement of the nose wheel. Refer to Chart 1 in Figure 103, TROUBLE SHOOTING FLOWCHART for corrective action. (7) After performing these Steps use the diagnostic codes and the TROUBLESHOOTING FLOWCHART to isolate any problems incurred. (8) After all checks and/or repairs are complete perform LOWERING THE AIRPLANE AFTER NOSE JACKING procedure (Ref. Chapter 07-10-00).

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E. Required Test Equipment (UC-1 and After) On the 1900 Airliner airplane, two test boxes are required to effectively troubleshoot the power steering system. The test boxes are installed in series with the power steering amplifier and airplane wire harness or between the power steering relay PCB and the airplane wire harness. Numbered test jacks on each test box accommodate checking voltages, measuring continuity and jumping system components. In addition, a nose centering tool may be used to closely measure nose wheel movement. Listed are the test boxes on the power steering system and nose centering tools to be used. Power Steering System Test Boxes (UC-1 and After) Part Number

Used On

114-380045/935

Power Steering Amplifier

114-380045/935-1

Power Steering Relay PCB

Nose Centering Tools (UC-1 and After) Part Number

Used On

114-820000/939

Nose Strut

114-820000/939-1

Nose Strut

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UE21B 991315AA.AI

Figure 101 Power Steering Amplifier Fault Display Codes (UC-1 and After)

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UE32B 991238AA.AI

Figure 102 Power Steering Schematic (UC-1 and After)

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Figure 103 (Sheet 1 of 7) Troubleshooting Flowchart (UC-1 and After)

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Figure 103 (Sheet 2 of 7) Troubleshooting Flowchart (UC-1 and After)

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Figure 103 (Sheet 3 of 7) Troubleshooting Flowchart (UC-1 and After)

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Figure 103 (Sheet 4 of 7) Troubleshooting Flowchart (UC-1 and After)

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Figure 103 (Sheet 5 of 7) Troubleshooting Flowchart (UC-1 and After)

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Figure 103 (Sheet 6 of 7) Troubleshooting Flowchart (UC-1 and After)

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Figure 103 (Sheet 7 of 7) Troubleshooting Flowchart (UC-1 and After)

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Figure 104 Power Steering Relay Panel (UC-1 and After)

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Figure 105 Power Steering Actuator Schematic (UC-1 and After)

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Figure 106 Troubleshooting - Power Steering Valves (7) (UA-1 and After; UB-1 and After)

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LANDING GEAR POWER STEERING (UC-1 AND AFTER) MAINTENANCE PRACTICES

200200

1. PROCEDURES WARNING: When an airplane has experienced abnormal landing gear procedures of any type, as a safety precaution, place the airplane on jacks prior to performing any inspection or maintenance. Ensure that all three landing gears are down and locked prior to removing the airplane from jacks. When performing maintenance on a hydraulically operated landing gear system, be aware that any movement of a hydraulic actuator cylinder can displace hydraulic fluid and cause unanticipated movement of other actuator cylinders in the landing gear retraction system. Servicing of the landing gear hydraulic accumulator can also result in unanticipated movement of an actuator. Either action can result in an unsafe, unlocked landing gear system. Therefore, place the airplane on jacks prior to performing any inspection or maintenance. Cycle the landing gear and ensure that all three landing gears are down and locked prior to removing the airplane from jacks. Any time maintenance is to be performed on the landing gear system, place the airplane on jacks. When jacking the airplane in an unsheltered area where winds in excess of 35 kts will be encountered, never jack more than one gear at a time clear of the ground. Any time the landing gear is only partially retracted during maintenance, always cycle the gear with the power pack through at least one complete cycle before removing the airplane from the jacks. For safety reasons, pull the 2-ampere control circuit breaker on the pilot's inboard subpanel and place a note on the circuit breaker panel that LANDING GEAR MAINTENANCE IS IN PROCESS during maintenance on the landing gear. The landing gear control handle must never be moved from the down-and-locked position while the airplane is on the ground. It is recommended that the area be roped off during extension or retraction of the landing gear. CAUTION: Jacking of an airplane for the purpose of landing gear operation, inspection, servicing or maintenance, should be accomplished within an enclosed building or hangar. In the interest of safety, should it become necessary to jack the airplane in the open, wind velocity in any direction and terrain variations, must be compensated for prior to jacking the airplane. The landing gear must not be cycled with the power pack if low on fluid or if the landing gear system is not properly rigged. Use the emergency extension hand pump, TK229/939 hydraulic hand pump or TK229-1/939 air-driven hydraulic pump to extend and retract the landing gear for maintenance and rigging (Ref. Chapter 91-00-00).

2. COMMAND POTENTIOMETER A. Removal (UC-1 and After) (1) Disconnect ground power and turn the battery OFF.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (2) Remove floor panel (1) under the pilot's rudder pedals (Ref. Chapter 06-50-00). (3) Disconnect the electrical connector from the command potentiometer (Ref. Figure 1, in the Description and Operation section). (4) Disconnect the spring, if installed, from the adjustable link. (5) Disconnect adjustable link, at the forward clevis, from the lever on the command potentiometer. (6) Move the lever on the potentiometer as required to gain access to the retaining pin securing lever to the input shaft. (7) Remove pin and lever. (8) Loosen the three screws securing the mounting clamps of the potentiometer. Remove one clamp and screw. (9) Remove potentiometer.

B. Installation (UC-1 and After) (1) Set the command potentiometer in place on the mounting plate and install the clamp and screw (Ref. Figure 1, in the Description and Operation section). (2) Adjust the three screws of the mounting clamps to support the potentiometer and allow rotational movement. (3) Install the lever on the command potentiometer. Secure with pin. (4) Connect the forward end of the adjustable link to the lever. (5) Perform COMMAND POTENTIOMETER RIGGING and COMMAND POTENTIOMETER ADJUSTMENTS found in this section. (6) Connect the electrical connector of the command potentiometer. (7) Connect spring, if removed, to adjustable link. (8) Install pilots floor panel (1) (Ref. Chapter 06-50-00). (9) Return airplane to service.

C. Rigging (UC-1 and After) (1) Disconnect ground power and turn the battery OFF. (2) Remove floor panel (1) under the pilot's rudder pedals (Ref. Chapter 06-50-00). (3) Set the rudder pedals at neutral and install a rig pin in the rudder pedal bellcrank. (4) Disconnect the electrical connector (E182J1) from the command potentiometer (Ref. Figure 1, in the Description and Operation section). (5) Loosen the forward and aft clamps securing the command potentiometer assembly on the rudder flight control torque tubes.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (6) Install a 3/16-inch-diameter rig pin through the aft clamps and into the forward clamps to align the aft lever assembly with the potentiometer support assembly. (7) Loosen the three screws securing the mounting clamps of the command potentiometer. (8) Measure between pins E182J1-2 and E182J1-3; adjust the potentiometer to a reading of 2500 ohms. NOTE: DO NOT overtighten the screws on the potentiometer mounting clamps. The adjustments made may move if the clamps are too tight. (9) Tighten the screws securing the mounting clamps. (10) Tighten the forward and aft clamps holding the potentiometer assembly on the torque tubes of the rudder flight control system. (11) Remove the rig pin from the forward and aft clamps. (12) Remove the rig pin from the rudder pedal bellcrank. (13) Install pilots floor panel (1) (Ref. Chapter 06-50-00). (14) Return airplane to service.

D. Command Potentiometer and Amplifier Adjustments (UC-1 and After) NOTE: To perform these adjustments, the power steering system must be ON and enabled. Refer to AIRCRAFT CONFIGURATION and SYSTEM CONFIGURATION in the troubleshooting procedures section. (1) Remove floor panel (1) under the pilot's rudder pedals (Ref. Chapter 06-50-00). (2) Set the engine power levers to IDLE (Ref. Figure 1, in the Description and Operation section). (3) Connect the 114-380045/935 test box between the power steering amplifier and the airplane wire harness. (4) Install the nose centering tool, P/N 114-820000/939, on the nose strut. (5) Remove the spring, if installed, from the adjustable link. (6) Remove the safety clips from the adjustable link of the command potentiometer. (7) Remove the rig pin from the rudder pedal bellcrank. (8) Remove the stop pins from the nose centering tool. Do not allow the actuator to travel and hit the stops. (9) Measure and record the voltage between test jacks 22 and 14 on the test box with the rudder pedals in neutral. (10) Measure and record the voltage with the rudder pedals hard left. (11) Measure and record the voltage with the rudder pedals hard right.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (12) If the voltage measurements from neutral to hard left and neutral to hard right are within 0.015 volt, secure the adjustable link with the safety clips. (13) If the neutral to right voltage is greater than the left voltage reading, shorten the adjustable link. (14) If the neutral to right voltage is less than the left voltage reading, lengthen the adjustable link. (15) Install the safety clips in the adjustable links. (16) Install the spring, if removed, on the adjustable link. (17) Set the rudder pedals at neutral. (18) Toggle the PARK/TAXI switch back and forth between positions and observe any nose wheel movement. Adjust potentiometer until there is no movement of the nose wheel when the selector switch is switched between TAXI and PARK mode. NOTE: DO NOT overtighten the screws on the potentiometer mounting clamps. The adjustments made may move if the clamps are too tight. (19) Tighten the screws of the mounting clamps of the command potentiometer. (20) Check for nose wheel movement again after tightening the command potentiometer clamps. (21) Install a rig pin in the rudder pedal bellcrank. (22) Check the position of the nose wheel. The nose wheel may not be at true center position. Adjust the power steering amplifier centering potentiometer through the opening placarded “CENTER” adjustment location of the applicable amplifier. (23) Install pilots floor panel (1) (Ref. Chapter 06-50-00). (24) Return the airplane to service.

3. HYDRAULIC FILTER A. Servicing Replace the system filter (10) every 100 hours of service for the first 300 hours and every 600 hours there after. The filter is accessible through the nose wheel well (Ref. Figure 201).

4. SUPPLY AND RETURN FILTER UNION (UC-143 AND AFTER; AND EARLIER UC SERIALS WITH KIT NO. 114-8021-1S INSTALLED) A. Servicing Remove and clean or replace the filter unions 100 hours after any maintenance on the power steering system which required opening any hydraulic line, fitting or port in the actuator between the primary system filter and the actuator, or if power steering operation indicates filters have possibly become clogged (Ref. Figure 201). The filter unions are accessible in the nose wheel well, in the hydraulic supply and return ports in the actuator.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL CAUTION: Exercise extreme care to avoid introducing any contamination into the system when servicing these filters. (1) Carefully clean the exterior of the filter unions and nearby portions of the lines and the actuator body. (2) Disconnect the attaching lines and remove the filter unions. Tag the unions for correct port location to ensure proper installation. (3) Clean the filter unions with cleaning solvent (2, Table 1, Chapter 91-00-00) and blow dry with filtered compressed air. NOTE: It is recommended that the unions be agitated in an ultrasonic cleaner, then flushed with clean solvent. (4) Install the filter unions in the same ports from which they were removed. (5) Perform HYDRAULIC RESERVOIR SERVICING procedure in this section.

5. HYDRAULIC RESERVOIR (UC-1 AND AFTER) A. Servicing (1) Fill the hydraulic reservoir (1) (Ref. Figure 201) with hydraulic fluid (39, Table 1, Chapter 91-00-00). (2) Bleed the hydraulic system by loosening the B-nut at the hydraulic filter inlet and operating the pump and motor briefly to ensure that all air has been bled from the inlet and outlet lines to the pump. Tighten the fitting. (3) Fill the hydraulic reservoir (1) (Ref. Figure 201) with hydraulic fluid (39, Table 1, Chapter 91-00-00). (4) Operate the pump and motor until the air has been purged from the rest of the system. (5) Fill the hydraulic reservoir with hydraulic fluid.

6. ACTUATOR (UC-1 AND AFTER) A. Removal (1) Disconnect and tag the four electrical connectors (5), (6), (16), and (18) from the power steering actuator (7) (Ref. Figure 202). CAUTION: As hydraulic lines are disconnected, tag and plug or cap the openings. (2) Disconnect the hydraulic supply and return lines (4) and (11) from the actuator at the bracket (Ref. Figure 201). (3) Perform NOSE LANDING GEAR REMOVAL procedure (Ref. 32-20-00). (4) Loosen the nuts on the eccentric bolts (10 and 15). Turn the bolts slightly to move the torque stop adjustment blocks (9) away from the torque stops (11) (Ref. Figure 202). (5) Remove the six bolts (2) attaching the straightener arm (3) to the nose gear assembly.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (6) Remove the three bolts (4) and actuator (7) from the airplane.

B. Installation (1) Position the actuator on the mounting plate (8) and secure with the three bolts (4) (Ref. Figure 202). (2) Position the actuator and mounting plate on the nose gear. Position the slots in the plate over the torque stops (11). (3) Attach the actuator and straightener arm (3) to the nose gear with six bolts (2). Align the straightener roller and index mark (17) on the actuator with the nose gear wheel. (4) Loosen the nut on the eccentric bolt (15) and turn the bolt until the torque stop adjustment block (9) holds the mounting plate (8) solid against the torque stop (11). Tighten nut. (5) Adjust one block (9) to where it has minimal contact up to 0.001 inch clearance from the stop (11), then adjust the other block to have minimal contact up to 0.001 inch clearance. NOTE: Use care when tightening the eccentric bolts to avoid altering adjustments. (6) Tighten the nuts on eccentric bolts (10) and (15). Torque 225 to 250 inch-pounds. (7) Check that the shock strut rotates in the nose gear brace without binding. (8) Perform NOSE LANDING GEAR INSTALLATION procedure (Ref. 32-20-00). NOTE: Any time a fitting is loosened or removed, install new packings. (9) Connect the hydraulic supply and return lines (4) and (11) to the actuator (Ref. Figure 201). (10) Perform HYDRAULIC RESERVOIR SERVICING procedure in this section.

7. SERVO VALVE (UC-1 AND AFTER) A. Removal CAUTION: Before removing any component from the actuator, completely clean the actuator around the component with a clean, lint-free rag lightly moistened with MPK. Do not clean the actuator with high pressure water or steam since this may damage the electrical wiring and connectors. Ensure that the highest standards of cleanliness are maintained whenever the actuator or any line leading to the actuator is opened. (1) Disconnect the return fluid hose from the tee on the left side of the nose wheel well. Dispose of any fluid. Cap both the tee and hose (Ref. Figure 201). NOTE: Retain and note locations of all attaching hardware for later use. (2) Disconnect the electrical connector and remove the servo valve harness from all clamps and supports (Ref. Figure 203). (3) Remove safety wire on the four internally wrenching servo valve retaining screws. Remove the screws and the servo valve.

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B. Installation CAUTION: The servo valve may be damaged by rough handling. (1) Ensure that the four packings supplied with servo valve are properly seated in the valve ports and install the valve using the four internally wrenching screws. Install safety wire through the four screw heads (Ref. Figure 203). (2) Connect the servo valve electrical connector and install the harness into the clamps and supports. (3) Connect the return fluid hose to the tee on the left side of the nose wheel well. (4) Perform HYDRAULIC RESERVOIR SERVICING procedure in this section and check for leaks. (5) Any time a servo valve is removed the following tests must be performed: (a) Perform NOSE JACKING procedure (Ref. Chapter 07-10-00). (b) Connect external power. (c) Turn on and engage power steering. (d) With the system operating in park mode, check that the nose strut can travel smoothly to 55 ± 5° left and right from neutral or centered position. If travel is jerky or the rate varies appreciably through the range of travel, check all joints for binding and proper lubrication before proceeding. (e) Starting from a nose strut position of 45° right, abruptly swing the rudder pedals to full nose left position and record the time required for the nose strut to travel through 45° left position to the nearest ± 0.2 second. Repeat this procedure from left to right. The times shall be between 5.0 and 6.0 seconds at a fluid temperature of 80° to 90°F. Times outside this range indicate the servo valve may have been contaminated or damaged during installation. NOTE: Times below this range due to prolonged system operation or high fluid temperatures and times above this range due to low fluid temperatures are normal and acceptable as long as the system operation is smooth and the rate of nose wheel travel to the left is approximately the same as to the right. (f) Turn power steering off, turn master power off. (g) Disconnect ground power from the airplane. (h) Perform LOWERING THE (Ref. Chapter 07-10-00).

AIRPLANE

AFTER

NOSE

JACKING

procedure

8. ARMING VALVE (UC-1 AND AFTER) A. Removal CAUTION: Before removing any component from the actuator, completely clean the actuator around the component with a clean, lint-free rag lightly moistened with MPK. Do not clean the actuator with high pressure water or steam since this may damage the electrical wiring and connectors. Ensure that the highest standards of cleanliness are maintained whenever the actuator or any line leading to the actuator is opened.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (1) Disconnect the pressure and return fluid tubes from the actuator fittings. Dispose of any drained fluid. Cap both the tubes and actuator fittings (Ref. Figure 202). NOTE: Retain and note locations of all attaching hardware for later use. (2) Disconnect the servo valve, arming valve and mode indicating switch electrical connectors and remove the harnesses from all clamps and supports. (3) Remove safety wire on all eight of the internally wrenching servo valve and arming valve retaining screws. Remove the four screws at the ends of the manifold and remove the complete manifold assembly from the actuator. Plug the ports in the manifold and actuator assemblies and remove the manifold assembly to a clean, protected work area. CAUTION: The pressure and return seats between the arming valve solenoid and the manifold assembly are spring loaded. (4) Carefully remove the two remaining internally wrenching solenoid retaining screws so as not to lose the valve seats or ball. Remove the solenoid from the manifold assembly and remove the valve seats, ball, and seals from the bottom of the solenoid.

B. Installation NOTE: To assist assembly of the arming valve onto the manifold assembly, application of 28 VDC across pins A and B of the solenoid will hold the ball and valve seats into the solenoid cavity. (1) Using new seals, assemble the ball, valve seats and seals into the solenoid cavity and install the arming valve assembly onto the manifold assembly with the solenoid wire harness toward the servo valve. Install the two short internally wrenching solenoid retaining screws adjacent to the servo valve (Ref. Figure 202). (2) Remove the port plugs from the manifold and actuator assemblies, assemble two new packings into the manifold ports and install the complete manifold assembly onto the actuator assembly. Install safety wire through the eight screws. (3) Connect the servo valve, arming valve and mode-indicating switch electrical connectors, and install the harnesses into the clamps and supports. (4) Connect the pressure line to the actuator inlet fitting and uncap the outlet fitting. (5) Connect external power, fill the brake and power steering reservoir and bleed air from the plumbing and actuator by momentarily engaging the power steering system. (6) Whenever an arming valve solenoid or any of the arming valve seals or seats are replaced, the following test must be performed: (a) Unplug the arming valve electrical connector, turn on and engage the power steering system. Hold the engage button on the power lever if required to keep the system engaged. (b) Allow the leakage from the actuator port to stabilize and collect the leakage for one minute. The leakage shall not exceed 20 drops per minute. Leakage in excess of this amount indicates the seals or pressure seats may have been damaged during assembly and should be replaced. (c) Unplug and connect the actuator fluid return tube, fill and bleed the system.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (d) Perform NOSE JACKING procedure (Ref. Chapter 07-10-00). (e) Engage power steering and check for proper operation. (f) Perform LOWERING THE (Ref. Chapter 07-10-00).

AIRPLANE

AFTER

NOSE

JACKING

procedure

9. MANUAL STEERING FAIL SWITCH (UC-1 AND AFTER) A. Removal (1) Disconnect electrical connector of the fail switch and remove the switch harness from all clamps and supports (Ref. Figure 202). (2) Remove safety wire from the two internally wrenching screws through the switch mounting flange. Remove the screws and the switch.

B. Installation (1) Connect the electrical connector of the fail switch and install the switch harness into the clamps and supports (Ref. Figure 202). (2) Anytime a fail switch is replaced the following test must be performed: (a) With the fail switch connected electrically but not attached to the actuator, turn on electrical power and observe that the MAN STEER FAIL annunciator illuminates. (b) Press the switch plunger and observe that the MAN STEER FAIL annunciator extinguishes. Failing this tests indicates there is fault in the wiring or annunciation system. (3) Install the switch with the two internally wrenching screws. Install safety wire through the screws.

10. FEEDBACK POTENTIOMETER (UC-1 AND AFTER) A. Removal (1) Perform NOSE LANDING GEAR REMOVAL procedure (Ref. 32-20-00). (2) Remove safety wire from the two internally wrenching feedback potentiometer retaining screws. Remove the screws and the feedback potentiometer (Ref. Figure 203).

B. Installation (1) Rotate the feedback gear on the potentiometer assembly until the resistance between pins A and B of the connector is 2,000 ± 250 ohms. Without altering this adjustment install the feedback potentiometer onto the actuator assembly with the nose gear in neutral or center position and with the feedback potentiometer wire harness routed toward the manifold assembly (Ref. Figure 202). (2) Install the internally wrenching screws loosely so that the feedback potentiometer can be rotated. (3) Position the nose gear assembly near the airplane and connect the feedback potentiometer electrical connector to the ships harness. Disconnect the electrical connector from the power steering pump motor.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (4) Install breakout box (P/N 114-380045/935) at the power steering amplifier connector. (5) Connect external power and turn the power steering switch on. (6) With the neutral indicating marks on top of the actuator aligned, rotate the feedback potentiometer in its mount until the voltage between test points 17 and 33 is 2.5 ± 0.5 VDC. (7) Tighten the internally wrenching feedback potentiometer retaining screws without altering the feedback potentiometer adjustment. Install the safety wire. (8) Turn power steering switch off, disconnect external power, remove the breakout box and disconnect the feedback potentiometer electrical connector. (9) Perform NOSE LANDING GEAR INSTALLATION procedure (Ref. 32-20-00). (10) Perform COMMAND POTENTIOMETER AND AMPLIFIER ADJUSTMENTS in this section.

11. AMPLIFIER AR103 (UC-1 AND AFTER) A. Removal (1) Verify that power to amplifier AR103 is OFF (Ref. Figure 1, in the Description and Operation section). (2) Remove copilot's seat (Ref. Chapter 25-10-00). (3) Remove floorboard (5) beneath copilot's seat (Ref. Chapter 06-50-00). (4) Disconnect electrical connector J1 from amplifier. (5) Remove attaching screws and amplifier AR103 from airplane.

B. Installation (1) Position amplifier AR103 on panel and secure with attaching screws (Ref. Figure 1, in the Description and Operation section). (2) Connect electrical connector J1. (3) Install the floorboard (5) under the copilot's seat (Ref. Chapter 06-50-00). (4) Install the copilot's seat (Ref. Chapter 25-10-00). (5) Perform COMMAND POTENTIOMETER AND AMPLIFIER ADJUSTMENTS procedure in this section and SYSTEM CHECKS procedure found in the troubleshooting procedures section to align the nose wheel and verify proper operation.

12. RELAY PANEL A320 (UC-1 AND AFTER) A. Removal (1) Verify electrical power to A320 panel is OFF (Ref. Figure 1, in the Description and Operation section). (2) Remove copilot's seat (Ref. Chapter 25-10-00).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (3) Remove floorboard (5) beneath copilot's seat (Ref. Chapter 06-50-00). (4) Disconnect electrical connector from A320 panel. (5) Remove attaching screws and A320 panel from airplane.

B. Installation (1) Position A320 relay panel and secure with attaching screws (Ref. Figure 1, in the Description and Operation section). (2) Connect electrical connector to A320 panel. (3) Install the floorboard (5) under the copilot's seat (Ref. Chapter 06-50-00). (4) Install the copilot's seat (Ref. Chapter 25-10-00). (5) Apply power and verify proper operation.

13. HYDRAULIC PRESSURE SWITCH (UC-1 AND AFTER) A. Removal (1) Verify electrical power to pressure switch (8) is OFF (Ref. Figure 201). (2) Disconnect electrical connector from pressure switch (8). (3) Remove the pressure switch (8) from the elbow (14).

B. Installation (1) Install the pressure switch (8) in the fitting (14) (Ref. Figure 201). (2) Connect electrical connector to pressure switch (8). (3) Apply electrical power and verify proper operation. (4) Perform HYDRAULIC RESERVOIR SERVICING procedure in this section.

14. PUMP RELIEF VALVE (UC-1 AND AFTER) A. Removal (1) Remove the clamp screw on the relief valve (9) (Ref. Figure 201). CAUTION: As hydraulic lines are disconnected, tag and plug or cap the openings. (2) Disconnect the hydraulic line from the relief valve. (3) Unscrew relief valve from elbow and remove from airplane.

B. Installation (1) Screw the relief valve (9) into the elbow (14) (Ref. Figure 201).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL NOTE: Any time a fitting is loosened or removed, install new packings. (2) Connect the hydraulic line to the relief valve. (3) Install the screw through the clamp securing the valve. (4) Perform HYDRAULIC RESERVOIR SERVICING procedure in this section.

15. PUMP AND MOTOR (UC-1 AND AFTER) A. Removal (1) Disconnect electrical connector (5) from power steering pump motor (Ref. Figure 201). CAUTION: As hydraulic lines are disconnected, tag and plug or cap the openings. (2) Disconnect hydraulic lines from hydraulic pump. (3) Remove attaching bolts and hydraulic pump and motor (3 and 6) from airplane. NOTE: If pump and motor need to be disassembled for further repair, refer to the supplier components manual.

B. Installation (1) Position the pump and motor (3) and (6), and secure to the bracket with attaching bolts (Ref. Figure 201). NOTE: Any time a fitting is loosened or removed, install new packings. (2) Connect hydraulic lines to hydraulic pump. (3) Connect electrical connector (5) to pump motor. (4) Perform HYDRAULIC RESERVOIR SERVICING procedure in this section.

16. HYDRAULIC RESERVOIR (UC-1 AND AFTER) A. Removal (1) Open hydraulic reservoir access door (Ref. Figure 201). (2) Siphon hydraulic fluid from reservoir with vacuum type bulb into suitable container. (3) Remove the attaching hardware and shield from the reservoir. CAUTION: As hydraulic lines are disconnected, tag and plug or cap the openings. (4) Disconnect lines from hydraulic reservoir. (5) Remove attaching bolts and hydraulic reservoir (1) from airplane.

B. Installation (1) Position hydraulic reservoir (1) and secure with attaching bolts (Ref. Figure 201).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (2) Position the shield on the reservoir and secure with the attaching hardware. NOTE: Anytime a fitting is loosened or removed, install new packings. (3) Connect hydraulic lines to reservoir. (4) Perform HYDRAULIC RESERVOIR SERVICING procedure in this section.

17. NOSE GEAR CENTERING MECHANISM (UC-1 AND AFTER) A. Removal (1) Straightener Link (Ref. Figures 204 and 205) (a) Remove the nut, washer and bolt attaching the aft end of the link to the bracket on the left wheel well keel. (b)

Remove the nut, washer and bolt attaching the forward end of the link to the steering bellcrank.

(2) Steering Bellcrank (a) Remove the nut, washers, grip bushing and bolt attaching the steering bellcrank to the airplane structure. (b) The straightener guide can be removed from the bellcrank with removal of the two attaching bolts.

B. Installation (1) Steering Bellcrank (Ref. Figures 204 and 205) (a) Secure the straightener guide to the bellcrank with the two bolts, nuts, and washers. (b) Attach the bellcrank to the airplane structure with the bolt, grip bushing, washers and nut. Install up to a maximum of three washers (P/N 1-0095-X-032-XH) to provide 0.09-inch clearance between the straightener guide and the straightener bracket when the nose wheel is rotated as far as possible to the left and to the right. Lubricate the bellcrank. (2) Straightener Link (a) Install the bolt, washer and nut attaching the forward end of the straightener link to the steering bellcrank. (b) Install the bolt, washer and nut attaching the aft end of the straightener link to the bracket on the left wheel well keel. (c) Following installation, check and adjust the nose wheel centering. Refer to NOSE WHEEL CENTERING ADJUSTMENT in this section.

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18. NOSE WHEEL CENTERING (UC-1 AND AFTER) A. Adjustment (1) Perform THREE-POINT JACKING (PREFERRED PROCEDURE) procedure (Ref. Chapter 07-10-00). All tires must be clear of the floor. Strut limiters will reduce strut extension (Ref. Chapter 91-00-00, Figure 1, Sheet 1). (2) Check the straightener link for a nominal length of 29.95 inches (Ref. Figure 205). Adjust the link length by rotating each rod end in or out as required to obtain the correct length. (3) Slowly retract the landing gear and observe the straightener roller on the nose gear as it makes contact with the straightener bracket on the airplane structure. During the retraction of the nose gear, the straightener roller should travel into the center slot of the straightener bracket without contacting the ramp portion of the bracket. Adjust the length of the straightener link as required to center the roller in the straightener bracket during retraction of the nose gear. (4) After all adjustments are complete perform LOWERING THE AIRPLANE AFTER THREE-POINT JACKING procedure (Ref. Chapter 07-10-00).

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Figure 201 Power Steering Hydraulic Plumbing (UC-1 and After)

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Figure 202 Power Steering Hydraulic Plumbing (UC-1 and After)

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Figure 203 Power Steering Actuator Components (UC-1 and After)

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Figure 204 Nose Wheel Centering Mechanism

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Figure 205 Nose Wheel Centering and Position Switch

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LANDING GEAR LANDING GEAR POSITION AND WARNING DESCRIPTION AND OPERATION

32-60-00 00

1. GENERAL WARNING: When an airplane has experienced abnormal landing gear procedures of any type, as a safety precaution, place the airplane on jacks prior to performing any inspection or maintenance. Ensure that all three landing gears are down and locked prior to removing the airplane from jacks. When performing maintenance on a hydraulically operated landing gear system, be aware that any movement of a hydraulic actuator cylinder can displace hydraulic fluid and cause unanticipated movement of other actuator cylinders in the landing gear retraction system. Servicing of the landing gear hydraulic accumulator can also result in unanticipated movement of an actuator. Either action can result in an unsafe, unlocked landing gear system. Therefore, place the airplane on jacks prior to performing any inspection or maintenance. Cycle the landing gear and ensure that all three landing gears are down and locked prior to removing the airplane from jacks. Any time maintenance is to be performed on the landing gear system, place the airplane on jacks. When jacking the airplane in an unsheltered area where winds in excess of 35 kts will be encountered, never jack more than one gear at a time clear of the ground. Any time the landing gear is only partially retracted during maintenance, always cycle the gear with the power pack through at least one complete cycle before removing the airplane from the jacks. For safety reasons, pull the 2-ampere control circuit breaker on the pilot's inboard subpanel and place a note on the circuit breaker panel that LANDING GEAR MAINTENANCE IS IN PROCESS during maintenance on the landing gear. The landing gear control handle must never be moved from the down-and-locked position while the airplane is on the ground. It is recommended that the area be roped off during extension or retraction of the landing gear. CAUTION: Jacking of an airplane for the purpose of landing gear operation, inspection, servicing or maintenance, should be accomplished within an enclosed building or hangar. In the interest of safety, should it become necessary to jack the airplane in the open, wind velocity in any direction and terrain variations, must be compensated for prior to jacking the airplane. The landing gear must not be cycled with the power pack if low on fluid or if the landing gear system is not properly rigged. Use the emergency extension hand pump, TK229/939 hydraulic hand pump or TK229-1/939 air-driven hydraulic pump to extend and retract the landing gear for maintenance and rigging (Ref. Chapter 91-00-00). Visual indication of the landing gear position is provided by two red in-transit lights located in the landing gear control switch handle and a green GEAR DOWN indicator light assembly, labeled NOSE - L - R, located adjacent to the landing gear control switch on the pilot's inboard subpanel (Ref. Figure 1). Illumination of the red in-transit lights indicates when the landing gear is in transit; gear-up is indicated when they go out.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Illumination of the green GEAR DOWN light indicates when the landing gear is down and locked. The red control handle lights can be checked by pressing the HDL LT TEST push-button switch located to the right of the landing gear control switch. To check the GEAR DOWN lights press the light assembly case. Power for the in-transit and GEAR DOWN lights is drawn from a 5 ampere LANDING GEAR - IND circuit breaker located on the RH circuit breaker panel. A circuit to the in-transit lights is completed through the downlock and up-position switches of each landing gear and through the automatic light dimming circuit. A circuit to the GEAR DOWN lights is completed through the downlock and actuator downlock switches of each landing gear. An up-position switch is located in the upper part of each wheel well. When the landing gear is in the fully retracted position, each landing gear strut actuates its respective up-position switch to open the circuit to the in-transit lights. The in-transit lights will remain illuminated until all three landing gears are fully retracted. As soon as the landing gear moves from the fully retracted position, each landing gear strut actuates its respective up-position switch to illuminate the in-transit lights. The in-transit lights go out when the drag brace of each landing gear actuates its respective downlock switch to open the circuit to the in-transit lights and to complete the circuit to the actuator downlock switch. In this position the internal locking mechanism of the landing gear actuator will actuate the actuator downlock switch mounted in the actuator to illuminate its respective GEAR DOWN light. A landing gear warning system is provided to warn the pilot that the landing gear is not down and locked during specific flight regimes. Various warning modes result, depending upon the position of the flaps. With the flaps in the UP, TAKE OFF or APPROACH position and either or both power levers retarded below 84% to 86% N1, the warning horn will sound intermittently and the landing gear control switch handle lights will illuminate. The horn can be silenced by pressing the WARN HORN SILENCE button adjacent to the landing gear switch handle; the lights in the landing gear control switch handle cannot be cancelled. The landing gear warning system will be rearmed if the power lever(s) are advanced sufficiently. With the flaps beyond the APPROACH position, the warning horn and landing gear control switch handle lights will be activated regardless of the power settings, and neither can be cancelled. To prevent accidental gear retraction on the ground, the safety switch on the right main strut breaks the control circuit whenever the strut is compressed. CAUTION: Never rely on the safety switch to keep the gear down while taxiing or on landing or take-off roll. Always check the position of the landing gear control switch handle. The safety switch also actuates a solenoid-operated downlock hook, which prevents the landing gear handle from being raised when the airplane is on the ground. The hook automatically unlocks when the airplane leaves the ground, but it can be manually overridden by pressing down on the red button placarded DN LCK REL. The right safety switch will ground the in-transit light relay to illuminate the in-transit lights anytime the landing gear control handle is moved to the UP position while the airplane is on the ground with the shock strut compressed.

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Figure 1 Landing Gear Position and Warning System

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LANDING GEAR LANDING GEAR POSITION AND WARNING TROUBLESHOOTING

100100

1. PROCEDURE WARNING: When an airplane has experienced abnormal landing gear procedures of any type, as a safety precaution, place the airplane on jacks prior to performing any inspection or maintenance. Ensure that all three landing gears are down and locked prior to removing the airplane from jacks. CAUTION: Jacking of an airplane for the purpose of landing gear operation, inspection, servicing or maintenance, should be accomplished within an enclosed building or hangar. In the interest of safety, should it become necessary to jack the airplane in the open, wind velocity in any direction and terrain variations, must be compensated for prior to jacking the airplane. Table 101 TROUBLESHOOTING - LANDING GEAR WARNING HORN AND SWITCHES PROBLEM 1. Warning horn inoperative.

2. Circuit breaker tripping.

PROBABLE CAUSE

CORRECTIVE ACTION

a. Horn circuit breaker tripped.

a. Reset circuit breaker.

b. Open or grounded circuit.

b. Check continuity: Circuit breaker-to-horn throttle switches, landing gear downlock switches and ground.

c. Throttle switches fail to operate properly or out of adjustment.

c. Adjust or replace.

d. Main or nose gear downlock switches fail to operate properly or out of adjustment.

d. Adjust or replace.

e. Flasher stuck open.

e. Check flasher continuity.

a. Grounded circuit.

a. Check for ground between circuit breaker and warning horn.

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Table 102 TROUBLESHOOTING - LANDING GEAR POSITION INDICATORS AND SWITCHES PROBLEM 1. All indicators inoperative.

2. One indicator completely inoperative.

Page 102 Nov 1/09

PROBABLE CAUSE

CORRECTIVE ACTION

a. Landing gear indicator circuit breaker tripped.

a. Reset.

b. Open circuit between circuit breaker and indicators.

b. Check continuity.

a. Indicator fails to operate properly.

a. Repair or replace.

b. Open circuit.

b. Check continuity.

c. Landing gear downlock, up-position or actuator downlock switches fail to operate properly or out of adjustment.

c. Adjust, repair or replace.

d. The nose gear down-position switch adjustable stop bracket is bowed or bent.

d. Straighten the bracket and shim or replace the bracket.

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LANDING GEAR LANDING GEAR POSITION AND WARNING MAINTENANCE PRACTICES

32-60-00 200200

1. PROCEDURES WARNING: When performing maintenance on a hydraulically operated landing gear system, be aware that any movement of a hydraulic actuator cylinder can displace hydraulic fluid and cause unanticipated movement of other actuator cylinders in the landing gear retraction system. Servicing of the landing gear hydraulic accumulator can also result in unanticipated movement of an actuator. Either action can result in an unsafe, unlocked landing gear system. Therefore, place the airplane on jacks prior to performing any inspection or maintenance. Cycle the landing gear and ensure that all three landing gears are down and locked prior to removing the aircraft from jacks.

2. NOSE GEAR UP-POSITION INDICATOR SWITCH A. Adjustment WARNING: Make all adjustments with the power off at the master switch and no external power connected. Any time maintenance is to be performed on the landing gear system, place the airplane on jacks. When jacking the airplane in an unsheltered area where winds in excess of 35 kts will be encountered, never jack more than one gear at a time clear of the ground. The landing gear control handle must never be moved from the down-and-locked position while the airplane is on the ground. It is recommended that the area be roped off during extension or retraction of the landing gear. CAUTION: Do not cycle the landing gear with the power pack if low on fluid or if the landing gear system is not properly rigged. Use the emergency extension hand pump, TK229/939 hydraulic hand pump or TK229-1/939 air-driven hydraulic pump (Ref. Figure 1, Sheet 8, Chapter 91-00-00) to extend and retract the landing gear for maintenance and rigging. NOTE: Always use a continuity checker to ascertain switch actuation. DO NOT depend upon the click of the switch as an indication of actuation. CAUTION: Excessive compression past 0.86 in. or bottoming out of the switch plunger may cause damage to the switch. Loosen the locking nuts that hold the switch in position. Adjust the switch in an additional two turns beyond the actuation point. However, the uplock switch should not be depressed more than 0.86 in. when the gear is in the up and locked position or the switch may be damaged. Tighten the locking nuts and safety wire. Check for proper indication of the in-transit lights through at least one complete cycle.

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3. MAIN GEAR UP-POSITION INDICATOR SWITCH A. Adjustment (UA-3, UB-1 thru UB-45, Except UB-37, Without Kit No. 114-8005 Installed) WARNING: Make all adjustments with the power off at the master switch and no external power connected. Any time maintenance is to be performed on the landing gear system, place the airplane on jacks. When jacking the airplane in an unsheltered area where winds in excess of 35 kts will be encountered, never jack more than one gear at a time clear of the ground. The landing gear control handle must never be moved from the down-and-locked position while the airplane is on the ground. It is recommended that the area be roped off during extension or retraction of the landing gear. CAUTION: Do not cycle the landing gear with the power pack if low on fluid or if the landing gear system is not properly rigged. Use the emergency extension hand pump, TK229/939 hydraulic hand pump or TK229-1/939 air-driven hydraulic pump (Ref. Figure 1, Sheet 8, Chapter 91-00-00) to extend and retract the landing gear for maintenance and rigging. NOTE: Always use a continuity checker to ascertain switch actuation. DO NOT depend upon the click of the switch as an indication of actuation. CAUTION: Excessive compression past 0.86 in. or bottoming out of the switch plunger may cause damage to the switch. Loosen the locking nuts that hold the switches in position. Adjust each switch in an additional two turns beyond the actuation point. Tighten the locking nuts and safety wire. Check for proper operation of the in-transit lights through at least one complete cycle.

B. Adjustment (UB-37, UB-46 and After, UC-1 and After and Prior Airplanes With Kit No. 114-8005 Installed) WARNING: Make all adjustments with the power off at the master switch and no external power connected. Any time maintenance is to be performed on the landing gear system, place the airplane on jacks. When jacking the airplane in an unsheltered area where winds in excess of 35 kts will be encountered, never jack more than one gear at a time clear of the ground. The landing gear control handle must never be moved from the down-and-locked position while the airplane is on the ground. It is recommended that the area be roped off during extension or retraction of the landing gear. CAUTION: Do not cycle the landing gear with the power pack if low on fluid or if the landing gear system is not properly rigged. Use the emergency extension hand pump, TK229/939 hydraulic hand pump or TK229-1/939 air-driven hydraulic pump (Ref. Figure 1, Sheet 8, Chapter 91-00-00) to extend and retract the landing gear for maintenance and rigging.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL NOTE: Always use a continuity checker to ascertain switch actuation. DO NOT depend upon the click of the switch as an indication of actuation. CAUTION: Excessive compression past 0.86 in. or bottoming out of the switch plunger may cause damage to the switch. With landing gear in the fully retracted position, adjust (H11-1541) switch by rotating the switch's locking nuts four additional turns after both switch sections have changed state electrically. However, the uplock switch should not be depressed more than 0.86 in. when the gear is in the up and locked position or the switch may be damaged. Tighten locking nuts to maintain switch position. Check for proper operation of the gear in-transit lights, while operating landing gear through at least one complete cycle. Safety wire the switch locking nuts to the bracket. CAUTION: For UC-1 and After, if the Main Gear Up-Position Indicator Switch/Wire Harness Assembly is replaced, ensure the harness is secured behind the inner fender to prevent chafing against the fuel line.

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LANDING GEAR NOSE GEAR DOWN-POSITION SWITCH MAINTENANCE PRACTICES

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1. PROCEDURES WARNING: When performing maintenance on a hydraulically operated landing gear system, be aware that any movement of a hydraulic actuator cylinder can displace hydraulic fluid and cause unanticipated movement of other actuator cylinders in the landing gear retraction system. Servicing of the landing gear hydraulic accumulator can also result in unanticipated movement of an actuator. Either action can result in an unsafe, unlocked landing gear system. Therefore, place the airplane on jacks prior to performing any inspection or maintenance. Cycle the landing gear and ensure that all three landing gears are down and locked prior to removing the aircraft from jacks.

A. Adjustment WARNING: Make all adjustments with the power off at the master switch and no external power connected. Any time maintenance is to be performed on the landing gear system, place the airplane on jacks. When jacking the airplane in an unsheltered area where winds in excess of 35 kts will be encountered, never jack more than one gear at a time clear of the ground. The landing gear control handle must never be moved from the down-and-locked position while the airplane is on the ground. It is recommended that the area be roped off during extension or retraction of the landing gear. CAUTION: Do not cycle the landing gear with the power pack if low on fluid or if the landing gear system is not properly rigged. Use the emergency extension hand pump, TK229/939 hydraulic hand pump or TK229-1/939 air-driven hydraulic pump (Ref. Figure 1, Sheet 8, Chapter 91-00-00) to extend and retract the landing gear for maintenance and rigging. NOTE: Always use a continuity checker to ascertain switch actuation. DO NOT depend upon the click of the switch as an indication of actuation. With the landing gear in the fully extended position and the actuator locked (use a continuity checker to verify locking of the actuator by a continuity indication through the normally open set of contacts in the switch), adjust the nose gear down-position switch as follows: NOTE: If the switch bracket becomes bowed, an indication of an unlocked nose gear could result. (1) Position the adjustable stop with approximately 0.25 inch of threads showing (Ref. Figure 201). (2) Loosen the locknuts and adjust the switch actuator housing until the switch plunger contacts the adjustable stop. (3) Tighten the locking nuts and safety wire. (4) Adjust the adjustable stop until the switch actuates as indicated by continuity through the normally open contacts.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL CAUTION: Excessive compression or bottoming out of the switch plunger may cause damage to the switch. (5) Increase the compression of the switch plunger by two (2) more turns of the adjustable stop while the switch remains actuated. (6) Check for proper operation of the GEAR DOWN light, located on the pilot's inboard subpanel, and the landing gear warning horn through at least one complete cycle.

Figure 201 Nose Gear Down-Position Switch Adjustment

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

LANDING GEAR MAIN GEAR DOWN-POSITION SWITCH MAINTENANCE PRACTICES

32-60-02 200200

1. PROCEDURES WARNING: When performing maintenance on a hydraulically operated landing gear system, be aware that any movement of a hydraulic actuator cylinder can displace hydraulic fluid and cause unanticipated movement of other actuator cylinders in the landing gear retraction system. Servicing of the landing gear hydraulic accumulator can also result in unanticipated movement of an actuator. Either action can result in an unsafe, unlocked landing gear system. Therefore, place the airplane on jacks prior to performing any inspection or maintenance. Cycle the landing gear and ensure that all three landing gears are down and locked prior to removing the aircraft from jacks.

A. Adjustment WARNING: Make all adjustments with the power off at the master switch and no external power connected. Any time maintenance is to be performed on the landing gear system, place the airplane on jacks. When jacking the airplane in an unsheltered area where winds in excess of 35 kts will be encountered, never jack more than one gear at a time clear of the ground. The landing gear control handle must never be moved from the down-and-locked position while the airplane is on the ground. It is recommended that the area be roped off during extension or retraction of the landing gear. CAUTION: Do not cycle the landing gear with the power pack if low on fluid or if the landing gear system is not properly rigged. Use the emergency extension hand pump, TK229/939 hydraulic hand pump or TK229-1/939 air-driven hydraulic pump (Ref. Figure 1, Sheet 8, Chapter 91-00-00) to extend and retract the landing gear for maintenance and rigging. With the landing gear in the fully extended position and the actuators locked (verify locking of the actuator by continuity through the normally open set of contacts in the actuator switch), adjust either main gear down-position switch as follows: (1) Loosen the locking nuts that hold the switch in position (Ref. Figure 201). CAUTION: Excessive compression or bottoming out of the switch plunger may cause damage to the switch. (2) Adjust the switch in an additional two (2) turns beyond the actuation point when the landing gear is fully extended. (3) Tighten the locking nuts to hold the switch in position and safety wire. (4) Check for proper operation of the GEAR DOWN light on the pilot's inboard subpanel and the landing gear warning horn through at least one complete cycle.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 201 Main Gear Down-Position Switch Adjustment

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

LANDING GEAR NOSE GEAR ACTUATOR DOWNLOCK SWITCH MAINTENANCE PRACTICES

32-60-03 200200

1. PROCEDURES WARNING: When performing maintenance on a hydraulically operated landing gear system, be aware that any movement of a hydraulic actuator cylinder can displace hydraulic fluid and cause unanticipated movement of other actuator cylinders in the landing gear retraction system. Servicing of the landing gear hydraulic accumulator can also result in unanticipated movement of an actuator. Either action can result in an unsafe, unlocked landing gear system. Therefore, place the airplane on jacks prior to performing any inspection or maintenance. Cycle the landing gear and ensure that all three landing gears are down and locked prior to removing the aircraft from jacks.

A. Adjustment WARNING: Make all adjustments with the power off at the master switch and no external power connected. Any time maintenance is to be performed on the landing gear system, place the airplane on jacks. When jacking the airplane in an unsheltered area where winds in excess of 35 kts will be encountered, never jack more than one gear at a time clear of the ground. The landing gear control handle must never be moved from the down-and-locked position while the airplane is on the ground. It is recommended that the area be roped off during extension or retraction of the landing gear. CAUTION: Do not cycle the landing gear with the power pack if low on fluid or if the landing gear system is not properly rigged. Use the emergency extension hand pump, TK229/939 hydraulic hand pump or TK229-1/939 air-driven hydraulic pump (Ref. Figure 1, Sheet 8, Chapter 91-00-00) to extend and retract the landing gear for maintenance and rigging. The actuator downlock switch is preadjusted and should not required further attention. If the switch must be replaced for any reason, proceed as follows: (1) Perform the NOSE GEAR ACTUATOR REMOVAL procedure (Ref. 32-30-14). (2) Apply 500 to 700 psig of hydraulic pressure to the primary extend port of the actuator to lock the actuator. Disconnect the hydraulic pump from the actuator (Ref. Figure 201). (3) Apply a 500-pound-column load to the actuator piston rod to ensure that the actuator is extended and locked. (4) Check continuity between pins C and F of the switch connector. (5) Turn the switch in until continuity is indicated through the switch contacts. Turn the switch in two full turns past this point. The switch may be turned in a maximum of one additional turn to align the electrical wiring toward the top of the actuator. (6) Apply 400 psig to the retract port of the actuator to unlock the actuator. When the actuator unlocks, there should be no continuity through the switch contacts.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (7) Tighten the jam nut and safety wire. (8) Cycle the actuator three times. Check that the switch closes when the actuator locks and opens when the actuator unlocks by checking for continuity at the switch connector. (9) Perform NOSE GEAR ACTUATOR INSTALLATION Procedure (Ref. 32-30-14), NOSE GEAR RIGGING (Ref. 32-30-13) and, HYDRAULIC SYSTEM FILLING AND BLEEDING procedure (Ref. 32-30-00).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 201 Nose Gear Actuator Downlock Switch

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LANDING GEAR MAIN GEAR ACTUATOR DOWNLOCK SWITCH MAINTENANCE PRACTICES

32-60-04 200200

1. PROCEDURES WARNING: When performing maintenance on a hydraulically operated landing gear system, be aware that any movement of a hydraulic actuator cylinder can displace hydraulic fluid and cause unanticipated movement of other actuator cylinders in the landing gear retraction system. Servicing of the landing gear hydraulic accumulator can also result in unanticipated movement of an actuator. Either action can result in an unsafe, unlocked landing gear system. Therefore, place the airplane on jacks prior to performing any inspection or maintenance. Cycle the landing gear and ensure that all three landing gears are down and locked prior to removing the aircraft from jacks.

A. Adjustment (All Airight Actuators Except P/N 114-380041-1) WARNING: Make all adjustments with the power off at the master switch and no external power connected. Any time maintenance is to be performed on the landing gear system, place the airplane on jacks. When jacking the airplane in an unsheltered area where winds in excess of 35 kts will be encountered, never jack more than one gear at a time clear of the ground. The landing gear control handle must never be moved from the down-and-locked position while the airplane is on the ground. It is recommended that the area be roped off during extension or retraction of the landing gear. CAUTION: Do not cycle the landing gear with the power pack if low on fluid or if the landing gear system is not properly rigged. Use the emergency extension hand pump, TK229/939 hydraulic hand pump or TK229-1/939 air-driven hydraulic pump (Ref. Figure 1, Sheet 8, Chapter 91-00-00) to extend and retract the landing gear for maintenance and rigging. The actuator downlock switch is preadjusted and should not require further attention. If the switch must be adjusted or replaced for any reason, proceed as follows (Ref. Figure 201): (1) Perform the MAIN GEAR ACTUATOR REMOVAL procedure (Ref. 32-30-10). (2) Place the actuator in a test fixture. Apply 500 to 700 psig of hydraulic pressure to the primary retract port of the actuator to lock the actuator. Remove the source of hydraulic pressure. (3) Apply a 500-pound tension load to the actuator piston rod to ensure that the actuator is retracted and locked. (4) Connect a continuity checker between pins A and C of the switch connector. The checker should indicate an open circuit. Note the position of the electrical wiring. (5) Remove the small screw that secures the tang washer located next to the switch. Loosen the switch jam nut to move the tang washer and allow a wrench to be positioned on the fitting located where the switch enters the upper body of the actuator. (6) Slowly turn the fitting clockwise until the continuity checker indicates continuity.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (7) Slowly turn the fitting counterclockwise until the continuity checker indicates an open circuit (loss of continuity). Continue turning the fitting counterclockwise one to two full turns past this point. Align the wiring toward the top of the actuator. (8) Apply 400 psig to the extend port of the actuator to unlock the actuator and extend the rod. The continuity checker should indicate continuity. (9) Position the tang washer and tighten the jam nut. Install the screw into the tang washer and safety wire. (10) Cycle the actuator three times, monitoring the continuity checker to ensure that the switch is indicating an open circuit when retracted and locked and continuity when unlocked and extending. (11) Disconnect the continuity checker from the actuator. (12) Perform the MAIN GEAR ACTUATOR INSTALLATION procedure (Ref. 32-30-10).

B. Adjustment (Phoenix Controls Actuators and Airight Actuator P/N 114-380041-1) WARNING: Make all adjustments with the power off at the master switch and no external power connected. Any time maintenance is to be performed on the landing gear system, place the airplane on jacks. When jacking the airplane in an unsheltered area where winds in excess of 35 kts will be encountered, never jack more than one gear at a time clear of the ground. The landing gear control handle must never be moved from the down-and-locked position while the airplane is on the ground. It is recommended that the area be roped off during extension or retraction of the landing gear. CAUTION: Do not cycle the landing gear with the power pack if low on fluid or if the landing gear system is not properly rigged. Use the emergency extension hand pump, TK229/939 hydraulic hand pump or TK229-1/939 air-driven hydraulic pump (Ref. Figure 1, Sheet 8, Chapter 91-00-00) to extend and retract the landing gear for maintenance and rigging. The actuator downlock switch is preadjusted and should not require further attention. If the switch must be adjusted or replaced for any reason, proceed as follows (Ref. Figure 201): (1) Perform the MAIN GEAR ACTUATOR REMOVAL procedure (Ref. 32-30-10). (2) Place the actuator in a test fixture. Apply 500 to 700 psig of hydraulic pressure to the primary retract port of the actuator to lock the actuator. Remove the source of hydraulic pressure. (3) Apply a 500-pound tension load to the actuator piston rod to ensure that the actuator is retracted and locked. (4) Connect a continuity checker between pins A and C of the switch connector. The checker should indicate an open circuit. Note the position of the electrical wiring. (5) Loosen the switch jam nut. (6) Slowly turn the switch counterclockwise until the continuity checker indicates continuity.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (7) Slowly turn the switch clockwise until the continuity checker indicates an open circuit (loss of continuity). Continue turning the switch clockwise one to two full turns past this point. Align the wiring toward the top of the actuator. (8) Apply 400 psig to the extend port of the actuator to unlock the actuator and extend the rod. The continuity checker should indicate continuity. (9) Tighten the jam nut and safety wire. (10) Cycle the actuator three times, monitoring the continuity checker to ensure that the switch is indicating an open circuit when retracted and locked and continuity when unlocked and extending. (11) Disconnect the continuity checker from the actuator. (12) Perform the MAIN GEAR ACTUATOR INSTALLATION procedure (Ref. 32-30-10).

C. Adjustment (Frisby Actuators) WARNING: Make all adjustments with the power off at the master switch and no external power connected. Any time maintenance is to be performed on the landing gear system, place the airplane on jacks. When jacking the airplane in an unsheltered area where winds in excess of 35 kts will be encountered, never jack more than one gear at a time clear of the ground. The landing gear control handle must never be moved from the down-and-locked position while the airplane is on the ground. It is recommended that the area be roped off during extension or retraction of the landing gear. CAUTION: Do not cycle the landing gear with the power pack if low on fluid or if the landing gear system is not properly rigged. Use the emergency extension hand pump, TK229/939 hydraulic hand pump or TK229-1/939 air-driven hydraulic pump (Ref. Figure 1, Sheet 8, Chapter 91-00-00) to extend and retract the landing gear for maintenance and rigging. The actuator downlock switch is preadjusted and should not require further attention. If the switch must be adjusted or replaced for any reason, proceed as follows (Ref. Figure 201): (1) Perform the MAIN GEAR ACTUATOR REMOVAL procedure (Ref. 32-30-10). (2) Place the actuator in a test fixture. Apply 500 to 700 psig of hydraulic pressure to the primary retract port of the actuator to lock the actuator. Remove the source of hydraulic pressure. (3) Apply a 500-pound tension load to the actuator piston rod to ensure that the actuator is retracted and locked. (4) If this procedure is being performed to adjust the switch currently installed in the actuator, back the switch out of the actuator, as required, to ensure it is not being activated. (5) Connect an ohmmeter between pins A and C of the switch connector. The ohmmeter should indicate an open circuit.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (6) Slowly screw the switch into the actuator until the ohmmeter indicates continuity. Unscrew the switch two full turns. The switch may be turned one additional full turn to align the tab washer with the slot in the head assembly of the actuator. The ohmmeter should indicate and open circuit. (7) Apply 500 psig to the extend port of the actuator to unlock the actuator. The ohmmeter should indicate continuity. (8) Torque the jam nut to 25 ±5 inch-pounds and safety wire. (9) Apply a bead of sealant (RTV-162) between the nut on the switch and the switch, and between the nut and the actuator head assembly. (10) Cycle the actuator three times, using a hydraulic source with a pressure of 500 psig. Monitor the ohmmeter to ensure that the switch is actuating properly. (11) Disconnect the ohmmeter from the actuator. (12) Perform the MAIN GEAR ACTUATOR INSTALLATION procedure (Ref. 32-30-10).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 201 Main Gear Actuator Downlock Switch

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LANDING GEAR MAIN GEAR SAFETY SWITCH MAINTENANCE PRACTICES

32-60-05 200200

1. PROCEDURES WARNING: When performing maintenance on a hydraulically operated landing gear system, be aware that any movement of a hydraulic actuator cylinder can displace hydraulic fluid and cause unanticipated movement of other actuator cylinders in the landing gear retraction system. Servicing of the landing gear hydraulic accumulator can also result in unanticipated movement of an actuator. Either action can result in an unsafe, unlocked landing gear system. Therefore, place the airplane on jacks prior to performing any inspection or maintenance. Cycle the landing gear and ensure that all three landing gears are down and locked prior to removing the aircraft from jacks.

A. Adjustment (444EN49-6 Switch) WARNING: Make all adjustments with the power off at the master switch and no external power connected. Any time maintenance is to be performed on the landing gear system, place the airplane on jacks. When jacking the airplane in an unsheltered area where winds in excess of 35 kts will be encountered, never jack more than one gear at a time clear of the ground. The landing gear control handle must never be moved from the down-and-locked position while the airplane is on the ground. It is recommended that the area be roped off during extension or retraction of the landing gear. CAUTION: Do not cycle the landing gear with the power pack if low on fluid or if the landing gear system is not properly rigged. Use the emergency extension hand pump, TK229/939 hydraulic hand pump or TK229-1/939 air-driven hydraulic pump (Ref. Figure 1, Sheet 8, Chapter 91-00-00) to extend and retract the landing gear for maintenance and rigging. This procedure is applicable to both the right and left hand safety switches (Ref. Figure 201). (1) Perform the THREE POINT JACKING (PREFERRED (Ref. Chapter 07-10-00). All tires must be clear of the floor.

PROCEDURE)

procedure

(2) Pull the 2-ampere control circuit breaker on the pilot's inboard subpanel and place a note on the circuit breaker panel that LANDING GEAR MAINTENANCE IS IN PROCESS during maintenance on the landing gear. (3) With the shock strut fully extended, check the actuator rod (3) and attached rod end for clearance with the upper torque knee (Ref. Figure 201). (4) If clearance is not sufficient: (a) Check the upper torque knee eye bolt for washers. There should be four 0.063 inch thick washers under the eye bolt head. If washers are added, check the eye bolt shank for sufficient length to allow nut engagement.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (b) Clearance can be increased a small amount by adjusting the rod end. The actuator rod (3) must be threaded into the rod end 0.38 to 0.50 inch. Verify the actuator rod covers the inspection hole in the rod end. (5) Press the strut air valve and completely deflate the strut. Remove valve core and attach a 1/4 inch I.D. hose to the air valve and connect a container to the other end to catch any possible fluid spillage. (6) Remove the cotter pin, nut (6), Washer (5) and bolt (4) to disconnect the actuator rod (3) (Ref. Figure 201). NOTE: Use a gear-type puller (obtain locally) to remove the switch arm (9) from the switch shaft to prevent damage to the internal mechanism of the switch. Do Not Pry Off With Screwdrivers. (7) Remove the retaining nut (8) and the switch arm (9) from the switch shaft. (8) Disconnect the switch wiring from the airplane wiring and connect the test box (30, Table 7, Chapter 91-00-00) to the switch wiring or connect the wire leads from an ohmmeter to pins “T” and “U” of the receptacle plug located in the upper rear of each wheel well. (9) With the shock strut fully extended, using a suitable marker, mark the shock strut piston at 0.38, 0.62 and 2.0 inches from the bottom of the brace assembly. (10) Position a jack under the landing gear. Jack the landing gear so the shock strut is compressed to 0.38 to 0.62 inch from the fully extended position. (11) Rotate switch shaft counterclockwise until the ohmmeter indicates a closed circuit or the test box red lights illuminate then continue rotating counterclockwise until the ohmmeter indicates an open circuit or the test box green lights illuminate. (12) Install the switch arm (9) on the switch shaft. Tighten retaining nut (8) on shaft to engage splines then back off the nut (Ref. Figure 201). (13) Install bolt (4), washer (5), nut (6) and cotter pin to connect the actuator rod (3) to the upper torque knee eye bolt. (14) Remove the safety wire from the lock screw (10) on the switch arm (9) and back off the lock screw. (15) Adjust the switch shaft clockwise at the adjusting screw (7) until the ohmmeter indicates a closed circuit on pins “T” and “U” or the test box red lights illuminate. (16) Tighten the retaining nut (8) on the shaft and tighten the lock screw (10) ensuring that the ohmmeter still indicates a closed circuit or the test box red lights are still illuminated. NOTE: The point at which the switch is actuated during compression and extension of the shock strut differs from the tolerances in the switch and its attendant linkage on the landing gear. (17) Check the safety switch rigging as follows: (a) Fully compress the shock strut. (b) Allow the shock strut to extend from the fully compressed position, ensure the green test lights are illuminated or the ohmmeter continues to indicate an open circuit until the shock strut reaches a position 0.38 to 2.0 inches from the fully extended position. At this point, the Page 202 Nov 1/09

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL red test lights illuminate or the ohmmeter will indicate a closed circuit up to and including the fully extended position. (c) From fully extended compress the strut. The red lights of the test box will extinguish and the green lights will illuminate or the ohmmeter indication will change from a closed circuit to an open circuit when the shock strut reaches a position 0.38 to 2 inches from the fully extended position. The green test lights will remain illuminated or the ohmmeter will continue to indicate an open circuit as the shock strut is fully compressed. WARNING: Before removing the airplane from the jacks, make sure that the landing gear emergency-extend hand pump handle is in the stowed position, the plunger on the service valve is pushed down with the hinged retainer in place, the landing gear control handle is in the DOWN position, the landing gear is down and locked and the accumulator is charged to 800 ± 50 psi. (18) For fine adjustment, use the adjusting screw (7) on the switch arm (9). NOTE: Ensure that the retaining nut (8) on the switch shaft and the lock screw (10) on the switch arm (9) are tight after each adjustment. (19) Safety wire the lock screw (10) to the switch arm (9). (20) Connect the switch wiring to the airplane by connecting the receptacle plug located in the upper rear of each wheel well. (21) Perform the MAIN LANDING GEAR SHOCK ABSORBER SERVICING procedure (Ref. 32-10-00). (22) Perform the LOWERING THE AIRPLANE AFTER THREE-POINT JACKING procedure (Ref. Chapter 07-10-00).

B. Adjustment (39EN6-6 Switch) WARNING: Make all adjustments with the power off at the master switch and no external power connected. Any time maintenance is to be performed on the landing gear system, place the airplane on jacks. When jacking the airplane in an unsheltered area where winds in excess of 35 kts will be encountered, never jack more than one gear at a time clear of the ground. The landing gear control handle must never be moved from the down-and-locked position while the airplane is on the ground. It is recommended that the area be roped off during extension or retraction of the landing gear. CAUTION: Do not cycle the landing gear with the power pack if low on fluid or if the landing gear system is not properly rigged. Use the emergency extension hand pump, TK229/939 hydraulic hand pump or TK229-1/939 air-driven hydraulic pump (Ref. Figure 1, Sheet 8, Chapter 91-00-00) to extend and retract the landing gear for maintenance and rigging. This procedure is applicable to both the right and left hand safety switches. (1) Perform the THREE POINT JACKING (PREFERRED (Ref. Chapter 07-10-00). All tires must be clear of the floor.

PROCEDURE)

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procedure

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (2) Pull the 2-ampere control circuit breaker on the pilot's inboard subpanel and place a note on the circuit breaker panel that LANDING GEAR MAINTENANCE IS IN PROCESS during maintenance on the landing gear. (3) With the shock strut fully extended, check the actuator rod (3) and attached rod end for clearance with the upper torque knee (Ref. Figure 201). (4) If clearance is not sufficient: (a) Check the upper torque knee eye bolt for washers. There should be four 0.063 inch thick washers under the eye bolt head. If washers are added, check the eye bolt shank for sufficient length to allow nut engagement. (b) Clearance can be increased a small amount by adjusting the rod end. The actuator rod (3) must be threaded into the rod end 0.38 to 0.50 inch. Verify the actuator rod covers the inspection hole in the rod end. (5) Press the strut air valve and completely deflate the strut. Remove valve core and attach a 1/4 inch I.D. hose to the air valve and connect a container to the other end to contain any possible fluid spillage. (6) Remove the cotter pin, nut (6), washer (5) and bolt (4) to disconnect the actuator rod (3) (Ref. Figure 201). NOTE: Use a gear-type puller (obtain locally) to remove the switch arm (9) from the switch shaft to prevent damage to the internal mechanism of the switch. Do Not Pry Off With Screwdrivers. (7) Remove the retaining nut (8) and the switch arm (9) from the switch shaft. (8) Disconnect the switch wiring from the airplane wiring and connect the test box (30, Table 7, Chapter 91-00-00) to the switch wiring or connect the wire leads from an ohmmeter to pins “T” and “U” of the receptacle plug located in the upper rear of each wheel well. (9) With the shock strut fully extended, using a suitable marker, mark the shock strut piston at 0.38, 0.62 and 2.0 inches from the bottom of the brace assembly. (10) Position a jack under the landing gear. Jack the landing gear so the shock strut is compressed to 0.38 to 0.62 inch from the fully extended position. (11) Rotate switch shaft counterclockwise until the ohmmeter indicates a closed circuit or the test box red lights illuminate then continue rotating counterclockwise until the ohmmeter indicates an open circuit or the test box green lights illuminate. (12) Install the switch arm (9) on the switch shaft. Tighten retaining nut (8) on shaft to engage splines then back off the nut (Ref. Figure 201). (13) Install bolt (4), washer (5), nut (6) and cotter pin to connect the actuator rod (3) to the upper torque knee eye bolt. (14) Remove the safety wire from the locking nuts (12) on the switch arm (9) and unlock the locking nuts.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (15) Using an allen wrench, adjust the gear lever counterclockwise until the ohmmeter indicates a closed circuit or the test box red lights illuminate then continue rotating counterclockwise until the ohmmeter indicates an open circuit or the test box red lights extinguish. (16) Using an allen wrench, adjust the hex head adjusting screw (13) clockwise in the switch arm (9) until the ohmmeter indicates a closed circuit on pins “T” and “U” or the test box red lights illuminate. (17) Tighten the locking nuts (12) on the switch arm (9) ensuring that the switch still indicates closed. Do not apply more than 20 to 25 inch pounds torque when tightening the lock nuts. NOTE: The point at which the switch is actuated during compression and extension of the shock strut differs from the tolerances in the switch and its attendant linkage on the landing gear. (18) Check the safety switch rigging as follows: (a) Fully compress the shock strut. (b) Allow the shock strut to extend from the fully compressed position, ensure the green test lights are illuminated or the ohmmeter continues to indicate an open circuit until the shock strut reaches a position 0.38 to 2.0 inches from the fully extended position. At this point, the red test lights illuminate or the ohmmeter will indicate a closed circuit up to and including the fully extended position. (c) From fully extended compress the strut. The red lights of the test box will extinguish and the green lights will illuminate or the ohmmeter indication will change from a closed circuit to an open circuit when the shock strut reaches a position 0.38 to 2 inches from the fully extended position. The green test lights will remain illuminated or the ohmmeter will continue to indicate an open circuit as the shock strut is fully compressed. WARNING: Before removing the airplane from the jacks, make sure that the landing gear emergency-extend hand pump handle is in the stowed position, the plunger on the service valve is pushed down with the hinged retainer in place, the landing gear control handle is in the DOWN position, the landing gear is down and locked and the accumulator is charged to 800 ± 50 psi. (19) For fine adjustment, use the hex head adjusting screw (13) on the switch arm (9). NOTE: Ensure that the retaining nut (8) on the switch shaft is tight after each adjustment. (20) Safety wire the locking nuts (12) on the switch arm (9). (21) Connect the switch wiring to the airplane by connecting the receptacle plug located in the upper rear of each wheel well. (22) Perform the MAIN LANDING GEAR SHOCK ABSORBER SERVICING procedure (Ref. 32-10-00). (23) Perform the LOWERING THE AIRPLANE AFTER THREE-POINT JACKING procedure (Ref. Chapter 07-10-00).

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2 9

10 SAFETY SWITCH P/N 444EN49-6

8

3

7

A 5

4

6

DETAIL

A

1 13 12

1. UPPER TORQUE KNEE 2. SWITCH 3. ACTUATOR ROD 4. BOLT 5. WASHER 6. NUT 7. ADJUSTING SCREW (UNDERSIDE) 8. RETAINING NUT 9. SWITCH ARM 10. LOCKING SCREW 11. *SWITCH ARM 12. *LOCK NUT 13. *HEX HEAD ADJUSTING SCREW

SAFETY SWITCH P/N 39EN6-6 11 8 12 13

*P/N39EN6-6 DETAIL

A UC32B 070150AA.AI

Figure 201 Main Landing Gear Safety Switch Adjustment

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LANDING GEAR LANDING GEAR WARNING HORN MAINTENANCE PRACTICES

32-60-06 200200

1. LANDING GEAR WARNING HORN WARNING: When performing maintenance on a hydraulically operated landing gear system, be aware that any movement of a hydraulic actuator cylinder can displace hydraulic fluid and cause unanticipated movement of other actuator cylinders in the landing gear retraction system. Servicing of the landing gear hydraulic accumulator can also result in unanticipated movement of an actuator. Either action can result in an unsafe, unlocked landing gear system. Therefore, place the airplane on jacks prior to performing any inspection or maintenance. Cycle the landing gear and ensure that all three landing gears are down and locked prior to removing the aircraft from jacks.

A. Check WARNING: Do not adjust the landing gear warning horn switches using this procedure. Perform the LANDING GEAR WARNING HORN ADJUSTMENT procedure in this section. (1) Perform the THREE POINT JACKING (PREFERRED (Ref. Chapter 07-10-00). All tires must be clear of the floor.

PROCEDURE)

procedure

(2) Perform APPLYING GROUND POWER procedure (Ref. Chapter 24-40-00). (3) Move the Engine Power levers fully forward. (4) Raise the landing gear electrically. (5) Verify that flaps are in the full UP position. (6) Retard one power lever until the warning horn sounds. The warning lights in the landing gear control handle should activate along with the horn. Advance the lever to silence the horn and to extinguish the warning lights. (7) Retard the opposite power lever until the warning horn sounds. The warning lights in the landing gear control handle should activate along with the horn. Advance the lever to silence the horn and to extinguish the warning lights. (8) Retard both power levers together until the warning horn sounds. Use the silence button next to the gear selector to silence the horn, then advance the levers back to full power. (9) Lower the flaps to the TAKEOFF and then to the APPROACH position. The warning horn will sound and the warning lights will activate when the flaps are lowered to these positions. Use the silence button to silence the horn for these flap positions. (10) Lower the flaps beyond the APPROACH position. The warning horn will activate and will continue to sound at any power lever position. (11) Attempt to silence the horn by using the silence button next to the gear selector. The horn shall continue to sound. Raise flaps to 0° UP and push the warning horn silence button. (12) Lower the landing gear electrically and make sure all three landing gear are down and locked.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (13) Perform the LOWERING THE AIRPLANE AFTER THREE-POINT JACKING procedure (Ref. Chapter 07-10-00). (14) Perform REMOVING GROUND POWER procedure (Ref. Chapter 24-40-00).

B. Adjustment A cam-operated switch is connected to each power lever linkage in the pedestal. The switches are wired so that moving either or both levers beyond the safe flight setting with the gear retracted will sound the warning horn intermittently. The airplane must be in flight when determining the position of the power levers, which correspond to 84% to 86% N1 (gas generator speed), as the ram air effect will alter the 84% to 86% N1 position indicated during ground operation (Ref. Figure 201). (1) With the airplane in flight, advance the power levers until 84% to 86% N1 is attained on each engine. (2) Mark the position of the power levers on the pedestal with a piece of tape. Land the airplane. (3) Remove the lower upholstery panels from both sides of the pedestal. (4) Place the power levers in alignment with the tape on the pedestal. (5) Adjust the landing gear warning horn switches by moving them in their slotted mounting brackets to just actuate in this position. (6) Move the power levers and listen for an audible “click” when the power levers are moved aft of the tape. (7) Remove the tape and install the pedestal upholstery. (8) Flight test the airplane with the landing gear retracted and the power levers retarded until the warning horn sounds at an N1 speed of 84% to 86%.

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B

DETAIL

A

A PROP CONTROL BELLCRANK PROP FEATHER DETENTS

BETA SWITCHES TO ACTUATE AT BETA DETENT 2 (PLACES)

THROTTLE CONTROL BELLCRANK

DETAIL

B

UC32B 063492AA.AI

Figure 201 Landing Gear Warning Horn Switches

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CHAPTER 33 - LIGHTS TABLE OF CONTENTS SUBJECT

PAGE

GENERAL INFORMATION 33-00-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1

FLIGHT COMPARTMENT 33-10-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Annunciator System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Troubleshooting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 101 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 101 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Control Wheel Map Light Bulb . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 206 Replacement . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 206 Annunciator Light Bulb . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 206 Replacement . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 206 Dimmer Switch Power Supply . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 206 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 206 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 206 Edgelighted Panel Assemblies . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 207 Replacement . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 207 Printed Circuit Board (PCB) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 210 Repair Procedure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 210 Cleaning . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 210 Conformal Coating . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 210 Extinguisher Push/Firewall Valve Switch . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 211 Lamp Replacement . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 211

LANDING GEAR INDICATOR LIGHTS 33-10-05 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .201 Landing Gear Downlock Lights Lamp . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .201 Replacement . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .201 Landing Gear Downlock Lights Annunciator. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .201 Functional Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .201

PASSENGER COMPARTMENT 33-20-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Entry Light Bulb . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Replacement . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Spar Cover Light Tube . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Replacement . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Cabin Reading Light Bulb . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202

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CHAPTER 33 - LIGHTS TABLE OF CONTENTS (CONTINUED) SUBJECT PAGE Replacement . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202 Passenger No Smoking/Fasten Seat Belt Sign Light Tray . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202 Replacement . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202

BAGGAGE AND CARGO COMPARTMENT 33-30-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Nose Baggage Compartment Light Bulb . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Replacement . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Aft Baggage Compartment Light Bulb . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Replacement . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201

EXTERIOR 33-40-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Strobe Light System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Recognition Lights . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Navigation Lights . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Taxi Light . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Landing Lights . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Wing Ice Lights . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Empennage Logo Lights . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2 Rotating Beacons . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2 Troubleshooting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 101 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 101 Strobe Light System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 101 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Wing Strobe Light . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Wing Strobe Light Flashtube . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202 Replacement . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202 Tail Strobe Light Flashtube . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202 Replacement . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202 Recognition Light Bulb . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 203 Replacement . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 203 Navigation Light Bulb . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 203 Replacement . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 203 Taxi Light Bulb . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 203 Replacement . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 203 Wing Ice Light Bulb . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 203 Replacement . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 203 Landing Light Bulb . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 204 Replacement . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 204 Empennage Logo Light Bulb . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 204 Replacement . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 204

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

CHAPTER 33 - LIGHTS TABLE OF CONTENTS (CONTINUED) SUBJECT PAGE Rotating Beacon Light Bulb . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 204 Replacement . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 204 Tail Strobe Light Power Supply . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 205 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 205 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 205

SELF-ILLUMINATED SIGNS 33-50-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Self-Illuminated Exit Sign (UA-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Replacement . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Self-Illuminated Exit Sign (UB-1 and After; UC-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 203 Replacement . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 203 Self-Illuminated Exit Sign Brightness Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 205

EMERGENCY EXIT LIGHTING SYSTEM (OPTIONAL INSTALLATION WITH EXTERNAL FLOODLIGHTS) (UC-1 AND AFTER) 33-50-01 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Emergency Exit Cabin Light Lamp . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Replacement . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Emergency Exit Exterior Light Lamp . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Replacement . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201

EMERGENCY EXIT LIGHTING 33-50-02 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 JAR-OPS Emergency Exit Lighting System (Optional on UC-1 and After, and with Kit 114-5313-3 Installed) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 JAR-OPS Emergency Exit Lighting System (Optional On UC-1 and After, and with Kit 114-5313-3) . . . . 201 Emergency Exit Cabin Light Lamp . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Replacement . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Emergency Exit Light . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202 Battery Replacement . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

LIGHTS GENERAL INFORMATION DESCRIPTION AND OPERATION

33-00-00 00

1. GENERAL Flight compartment lighting consists of annunciator panel lights, electroluminescent panels, indicator lights, map lights, an overhead light, lights installed in the landing gear control handle and lights in the fire extinguisher pull handle. Passenger compartment lighting consists of overhead fluorescent lights, reading lights, airstair door lights, entry way lights and spar cover lights. Lights located in the baggage and cargo compartments are controlled by switches located in the compartments. Visibility of the airplane exterior is improved by the use of strobe lights, rotating beacons, navigation lights, and recognition lights. Landing lights and taxi lights illuminate the runway. The empennage logo lights illuminate the identification numbers located on the vertical stabilizer.

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LIGHTS FLIGHT COMPARTMENT DESCRIPTION AND OPERATION

33-10-00 00

1. GENERAL The brightness of the electroluminescent panels, glareshield indirect instrument lights, engine instrument and flight instrument direct lights, avionics panel lights and overhead lights is controlled by dimmer switches located on the overhead light control panel. Power supplies for the electroluminescent panel dimmer circuits and the engine instrument lights dimmer circuit are located on the cabin forward electrical equipment panel. Refer to the Model 1900 Airliner Series Component Maintenance Manual, P/N 114-590021-11, for further information on the power supply for the electroluminescent panels. Flight compartment lights can be controlled by the master panel light switch located on the overhead light control panel. The indirect instrument lights on the glareshield can be controlled by the emergency lights switch located on the overhead meter panel. Annunciator system circuitry includes fault detection cards, time delay printed circuit boards, an annunciator control card, and circuit breakers. Refer to Chapter 39-20-00 for maintenance practices on these components.

A. Annunciator System The annunciator system consists of a warning annunciator panel located on the glareshield, a caution/ advisory panel located on the center subpanel, a press-to-test switch, pilot's and copilot's master warning flasher lights, and pilot's and copilot's master caution flasher lights. The warning annunciator lights are red, the caution lights are yellow, and the advisory lights are green. The press-to-test switch is located adjacent to the warning annunciator panel. The master warning flashers and master caution flashers are located on the glareshield. The master warning flashers are red and the master caution flashers are yellow. Individual warning and caution lights incorporate wording to indicate which airplane system is called to the operator's attention. If an airplane system fault requires the immediate attention of the pilot, the master warning light (red) on the glareshield will flash. The flashing warning light may be extinguished by pressing the face of the light, resetting the circuit. However, the warning annunciator light will illuminate and remain illuminated if the fault is not corrected or cannot be corrected. If an additional fault occurs, the appropriate annunciator panel light will illuminate and the master warning light will again flash. When an airplane system fault does not require the operator's immediate attention or action, the master caution light (yellow) will flash and the appropriate light on the caution/advisory annunciator panel will illuminate. The flashing master caution light may be extinguished by pressing the face of the light. An additional fault will illuminate the caution/advisory light and the master caution light again. The advisory annunciator lights (green) indicate that a system is operating under certain conditions. The master warning and master caution flasher lights are not associated with the advisory system. Generally, current necessary to illuminate the annunciator lights is taken from the appropriate system's control circuits; however, on airplanes having the ANN PWR SOURCE annunciator light in the caution/ advisory panel, certain annunciator lights receive their current from the annunciator power source Printed Circuit Board (PCB) (A141) (Ref. Figure 1). The annunciator power source PCB precludes the possibility of losing power to an annunciator should a fault develop within the control circuits of an affected system. Additionally, to further minimize the possibility of losing power to an annunciator, the annunciator power source PCB provides power to the associated annunciator lights from two separate sources: the triple fed bus and a generator bus.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL The circuitry to each light being fed from the annunciator power source PCB is protected by a fuse in each power source feeder. The fuses are mounted on the PCB and can be replaced by field service personnel. Should any one of the PCB fuses open or a power failure occur on either one of the power supplying buses, the yellow ANN PWR SOURCE annunciator light will be illuminated in the caution/ advisory panel. This is an announcement to the flight crew that an annunciator light being powered by the annunciator power source PCB has experienced a partial or complete loss of power. The lights that are powered by the annunciator power source PCB are as follows: R and L FUEL FEED, R and L FW VALVE, R and L BK DI OVHT, R and L ENG ICE FAIL, HYD FLUID LOW, BATT CHARGE, MAN STEER FAIL, PWR STEER FAIL, R and L ENVIR OFF, ANTI SKID FAIL, INVERTER, FWD CABIN DOOR, AFT CABIN DOOR, R and L OIL PRESS, CABIN ALT, R and L ENVIR FAIL, and ANN PWR SOURCE. The annunciator panel lights function in a “bright” or a “dim” mode. The system functions in the “dim” mode automatically when all of the following conditions exist: a generator is on the line; the overhead flood light is OFF; the pilot's instrument lights are ON; and the ambient light sensor senses light below a preset value. Unless all these conditions exist, the annunciator panel lights will function in the “bright” mode. The ambient light sensor is located on the overhead light control panel.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 1 Annunciator System Electrical Schematic

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

LIGHTS FLIGHT COMPARTMENT TROUBLESHOOTING

100100

1. PROCEDURES Refer to Chart 101 in Figure 101 and Figure 102 to troubleshoot the annunciator system.

Figure 101 Troubleshooting - Annunciator System

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 102 Annunciator System Electrical Schematic

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL ANNUNCIATOR PANELS NOMENCLATURE

COLOR

CAUSE OF ILLUMINATION WARNING ANNUNCIATOR PANEL

L FUEL PRESSURE

Red

Fuel pressure low on LH side.

CABIN ALTITUDE

Red

Cabin altitude exceeds 12,500 feet. (1)

BAGGAGE DOOR

Red

Nose baggage door is not latched.

INVERTER

Red

The inverter selected is inoperative. (1)

R FUEL PRESS

Red

Fuel pressure low on RH side.

L OIL PRESS

Red

Oil pressure low on LH side. (1)

L ENVIR FAIL

Red

LH bleed air overpressure or overtemperature condition detected. (1)

FWD CABIN DOOR

Red

Forward cabin door is not latched. (1)

R ENVIR FAIL

Red

RH bleed air overpressure or overtemperature condition detected. (1)

R OIL PRESS

Red

Oil pressure low on RH side. (1)

AFT CABIN DOOR

Red

Aft cabin door is not latched. (1)

L BL AIR FAIL

Red

LH bleed air fail switch has actuated.

A/P TRIM FAIL

Red

Improper trim or no trim from autopilot trim command. (2)

A/P DISC

Red

Autopilot is disconnected. (2)

R BL AIR FAIL

Red

RH bleed air fail switch has actuated. CAUTION/ADVISORY ANNUNCIATOR PANEL

L DC GEN

Yellow

LH generator is off the line.

L FUEL QTY

Yellow

Quantity of fuel in the LH system is low.

BATTERY CHARGE

Yellow

Excessive charge rate on battery. (1)

BAT TIE OPEN

Yellow

Battery isolated from the generator bus.

R FUEL QTY

Yellow

Quantity of fuel in RH system is low.

R DC GEN

Yellow

RH generator is off the line.

L FW VALVE

Yellow

LH fuel firewall valve has not reached its selected position. (1)

L FUEL FEED

Yellow

LH fuel feed level is low. (1)

L GEN TIE OPEN

Yellow

LH generator bus is isolated from center bus.

R GEN TIE OPEN

Yellow

RH generator bus is isolated from center bus.

R FUEL FEED

Yellow

RH fuel feed level is low. (1)

R FW VALVE

Yellow

RH fuel firewall valve has not reached its selected position. (1)

L ENG ICE FAIL

Yellow

LH ice vane electric actuator is not functioning properly. (1)

L BK DI OVRHT

Yellow

LH overheat pressure switch on brake deice system has actuated. (1) (2)

HYD FLUID LOW

Yellow

Oil level in landing gear power pack is low. (1)

MAN STEER FAIL

Yellow

Nose wheel is not free to caster when power steering is off. (1) (2)

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL ANNUNCIATOR PANELS (Continued) NOMENCLATURE

COLOR

CAUSE OF ILLUMINATION

CAUTION/ADVISORY ANNUNCIATOR PANEL (Continued) R BK DI OVHT

Yellow

RH overheat pressure switch on brake deice system has actuated. (1) (2)

R ENG ICE FAIL

Yellow

RH ice vane electric actuator is not functioning properly. (1)

L CHIP DETECT

Yellow

Metallic contamination is detected in the LH engine oil. (2)

PWR STEER FAIL

Yellow

Power steering system failure. (1) (2)

ANN PWR SOURCE

Yellow

Partial power loss to some annunciator lights. (1)

R CHIP DETECT

Yellow

Metal contamination is detected in the RH engine oil. (2)

ANTI SKID FAIL

Yellow

Anti skid system failure.

L NO FUEL XFER

Yellow

Senses no fuel transfer from the aux to main tank.

R NO FUEL XFER

Yellow

Senses no fuel transfer from the aux to main tank.

PWR STEER ARM

Green

Senses power steering is ready to operate (UC-1 and After).

L AUTOFEATHER

Green

LH autofeather is armed and power levers are advanced above 90% N1.

L IGNITION ON

Green

LH starter/ignition switch is in the engine ignition mode or the LH auto ignition system is armed and LH engine torque is below 400 ft-lbs.

TAXI LIGHT

Green

Taxi light switch is in the ON position and the landing gear is retracted.

EXTERNAL POWER

Green

External power is connected to the airplane.

R IGNITION ON

Green

RH starter/ignition switch is in the engine ignition mode or the RH engine torque is below 400 ft-lbs when the RH auto ignition system is armed.

R AUTOFEATHER

Green

RH autofeather is armed and power levers are advanced above 90% N1.

L ENG ANTI-ICE

Green

LH engine anti-ice vane is extended.

L BK DEICE ON

Green

LH brake deice system is energized. (2)

ELEC TRIM OFF

Green

Electric trim is de-energized by a trim disconnect switch on the control wheel when the system power switch on the pedestal is in the ON position. (2)

MAN TIES CLOSE

Green

Generator bus ties have closed manually.

R BK DEICE ON

Green

RH brake deice system is de-energized. (2)

R ENG ANTI-ICE

Green

RH engine anti-ice vane is in the extended position.

L ENVIR OFF

Green

LH environmental bleed air valve is closed. (1)

FUEL TRANSFER

Green

Fuel transfer valve is open.

AIR CON N1 LOW

Green

RH engine RPM is too low to allow air-conditioning load.

R ENVIR OFF

Green

RH environmental bleed air valve is closed. (1)

1. These annunciator lights are powered through the annunciator power source PCB A141. 2. These annunciators are for optional systems and may not be installed in all airplanes.

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LIGHTS FLIGHT COMPARTMENT MAINTENANCE PRACTICES

200200

1. PROCEDURES Refer to Table 201 and Figure 201 to replace bulbs in the flight compartment. Table 201 Bulb Replacement Guide LOCATION

P/N

Annunciator Panel Lights (Warning and Caution/Advisory)

327

Cabin Pressure Controls Electroluminescent Panel

71565

Circuit Breakers Electroluminescent Panel

71581

Electroluminescent Subpanel, Pilot's Outboard

71560

Electroluminescent Subpanel, Pilot's Inboard

71561

Electroluminescent Subpanel, Copilot's Outboard

71563

Electroluminescent Subpanel, Copilot's Inboard

71562

Engine Fire Warning Lights

327

Engine Fire Extinguisher Pull Handle Lights

327

Fuel Controls Upper Electroluminescent Panel

71558

Fuel Controls Lower Electroluminescent Panel

71559

Instrument Indirect Lights (Glareshield)

1864

Instrument Direct Lights

327

Landing Gear Control Handle Light

327

Landing Gear Indicator Lights

327

Map Light (Control Wheel)

MS25069-1495

Map Overhead Light

303

Master Caution Light (Yellow)

387

Master Warning Light (Red)

387

Outside Air Temperature Indicator Lights

327

Overhead Controls Electroluminescent Panel

71557

Overhead Instruments Electroluminescent Panel

71566

Pedestal Upper Electroluminescent Panel

71564

Pedestal Lower Electroluminescent Panel

71566

Post Light

327

Standby Compass Light

327

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Figure 201 (Sheet 1 of 4) Bulb Replacement Guide - Flight Compartment

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Figure 201 (Sheet 2 of 4) Bulb Replacement Guide - Flight Compartment

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Figure 201 (Sheet 3 of 4) Bulb Replacement Guide - Flight Compartment

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Figure 201 (Sheet 4 of 4) Bulb Replacement Guide - Flight Compartment

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2. CONTROL WHEEL MAP LIGHT BULB A. Replacement (1) Remove the attaching screws and the escutcheon from the control wheel (Ref. Figure 201). (2) Remove the bulb by pressing it into the socket and turning counterclockwise. (3) Install a new bulb in the socket. (4) Install the escutcheon and secure with the attaching screws.

3. ANNUNCIATOR LIGHT BULB A. Replacement (1) Press and release the lens assembly to partially eject it from the annunciator assembly (Ref. Figure 201). (2) Pull the lens assembly from the annunciator assembly. (3) Remove the bulb from the socket. (4) Install a new bulb. (5) Insert the lens assembly in the annunciator assembly.

4. DIMMER SWITCH POWER SUPPLY A. Removal (1) Remove the carpet (Ref. Chapter 25-20-01, CARPET REMOVAL AND INSTALLATION) and floorboards in the center of the passenger compartment adjacent to the forward airstair door (Ref. Chapter 6-50-00, FLOOR ACCESS PANELS). (2) Remove the electrical connector from the power supply (Ref. Figure 202). (3) Remove the attaching screws and the power supply from the electrical equipment panel.

B. Installation (1) Install the power supply on the electrical equipment panel and secure with the attaching screws (Ref. Figure 202). (2) Install the electrical connector on power supply. (3) Install the floorboards in the center of the passenger compartment. (4) Install the carpet previously removed (Ref. Chapter 25-20-01, CARPET REMOVAL AND INSTALLATION).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

5. EDGELIGHTED PANEL ASSEMBLIES The edgelighted panel assemblies consist of an outer plastic panel and a printed circuit board to which wheat lights are soldered. These panel assemblies are mounted strategically throughout the crew compartment to identify the function and mode of the various switches and systems. The inscriptions on the outer, counterbored, plastic panel of the edgelighted assemblies are illuminated of wheat light lamps soldered in parallel to a printed circuit board and powered by 28 vdc. Due to the fact that edgelighted panels can be damaged by inappropriate maintenance practices and contamination, usually moisture resulting from spills or storm windows left open in moist environments, they should be functionally checked during each routine checking of all lights.

A. Replacement (1) Disconnect the airplane power. (2) Remove any knobs that may interfere with the removal of the panel assemblies. NOTE: Exercise care not to scratch or damage the panel assemblies. (3) Remove the attaching screws and lift the edgelighted panel away from the structure (Ref. Figure 203). (4) Position the new edgelighted panel in place and install it with the attaching screws. DO NOT OVERTIGHTEN THE SCREWS. (5) Install all knobs which were removed for this operation. (6) Restore electrical power to the airplane.

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Figure 202 Dimmer Switch Power Supply Installation

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Figure 203 Edgelighted Panel Replacement

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6. PRINTED CIRCUIT BOARD (PCB) A. Repair Procedure Because of the bonded type construction of the board, repair of the PCB assembly is limited to replacement of the wheat light lamps. The following Steps should be observed when replacement of a lamp (or lamps) becomes necessary. (1) Utilizing a low wattage soldering iron, heat the solder holding each of the lamp leads to the PCB. As the solder begins to soften, lift the lead away from its mounting pad. CAUTION: The lamp mounting pads and leads may lift away from the board if excessive heat is applied to the solder attaching the lamp leads to the mounting pads. If this occurs, the PCB must be replaced. (2) Solder the new wheat light lamp to the appropriate pads on the PCB. Do not apply excessive solder to the assembly, and ensure that each connection has a good fillet between the pad and the lead. The contour of each lamp lead should be visible in the solder. (3) After completing repairs, check that all lamps match the counterbores in the outer plastic panel. (4) Connect a regulated 28 vdc power source to the PCB and verify that all lamps illuminate. When normal panel function is attained, the panel must be cleaned and a conformal coating applied before final assembly (Ref. PRINTED CIRCUIT BOARD CLEANING and PRINTED CIRCUIT BOARD - CONFORMAL COATING).

B. Cleaning (1) Immerse the entire PCB in a container of isopropyl alcohol (30, Table 1, Chapter 91-00-00). NOTE: The PCB may be lightly scrubbed with a soft bristled brush as necessary to remove stubborn contaminants. (2) Assure that all traces of flux and other contaminants are washed away. (3) Remove the PCB from the container and rinse it by pouring clean isopropyl alcohol over the panel assembly. NOTE: After cleaning, care must be exercised to protect from the environment. All traces of oil, moisture or any other contaminants must be kept away from the PCB. Clean, lint-free, white gloves or clean rubber gloves must be worn at all times during and after the cleaning of a PCB assembly. (4) Immediately after rinsing, place the assembly in a vertical position (to draw off any residual moisture) and allow to dry.

C. Conformal Coating (1) As soon as possible after the PCB is fully dry, the conformal coating should be applied. The coating must completely cover the side of the PCB on which the light lamps are installed.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL NOTE: Hawker Beechcraft Corporation recommends coating (108 or 109, Table 1, Chapter 91-00-00) thinned 25% with solvent (110, Table 1, Chapter 91-00-00). CE1155 is a two part product and should be mixed in a ratio of 10 parts of component A to 7 parts component B by weight. CE1155 is the preferred coating agent; however, any coating that conforms to MIL-I-46058 may be used. (2) The coating may be applied by dipping, brush or spray. Whichever coating method is used, care must be taken to avoid getting any coating in the area of the electrical connector. After application, the coating must be cured as recommended in the manufacturer's instructions. (3) Inspect the coating for complete coverage, fisheyes, bubbles, dry spots, peeling, blisters, wrinkles, and cracks. If any coating defects are found, the panel must be cleaned and have the conformal coating applied. (4) When the conformal coating has been satisfactorily applied and cured, the edgelighted panel may be assembled and installed. Replace any knobs that were removed and connect the airplane power.

7. EXTINGUISHER PUSH/FIREWALL VALVE SWITCH A. Lamp Replacement CAUTION: Failure to disconnect the battery from the “hot” battery bus before lamp replacement may result in accidental discharge of the engine fire extinguishers or closing of the firewall valves. (1) Remove all electrical power from the airplane and disconnect the airplane battery. (2) Remove the copper safety wire from the clear plastic guard. Flip the clear plastic guard up, grasp the finger grips on the sides of the legend face, and pull outward (Ref. Figure 205). (3) While fully extended, rotate the legend face a 1/4 turn in either direction. (4) Push in on the legend face to release the lamp holder tabs from their slots. (5) Pull the lamp holder out and swivel it down to access the lamps. (6) Remove colored boots from the lamps as required. (7) Extract the four lamps from the lamp holder with fingers or a small screwdriver. Always replace the lamps as a set. (8) Install the new lamps. (9) Fully extend the legend face to engage the tab cam. (10) Retract the tabs with the tab cam by rotating the legend face 1/4 turn in either direction. (11) Once the tabs are retracted, swivel the lamp holder up and slide it all the way into the switch body. NOTE: Install the legend face with the nomenclature facing upright. The lamp holder will not seat properly if the legend face is inverted. (12) Extend the tabs by rotating the legend face 1/4 turn. The lamp holder will lock into place as the tabs extend.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (13) Push the legend face into the switch body until it clicks. An audible click indicates the legend face has locked in place. (14) Safety wire the clear plastic guard closed with 30 gage (0.010 diameter) copper safety wire. (15) Connect the airplane battery.

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Figure 204 Edgelighted Panel Printed Circuit Board Lamp Replacement

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Figure 205 Fire Extinguisher and Firewall Valve Switch

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LIGHTS LANDING GEAR INDICATOR LIGHTS MAINTENANCE PRACTICES

33-10-05 200200

1. LANDING GEAR DOWNLOCK LIGHTS LAMP A. Replacement WARNING: When replacing lamps, never move the landing gear handle from the extended (down) position. (1) Remove electrical power from the airplane. (2) Pull the annunciator module from the annunciator assembly (Ref. Figure 201). CAUTION: Any time this annunciator is removed, perform LANDING GEAR DOWNLOCK LIGHT ANNUNCIATOR FUNCTIONAL CHECK. (3) Remove and replace the defective lamp (Ref. Table 1, 33-10-00). (4) Ensure proper orientation of the annunciator module and press it firmly into the annunciator assembly. WARNING: Verify the landing gear switch handle is in the extended (down) position before restoring electrical power to the airplane. (5) Apply power to the airplane and verify operation of the annunciator.

2. LANDING GEAR DOWNLOCK LIGHT ANNUNCIATOR A. FUNCTIONAL CHECK WARNING: When replacing lamps, never move the landing gear handle from the extended (down) position. (1) Remove electrical power from the airplane. (2) Pull the annunciator module from the annunciator assembly (Ref. Figure 201). (3) Remove one bulb at a time and shine a flashlight from the back of the annunciator forward. (4) Ensure that the appropriate portion of the annunciator illuminates at the location where the bulb is removed. CAUTION: The annunciator has dividers installed to block the light between the LH and the RH portion of the annunciator. The nose location, either bulb, will light the entire NOSE section. (5) Continue checking until all of the removed bulb locations are checked. (6) Install all bulbs and install annunciator assembly. (7) Restore power to the airplane and verify all bulbs are illuminated.

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A

PILOT’S INBOARD SUBPANEL

LANDING GEAR INDICATORS

LANDING GEAR HANDLE

DETAIL

A UB33B 111542AA.AI

Figure 201 Landing Gear Downlock Annunciator Location

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LIGHTS PASSENGER COMPARTMENT DESCRIPTION AND OPERATION

33-20-00 00

1. GENERAL Passenger compartment lighting consists of entry way and aisle lighting, overhead lighting, reading lights and warning lights. Lights located in the airstair door and doorpost can be controlled by a switch located on the side of the airstair door step. A door-lock light may be illuminated by pressing a switch located adjacent to the entry lights switch. The fluorescent overhead lights, reading lights and the lights in the no smoking/fasten seat belt signs are controlled by switches located on the overhead light control panel in the flight compartment. The reading lights may be controlled by push button switches located adjacent to the lights.

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LIGHTS PASSENGER COMPARTMENT MAINTENANCE PRACTICES

200200

1. PROCEDURES Refer to Table 201 and Figure 201, 202 and 203 for replacement of bulbs located in the passenger compartment.

2. ENTRY LIGHT BULB A. Replacement (1) Remove the attaching screws and lens from the aft door post on the airstair door (Ref. Figure 201 and 202). (2) Remove the bulb by pressing it into the socket and turning counterclockwise. (3) Install a new bulb in the socket. (4) Install the lens and secure with the attaching screws.

3. SPAR COVER LIGHT TUBE A. Replacement (1) Remove the attaching screws and the lens from the fluorescent light fixture (Ref. Figure 203). (2) Remove the fluorescent tube from the fixture. (3) Install a new fluorescent tube in the fixture. (4) Install the lens and secure with the attaching screws.

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Table 201 BULB REPLACEMENT GUIDE Location

P/N

Aft Dome Light

303

Airstair Doorpost Light

1864

Airstair Door-Lock Light

1864

Aisle Entry Light

303

LH Forward Partition Entry Light

1495

No Smoking/Fasten Seat Belt Signs

1202-300 (UA, UB, & UC-1 thru UC-57)

No Smoking/Fasten Seat Belt Signs

OL6839BPE (UC-58 & after and earlier serials with Kit No. 114-3015-1S installed)

Overhead Fluorescent Light

5108 WW

Reading Light

303

Spar Cover Fluorescent Light

5104 WW

4. CABIN READING LIGHT BULB A. Replacement (1) Remove the threaded sleeve from the light receptacle (Ref. Figure 204). (2) Remove the lens and the bulb from the socket. (3) Install a new bulb in the socket. (4) Install the lens and secure with the threaded sleeve.

5. PASSENGER NO SMOKING/FASTEN SEAT BELT SIGN LIGHT TRAY A. Replacement (1) Snap off the cover. The sign may remain mounted on the bulkhead while this is being accomplished. (2) Remove the two screws which retain the lamp tray in the light wedge. (3) Replace the lamp tray with a new lamp tray and secure it with the two screws. (4) Replace the snap-on cover.

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Figure 201 Forward Partition Entry Light - Passenger Compartment

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Figure 202 Airstair Door Entry Lights

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Figure 203 Spar Cover Light - Passenger Compartment

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Figure 204 Overhead Fluorescent/ Light Cabin Reading Light - Passenger Compartment

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LIGHTS BAGGAGE AND CARGO COMPARTMENT DESCRIPTION AND OPERATION

33-30-00 00

1. GENERAL The nose baggage compartment light is controlled by a switch located adjacent to the light. The aft baggage compartment light is controlled by a switch located adjacent to the lower aft corner of the aft airstair door. The cargo compartment dome light is controlled by a switch located adjacent to the lower forward corner of the cargo door.

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LIGHTS BAGGAGE AND CARGO COMPARTMENT MAINTENANCE PRACTICES

200200

1. PROCEDURES Refer to Table 201 and Figure 201 and 202 to replace bulbs installed in the airplane baggage and cargo compartments. Table 201 Bulb Replacement Guide Location

P/N

Aft Baggage Compartment Light

303

Cargo Compartment Light

303

Nose Baggage Compartment Light

303

2. NOSE BAGGAGE COMPARTMENT LIGHT BULB A. Replacement (1) Open the nose baggage compartment door. (2) Remove the attaching screws and lens from the light (Ref. Figure 201). (3) Remove the bulb by pressing it into the socket and turning counterclockwise. (4) Install a new bulb in the socket. (5) Install the lens and secure it with the attaching screws. (6) Install the cover on the lens. (7) Close the nose baggage compartment door.

3. AFT BAGGAGE COMPARTMENT LIGHT BULB A. Replacement (1) Remove the attaching screws, lens cover and lens from the light (Ref. Figure 202). (2) Remove the bulb by pressing it into the socket and turning counterclockwise. (3) Install a new bulb in the socket. (4) Install the lens and lens cover then secure them with the attaching screws.

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Figure 201 Nose Baggage Compartment Light

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Figure 202 Aft Baggage Compartment Light

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LIGHTS EXTERIOR DESCRIPTION AND OPERATION

33-40-00 00

1. GENERAL A. Strobe Light System Strobe lights installed on the aft end of the vertical stabilizer and on the wing tips increase visibility of the airplane during night flight. A toggle switch located on the LH inboard subpanel controls the strobe light system. The toggle switch incorporates a 5-amp circuit breaker. The wing strobe lights power supply units are installed adjacent to the strobe lights. The vertical stabilizer strobe light power supply is installed on a bulkhead located forward of the stabilizer strobe light. Transistorized circuitry in the power supply units steps up 28-vdc airplane power, producing between 400 and 600 volts. The stepped-up voltage is stored in two capacitors until a timer releases the voltage to the strobe light assembly. The high voltage stored in the capacitor then surges through the gas in the flashtube to produce a brilliant burst of light.

B. Recognition Lights The recognition lights are installed in each wing tip, forward of the strobe lights. The lights are aimed forward and outboard of the airplane. The recognition lights are controlled by a toggle switch located on the pilot's inboard subpanel. The switch incorporates a 7.5-amp circuit breaker.

C. Navigation Lights Navigation lights are installed in each wing tip and on the aft end of the vertical stabilizer fairing. The lights are controlled by a toggle switch located on the pilot's inboard subpanel. The toggle switch incorporates a 5-amp circuit breaker.

D. Taxi Light The taxi light is installed on the nose landing gear strut and illuminates the runway when the landing gear is extended. A toggle switch incorporating a 15-amp circuit breaker is located on the pilot's inboard subpanel and controls the taxi light. If the taxi light control switch is in the ON position when the landing gear is retracted, an advisory panel light will illuminate.

E. Landing Lights Landing lights are installed in the leading edge of each wing and are controlled by LH and RH toggle switches. The switches are located on the pilot's inboard subpanel and incorporate circuit breakers.

F. Wing Ice Lights Wing ice lights aid the pilot in detecting the formation of ice on the wing leading edge. The lights are installed on the outboard side of each engine nacelle and are controlled by a toggle switch located on the pilot's inboard subpanel. The switch incorporates a 5-amp circuit breaker.

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G. Empennage Logo Lights Logo lights are installed on the underside of the horizontal stabilizer and illuminate the identification logo on the vertical stabilizer. A toggle switch located on the pilot's inboard subpanel allows control of the lights. The switch incorporates a 15-amp circuit breaker.

H. Rotating Beacons The rotating beacons are anti-collision lights installed on the underside of the fuselage and on the vertical stabilizer fairing. The beacons are controlled by a toggle switch located on the pilot's inboard subpanel. The switch incorporates a 10-amp circuit breaker. The bulbs are mounted in a motor that rotates each bulb approximately 180°. These rotating lights increase airplane visibility from all directions.

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LIGHTS EXTERIOR TROUBLESHOOTING

100100

1. PROCEDURES A. Strobe Light System The following procedures are recommended when a strobe light flashtube fails to illuminate: (1) Ensure that the flashtube is not broken and that the electrical connectors are installed securely. WARNING: Although a bleed-off resistor is incorporated in the power supply circuit, high voltage is involved in the circuit between the power supply and light assemblies. For this reason, turn the control switch for the strobe lights off and wait at least ten minutes before disconnecting the cables at the power supply and before handling or disassembling these units in any way. Failure to observe these precautions may result in physical injury or shock. (2) Disconnect the power cable from the inoperative light. Connect the power cable to a light known to be good. If the flashtube of the good light fails to illuminate, the strobe light circuit breaker has tripped or the power supply unit has failed. CAUTION: Never connect the wiring of a good power supply to the circuit of an inoperative light, because a short in the defective light would damage the power supply. The most likely reasons for malfunctions of this system are: (1) shorts in the power supply unit or the lamp assembly, (2) shorts caused by contact of the power supply unit with a foreign object during operation, (3) moisture in the connectors and (4) excessive heat within the power supply unit.

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LIGHTS EXTERIOR MAINTENANCE PRACTICES

200200

1. PROCEDURES Refer to Table 201 and Figure 201 through 207 to replace bulbs located on the airplane exterior. Table 201 BULB REPLACEMENT GUIDE Location

P/N

Empennage Logo Light

1982

Nose Taxi Light

4587

Recognition Light

1982

Rotating Beacon (Upper)

A7079B-24

Rotating Beacon (Lower)

A7079B-24

Strobe Flashtubes (Wing and Vertical Stabilizer)

55-0221-1

Tail Navigation Light

1683

Wing Ice Light

A7079B-24

Wing Landing Lights

4553

Wing Navigation Lights

A7512-24

2. WING STROBE LIGHT A. Removal WARNING: Although a bleed-off resistor is incorporated in the power supply circuit, high voltage is involved in the circuit between the power supply and light assemblies. For this reason, turn the control switch for the strobe lights off and wait at least ten minutes before disconnecting the cables at the power supply and before handling or disassembling these units in any way. Failure to observe these precautions may result in physical injury or shock. (1) Remove the transparent wing tip cover. (2) Remove the attaching screws and lift out the light assembly (Ref. Figure 201). (3) Disconnect the electrical connector from the power supply and remove the complete unit from the airplane.

B. Installation (1) Connect the electrical connector at the aft end of the power supply to the wing wiring (Ref. Figure 201). (2) Position the light assembly in the wing tip and secure with the attaching screws.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (3) Install the transparent wing tip cover and secure with the attaching screws.

3. WING STROBE LIGHT FLASHTUBE A. Replacement WARNING: Although a bleed-off resistor is incorporated in the power supply circuit, high voltage is involved in the circuit between the power supply and light assemblies. For this reason, turn the control switch for the strobe lights off and wait at least ten minutes before disconnecting the cables at the power supply and before handling or disassembling these units in any way. Failure to observe these precautions may result in physical injury or shock. (1) Remove the transparent wing tip cover (Ref. Figure 201). (2) Remove two screws from the strobe light end plate and remove the end plate. (3) Slide the strobe light lens outboard and off the light assembly. (4) Remove the flashtube from the clips and discard the flashtube. (5) Place a clean, lint free cloth around a new flashtube to prevent fingers from contacting the glass and install the new flashtube with the indexing clip pointing outward. WARNING: If fingers contact the flashtube, clean the area thoroughly with a clean, lint free cloth and isopropyl alcohol and allow to air dry. (6) Install the strobe light lens, the end plate and the wing tip cover.

4. TAIL STROBE LIGHT FLASHTUBE A. Replacement WARNING: Although a bleed-off resistor is incorporated in the power supply circuit, high voltage is involved in the circuit between the power supply and light assemblies. For this reason, turn the control switch for the strobe lights off and wait at least ten minutes before disconnecting the cables at the power supply and before handling or disassembling these units in any way. Failure to observe these precautions may result in physical injury or shock. (1) Remove the transparent cover from the aft end of the vertical stabilizer fairing (Ref. Figure 201). (2) Remove the screws which attach the strobe light assembly to the bulkhead and pull the strobe light assembly out of the opening in the fairing bulkhead. (3) Remove the mounting bracket and end plate from one end of the strobe light assembly and slide the lens off the strobe light housing. (4) Remove and discard the flashtube. WARNING: If fingers contact the flashtube, clean the area thoroughly with a clean, lint free cloth and isopropyl alcohol and allow to air dry.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (5) Place a clean, lint free cloth around a new flashtube to keep fingers from contacting the glass and install the new flashtube with the indexing clip pointing aft. (6) Install the strobe light lens, end plate and bracket, install the assembly in the tail fairing and install the transparent cover on the aft end of the tail fairing.

5. RECOGNITION LIGHT BULB A. Replacement (1) Remove the wing tip light transparent cover (Ref. Figure 201). (2) Remove the two attaching screws and pull the reflector assembly off the light bulb. (3) Press the bulb into the socket and turn counterclockwise to remove. (4) Install a new bulb in the lamp socket, holding the bulb with a clean, lint free cloth to keep fingers from contacting the glass. WARNING: If fingers contact the flashtube, clean the area thoroughly with a clean, lint free cloth and isopropyl alcohol and allow to air dry. (5) Install the reflector assembly and the transparent cover on the wing tip.

6. NAVIGATION LIGHT BULB A. Replacement NOTE: The following replacement procedures apply to the wing navigation lights and the tail navigation light. (1) Remove the transparent light cover from the stabilizer fairing (Ref. Figure 201). (2) Remove the screws, lens retainer, and lens. (3) Press the lamp into its socket and turn counterclockwise to remove. (4) Install the retaining ring and secure with the attaching screw.

7. TAXI LIGHT BULB A. Replacement (1) Remove the screw and the retaining ring from the light housing (Ref. Figure 203). (2) Lift the bulb out of its housing and disconnect the electrical wiring from bulb. (3) Connect the wiring to a new bulb and install the bulb in the housing. (4) Install the retaining ring and secure with the attaching screw.

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8. WING ICE LIGHT BULB A. Replacement (1) Remove the upper center cowling (Ref. Chapter 71-10-00, COWLING REMOVAL) (Ref. Figure 204). (2) Press the bulb into its socket and turn counterclockwise to remove. (3) Install a new bulb in the socket, holding the bulb with a clean, lint free cloth. (4) Install the upper center cowling (Ref. Chapter 71-10-00, COWLING INSTALLATION).

9. LANDING LIGHT BULB A. Replacement (1) Remove the attaching screws and lens assembly from the wing leading edge (Ref. Figure 205). (2) Remove the attaching screws and nuts securing the retaining ring. Remove the retaining ring. (3) Remove the bulb and disconnect the electrical wiring from the bulb. (4) Connect the electrical wiring to a new bulb. Install the bulb and secure with the retaining ring and the attaching screws and nuts. (5) Install the lens assembly and secure with the attaching screws.

10. EMPENNAGE LOGO LIGHT BULB A. Replacement (1) Remove the attaching screws, retainer and lens from the underside of the horizontal stabilizer (Ref. Figure 206). (2) Press the bulb into its socket and turn counterclockwise to remove. (3) Install a new bulb in the socket, holding the bulb with a clean, lint free cloth. (4) Install the lens and retainer and secure with the attaching screws. Seal around the periphery of the retainer using sealer (60, Table 1, Chapter 91-00-00).

11. ROTATING BEACON LIGHT BULB A. Replacement (1) Remove the attaching screw and washer from the lens of the beacon. Loosen the screw on the aft end of the beacon and remove the lens (Ref. Figure 202). (2) Remove the bulb from the rotation motor socket. (3) Install a new bulb in the socket. (4) Install the lens and seal and secure with the attaching screw and washer. Tighten the screw on the aft end of the beacon. Page 204 Nov 1/09

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12. TAIL STROBE LIGHT POWER SUPPLY A. Removal WARNING: Although a bleed-off resistor is incorporated in the power supply circuit, high voltage is involved in the circuit between the power supply and light assemblies. For this reason, turn the control switch for the strobe lights off and wait at least ten minutes before disconnecting the cables at the power supply and before handling or disassembling these units in any way. Failure to observe these precautions may result in physical injury or shock. (1) Remove the attaching screws and the aft section of the vertical stabilizer fairing (Ref. Figure 207). (2) Remove the electrical connector from the power supply. (3) Remove the attaching parts and power supply from the airplane.

B. Installation (1) Install the power supply and secure with the attaching screws and washers (Ref. Figure 207). (2) Install the electrical connector on the power supply. (3) Install the vertical stabilizer fairing and secure with the attaching screws.

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Figure 201 Wing and Tail Strobe Light

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Figure 202 Beacon Light Bulb Replacement

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Figure 203 Taxi Light

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Figure 204 Wing Ice Light

33-40-00

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Figure 205 Landing Light

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Figure 206 Empennage Logo Light

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Figure 207 Vertical Stabilizer Strobe Light Power Supply

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LIGHTS SELF-ILLUMINATED SIGNS DESCRIPTION AND OPERATION

33-50-00 00

1. GENERAL Airstair door and emergency door exit signs are manufactured with T-lights. These lights are sealed glass tubes filled with tritium gas and are phosphor-coated internally. In principle, T-lights produce light much the same way as television tubes or luminous watch dials; no exposure to daylight is necessary. T-lights have a useful life of many years; however, FAA regulations require a minimum operational brightness of 100 microlamberts. Periodic brightness checks must be conducted to insure that the brightness level remains at or above 100 microlamberts. Refer to Chapter 4-00-00 for inspection interval information.

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LIGHTS SELF-ILLUMINATED EXIT SIGNS MAINTENANCE PRACTICES

200200

1. PROCEDURES 2. SELF-ILLUMINATED EXIT SIGN (UA-1 AND AFTER) A. Replacement Self-illuminated exit signs are located on all three emergency escape hatches (Ref. Figure 201, View A). A self-illuminated exit sign is located in the vestibule area on either the forward LH partition (View C), mounted to the forward escutcheon (View D) or mounted to the headliner centered above the forward airstair door (View B). Another self-illuminated sign is located above and centered over the aft airstair door headliner (View B). The self-illuminated exit signs can be replaced as follows: (1) Carefully remove the self-illuminated exit sign from its mounting surface. (2) Remove all adhesive residue from the mounting surface. (3) Scuff both the new self-illuminating sign and the mounting surface bonding area using 240 grit sandpaper or scrubbing pad (150, Table 1, Chapter 91-00-00). (4) Clean the bonding surfaces using solvent (30 or 54, Table 1, Chapter 91-00-00). (5) Bond the self-illuminating exit sign to the mounting surfaces using adhesive (186, Table 1, Chapter 91-00-00) per BS201.

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BL 20.000 REF 5.00 INCHES FROM TOP OF PARTITION

1.00 INCH LOCATED AT FS 150.60

VIEW

C

C

B

D

VIEW

D

WL 135.00 REF

A A

B

VIEW

B

VIEW

UA33B 071355AA.AI

Figure 201 Self-Illuminated Sign Installation (UA-1 and After)

Page 202 Nov 1/09

A

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3. SELF-ILLUMINATED EXIT SIGN (UB-1 AND AFTER; UC-1 AND AFTER) A. Replacement Self-illuminated exit signs are located on all three emergency escape hatches (Ref. Figure 202, View A). A self-illuminated exit sign is located in the vestibule area on either the forward LH partition (View C), mounted to the forward escutcheon (View B) or mounted to the headliner centered above the forward airstair door (View D). The self-illuminated exit signs can be replaced as follows: (1) Carefully remove the self-illuminated exit sign from its mounting surface. (2) Remove all adhesive residue from the mounting surface. (3) Scuff both the new self-illuminating sign and the mounting surface bonding area using 240 grit sandpaper or scrubbing pad (150, Table 1, Chapter 91-00-00). (4) Clean the bonding surfaces using solvent (30 or 54, Table 1, Chapter 91-00-00). (5) Bond the self-illuminating exit sign to the mounting surfaces using adhesive (186, Table 1, Chapter 91-00-00) per BS201.

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BL 20.000 REF 5.00 INCHES FROM TOP OF PARTITION

1.00 INCH LOCATED AT FS 150.60

VIEW

C

C

D

B

VIEW

B

WL 135.00 REF

A A

VIEW VIEW

A

D

Figure 202 Self-Illuminated Sign Installation (UB-1 and After; UC-1 and After)

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

4. SELF-ILLUMINATED EXIT SIGN The equipment used in this check consists of DB-45-B1 T-light Comparator (Ref. Figure 203). This equipment is available from Self-Powered Lighting Inc., 8 Westchester Plaza, Elmsford, New York 10523. Self-illuminated exit signs can be inspected on the airplane under all types of lighting conditions without the need for removing the sign. Checking for brightness requires the use of a T-light Comparator. The test is performed as follows:

A. Brightness Check (1) Place the comparator base plate flat on the surface of the sign and move it gently until the viewing aperture is over the portion of the sign to be inspected. (2) Compare the brightness of the sign portion seen through the aperture and the calibrated source; if the sign is as bright or brighter than the calibrated source, the sign brightness is equivalent to, or better than, the 100 microlamberts required. If the sign brightness is less than that of the comparator, the sign must be replaced. (3) Move the comparator over all the illuminated portions of the sign to check minimum brightness at all points. (4) If the sign at any point appears less bright than the comparator, the sign is due for replacement.

Figure 203 T-Light Comparator

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LIGHTS SELF-ILLUMINATED SIGNS DESCRIPTION AND OPERATION

33-50-00 00

1. GENERAL Airstair door and emergency door exit signs are manufactured with T-lights. These lights are sealed glass tubes filled with tritium gas and are phosphor-coated internally. In principle, T-lights produce light much the same way as television tubes or luminous watch dials; no exposure to daylight is necessary. T-lights have a useful life of many years; however, FAA regulations require a minimum operational brightness of 100 microlamberts. Periodic brightness checks must be conducted to insure that the brightness level remains at or above 100 microlamberts. Refer to Chapter 04-00-00 of the 1900 Airliner Series Airworthiness Limitations Manual P/N 129-590000-133 for inspection interval information.

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LIGHTS EMERGENCY EXIT LIGHTING SYSTEM (OPTIONAL INSTALLATION WITH EXTERNAL FLOODLIGHTS) (UC-1 AND AFTER) MAINTENANCE PRACTICES 1. PROCEDURES 2. EMERGENCY EXIT CABIN LIGHT LAMP A. Replacement (1) Ensure the emergency light switch placarded OFF-ARM-ON, in the overhead panel, is in the OFF position. (2) Remove the lens by unsnapping it from the overhead emergency exit cabin light. NOTE: The emergency exit light has one large lamp and two small lamps located at the bottom side of the floodlight reflector, and four miniature lamps located around the perimeter of the lens assembly. (3) Gently press the flood lamp into the lamp socket and turn counterclockwise (approximately a 1/8 rotation), then pull the lamp from the socket. (4) Press a new lamp into the lamp socket and turn clockwise until movement stops. Lightly pull the lamp up to assure that it is properly locked in place. (5) To remove the smaller bulbs, unscrew the cap counterclockwise. (6) To install the smaller bulbs, screw the cap clockwise. (7) Position the lens on the overhead emergency light and snap in place.

3. EMERGENCY EXIT EXTERIOR LIGHT LAMP A. Replacement (1) Remove the four screws from the lens assembly. (2) Remove the lens assembly from the emergency exit light. (3) Press the lamp into the lamp socket and turn it counterclockwise (approximately 1/8 rotation), then pull the lamp from the socket. (4) Press a new lamp Into the lamp socket and turn it clockwise (approximately 1/8 rotation). (5) Lightly pull on the lamp to assure that it is properly locked in place. (6) Position the lens assembly on the emergency exit light and install the four attaching screws.

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LIGHTS EMERGENCY EXIT LIGHTING DESCRIPTION AND OPERATION

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1. GENERAL A. JAR-OPS Emergency Exit Lighting System (Optional on UC-1 and After, and with kit 114-5313-3 Installed) NOTE: This equipment is installed on airplanes operating in accordance with Joint Aviation Requirements (JAR-OPS 1). It is installed by Kit No. 114-5313-3. The emergency exit lighting system is a self-activating self-contained lighting system designed to provide emergency lighting to the interior areas around the cabin door and three emergency exits (Ref. Figure 1). Three lighting modules are located in the headliner of the cabin: one opposite the cabin door, one opposite the right forward emergency exit, and one between the left and right emergency exits. Each module is equipped with two light sources. A dim light source, which is illuminated during normal operations is powered by the 28 volt aircraft battery through a 5 amp circuit breaker on the right-hand circuit breaker panel. The other, brighter light source is powered by four internal alkaline batteries and is controlled by two switches (Ref. Figure 2). One is an internal “g” switch which activates upon sensing rapid deceleration in the forward direction. The other is an external three-position rocker switch (spring-loaded to the center position) placarded ON-TEST-/OFF-RESET. The light illuminates when the switch is momentarily placed in the ON-TEST position. The light extinguishes when the switch is momentarily placed in the OFF-RESET position.

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Figure 1 JAR-OPS Emergency Exit Lighting System (With Kit 114-5313-3 Installed)

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Figure 2 JAR-OPS Emergency Exit Lights Schematic

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LIGHTS EMERGENCY EXIT LIGHTING MAINTENANCE PRACTICES

200200

1. PROCEDURES 2. JAR-OPS EMERGENCY EXIT LIGHTING SYSTEM (OPTIONAL ON UC-1 AND AFTER, AND WITH KIT 114-5313-3) NOTE: This equipment is installed on airplanes operating in accordance with Joint Aviation Requirements (JAR-OPS 1).

3. EMERGENCY EXIT CABIN LIGHT LAMP A. Replacement (1) Ensure that the emergency light switch (1) placarded ON-TEST/OFF-RESET is OFF (Ref. Figure 201). (2) Unsnap the lens (3) from the overhead emergency exit cabin light and remove it. (3) Gently press the large lamp (9) into the lamp socket and turn counterclockwise (approximately a 1/8 turn), then pull the lamp from the socket. (4) Press a new lamp (9) into the lamp socket and turn clockwise until movement stops. Lightly pull the lamp to assure that it is properly locked in place. (5) To remove the smaller bulbs from the light assembly (10), unscrew the cap counterclockwise. (6) To install the smaller bulbs, screw the cap clockwise. (7) Position the lens (3) on the overhead emergency light and snap in place. Table 201 Lamp Replacement Guide For JAR-OPS Emergency Exit Lighting System LAMP

LAMP P/N

Overhead Emergency Exit Cabin Light (Large Lamp)

425

Overhead Emergency Exit Cabin Light (Small Lamp)

123-C-STC-FB59

4. EMERGENCY EXIT LIGHT A. Removal (1) Ensure that the emergency light switch (1) placarded ON-TEST/OFF-RESET is OFF (Ref. Figure 201). (2) Unsnap the lens (3) from the overhead emergency exit cabin light. NOTE: It may be desirable to detach the switch (1) from the lens (3) to gain easier access to the light unit.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (3) To detach the switch (1) from the lens (3), push the switch through the mounting hole in the lens and remove the mounting plate (2) from the switch. This will allow the switch to pass through the hole in the lens. (4) Disconnect the three 1/4-turn fasteners (5) and lower the exit light unit from the headliner. (5) Disconnect the exit light electrical connector (4) from the wire harness behind the headliner. (6) Remove the exit light unit from the airplane.

B. Installation (1) Connect the exit light electrical connector (4) to the wire harness behind the headliner (Ref. Figure 201). (2) Mount the exit light unit in the headliner by connecting the three 1/4-turn fasteners (5). (3) If the exit light switch (1) has been detached from the lens (3): insert the switch through the mounting hole in the lens, install the switch mounting plate (2) on the switch, and push the switch mounting plate into the hole in the lens. (4) Snap the lens (3) into place over the exit light unit.

C. Battery Replacement (1) Perform the EMERGENCY EXIT LIGHT REMOVAL, in this section. NOTE: The proper placement of the exit light batteries (8) is illustrated on the exit light cover (6) (Ref. Figure 201). (2) Remove the screw and washer (7) attaching the exit light cover (6). (3) Remove the old batteries (8) from the battery holder and install new ones. (4) Attach the exit light cover (6) with the screw and washer (7). (5) Perform the EMERGENCY EXIT LIGHT INSTALLATION, in this section.

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Figure 201 JAR-OPS Emergency Exit Lights

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CHAPTER 34 - NAVIGATION AND PITOT/STATIC TABLE OF CONTENTS SUBJECT

PAGE

FLIGHT ENVIRONMENT DATA 34-10-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Pitot And Static Pressure System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Outside Air Temperature (OAT) Indicator . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Troubleshooting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 101 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 101 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Pitot/Static Mast . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Electrical Continuity Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202 Heat Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202 Pitot System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 203 Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 203 Static System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 204 Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 204 Pitot System Hoses . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 205 Inspecting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 205 Pitot Heads . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 205 Inspecting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 205 Outside Air Temperature (OAT) Indicator . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 212 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 212 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 212 Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 212

MAGNETIC COMPASS 34-20-00 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Calibration . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201

RADIO MAGNETIC INDICATOR SYSTEM 34-50-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201

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NAVIGATION AND PITOT/STATIC FLIGHT ENVIRONMENT DATA DESCRIPTION AND OPERATION

34-10-00 00

1. GENERAL A. Pitot And Static Pressure System The pitot and static pressure system provides a source of impact air pressure and static air for operation of the instruments (Ref. Figure 1). The pitot portion of the system utilizes the pitot/static masts, located on each side of the upper fuselage nose (Ref. Figure 2). The impact air pressure entering the masts is transmitted through separate tubing to the dual airspeed indicators mounted on the instrument panel. Since the pitot/static mast is the lowest point in each line from the instrument panel, the resultant natural drainage eliminates the need for drain valves. Two circuit breaker switches on the left subpanel control the heating elements that prevent the openings in the pitot/static masts from becoming clogged with ice which would cause the indicators to register erroneous readings. The static portion of the system includes the pitot/static masts, located on each side of the upper fuselage nose (Ref. Figure 2). Static air is routed through separate tubing to the pilot and copilot vertical speed indicator, altimeter, and airspeed indicator on the instrument panel. Copilot static air is also routed to pneumatic pressure, and cabin differential pressure gage on the subpanel. The need for drain valves is eliminated since the pitot/static mast is the lowest point in each line from the instrument panel. Should abnormal or erratic instrument readings indicate that the normal static source is restricted, an alternate static air source is provided. Alternate air static buttons are located on each side of the lower fuselage at FS 120.00. Alternate static air is routed from these static buttons through tubing to the pilot and copilot alternate static air selector valves. The alternate static air selector valve for the pilot is located on the LH side of the instrument panel, below and to the left of the airspeed indicator. The copilot's alternate static air selector valve is located on the RH side of the instrument panel, just to the right of the vertical speed indicator. When the alternate air source is required, the toggle switch is moved from the NORMAL to the ALTERNATE position on the pilot's or copilot's alternate static air selector valve. The need for drain valves is eliminated since the alternate static buttons are located in the lowest point in each line from the instrument panel.

B. Outside Air Temperature (OAT) Indicator The outside air temperature indicator is installed in the left sidewall panel in the pilot's compartment. The indicator dial is on the inside of the compartment with the stem of the instrument protruding through the skin of the airplane to the outside air. The unit is hermetically sealed against dust and moisture. A post light adjacent to the indicator face provides illumination for low light conditions and night time operation.

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Figure 1 Pitot and Static System Schematic

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Figure 2 Pitot and Static System Test Valve

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NAVIGATION AND PITOT/STATIC FLIGHT ENVIRONMENT DATA TROUBLESHOOTING

100100

1. PROCEDURES The following are some sample fault conditions and corrective actions for the pitot and static system. NOTE: Perform pitot static pressure check on system prior to service after any lines have been open. Table 101 Troubleshooting - Pitot and Static System PROBLEM 1. Heating element fails to operate properly.

PROBABLE CAUSE

CORRECTIVE ACTION

a. Switch fails to operate properly.

a. Replace the switch.

b. Grounded or open circuit.

b. Check for continuity and repair.

c. Heating element in pitot head fails to operate properly.

c. Replace the pitot/static mast.

2. Circuit breaker keeps tripping.

a. Grounded wire.

a. Check for continuity and repair.

3. Instruments inoperative or erratic in operation.

a. Lines clogged.

a. Disconnect lines at instruments and blow out with low pressure air.

b. Line leaks.

b. Check lines for loose connections at all connection points.

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1. PROCEDURES The removal and installation procedure for the left and right masts are almost identical. Both are accessed from outside of the airplane. Two index pins on the mast base plate engage with two index pin holes on the mast mounting plate to ensure approximate pitot/static mast alignment.

2. PITOT/STATIC MAST A. Removal (1) Make sure the BATT switch is set to the OFF position and attach a red tag indicating “DO NOT APPLY POWER”. (2) Perform BATTERY DISCONNECTION procedures (Ref. Chapter 24-31-00). (3) Remove the upper (6) and lower (7) half doors around the pitot/static mast (5), by removing the door screws (10 and 11). Clean any residual sealant from doors (Ref. Figure 203, Sheet 2 of 2). CAUTION: When removing the pitot/static mast (5) be very careful not to dislodge the Wet Shim (15) or the surface plate (14) from the surface of the mount assembly (8) (Ref. Figure 203, Sheet 2 of 2). If this happens, keep the pieces as intact as possible and contact Hawker Beechcraft Technical Support at 800.429.5372 or 316.676.3140. Assistance is available at any time. (4) Remove four screws (12) from the base of the pitot/static mast (5). Without breaking the wet shim (15) carefully pull the mast (5) and the gasket (13) (Ref. Figure 203, Sheet 2 of 2) away from the airplane just far enough to expose the hoses and electrical connector (4) (Ref. Figure 203, Sheet 1 of 2, Detail D). NOTE: On the right-hand pitot/static mast the aft most static line is the S2. On the left-hand pitot/ static mast the aft most static line is the S1. (5) Disconnect the pitot pressure hose (1), the S2 and S1 static pressure hoses (2 and 3) and the electrical connector (4) (Ref. Figure 203, Sheet 1 of 2). (6) Remove the gasket (13) from the base of the pitot/static mast (5) (Ref. Figure 203, Sheet 2 of 2). (7) Inspect gasket for pliability, cracks, tears or excess wear of any kind. Replace if necessary. (8) Install protective covers on the pitot/static hoses (1, 2 and 3) and the electrical connector (4) also cover the opening on the airplane fuselage (Ref. Figure 203, Sheet 1 of 2). (9) Install protective covers on the pitot/static mast (5), if it is to be put back on the airplane.

B. Installation (1) Perform PITOT/STATIC MAST ELECTRICAL CONTINUITY TEST procedures in this chapter prior to installation.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (2) Remove the protective covers from airplane fuselage, pitot/static hoses (1, 2 and 3) and electrical connector (4) (Ref. Figure 203, Sheet 1 of 2). (3) Place the gasket (13) on the base of the pitot/static mast (5) (Ref. Figure 203, Sheet 2 of 2). (4) Connect the airplane pitot/static hoses (1, 2 and 3) and electrical connector (4) to the pitot/static mast (5) (Ref. Figure 203, Sheet 1 of 2). (5) Install pitot/static mast (5) through the opening in the fuselage and secure with four screws (12) (Ref. Figure 203, Sheet 2 of 2). (6) Perform PITOT/STATIC MAST HEAT TEST procedure in this chapter. NOTE: Allow sufficient time for the Pitot/Static Mast to cool before proceeding with Step (7). (7) Temporarily install the upper (6) and lower (7) half doors. Use masking tape to mark a line where the outside contour of the door halves meet with the pitot/static mast (5). Remove the door halves and apply a bead of sealant (188, Table 1, Chapter 91-00-00) to the pitot/static mast (5) as shown (Ref. Details B and C- C Figure 203, Sheet 1 of 2). (8) Install upper (6) and lower (7) half doors, ensuring the sealant (188, Table 1, Chapter 91-00-00) on the pitot/static mast does not protrude beyond the outside edge of the doors. Secure door halves with door screws (10 and 11). Remove masking tape from pitot/static mast (Ref. Figure 203, Sheet 2 of 2). (9) Remove the red tag from BATT switch and perform BATTERY CONNECTION procedure (Ref. Chapter 24-31-00). (10) Perform PITOT SYSTEM TEST and STATIC SYSTEM TEST procedures in this chapter.

C. Electrical Continuity Test The purpose of this test is to ensure continuity of the heating element and to verify that no short exists in the circuit. NOTE: If any one of the following three Steps fail, DO NOT INSTALL this pitot/static mast (5) in the airplane (Ref. Figure 203). (1) With a suitable ohm meter check that continuity exists between pins A and B of the pitot/static mast (5). (2) Check from pin A to the case of the pitot static mast (5) to ensure that continuity DOES NOT exist. (3) Check from pin B to the case of the pitot static mast (5) to ensure that continuity DOES NOT exist.

D. Heat Test WARNING: Application of electrical power to the heating elements should not exceed ten seconds at a time. To prevent possible burns do not directly touch the heated areas. If testing cannot be completed within the ten seconds, allow the elements to cool before continuing. Ensure Pitot/Static Masts are not covered and are free of any foreign material that may burn or melt due to the extreme heat.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (1) Apply ground power to the airplane (Ref. Chapter 24-40-00). (2) Position pitot heat switches (left, right or both depending on mast being tested) to ON (Ref. Figure 204). (3) Verify that the corresponding left-hand or right-hand pitot/static mast surface gets hot (Ref. Figure 201). (4) Position the pitot/static heat switch or switches to OFF (Ref. Figure 204). (5) Remove ground power from the airplane (Ref. Chapter 24-40-00).

3. PITOT SYSTEM A. Test The purpose of these tests is to verify that no leaks exist in the entire system sufficient to induce a significant error in the airspeed or altitude indications. Note that NO instrument may be disconnected or any break made in the plumbing which would result in an untested point. (1) Cap both the pitot/static masts with Pitot/Static Adapter (4, Table 7, Chapter 91-00-00). The pitot/ static masts are located on the left and right upper side of fuselage nose. Do not cover the static ports on the pitot/static masts while performing this test. (2) Position the Pitot/Static Tester (3, Table 7, Chapter 91-00-00) in the pilot's compartment so the operator can see the instruments. (3) Connect a shop air source to the test equipment. (4) Remove the dust cap from the pilot's pitot system test port and connect the hose from the test equipment to the port. The pitot/static system test ports are located on the forward side of the bulkhead at FS 84.00 on either the pilot's or copilot's side. (5) Open the pilot's pitot test valve, close the coupler valve on the test box, then open the pressure needle valve to slowly increase pressure in the pitot system until the airspeed indicator registers 90% of the maximum reading. (6) Close the pressure valve to isolate the pitot system and observe the airspeed indicator. If the system retains the pressure for a period of five minutes, there are no leaks in the pitot lines. CAUTION: The pressure must be released SLOWLY to avoid damaging the airspeed indicator. (7) Open the pressure bleed valve and slowly reduce the pressure to zero. If a leak in the system is indicated, repair as necessary and repeat Steps (5) thru (7). (8) With the system at ambient pressure and no leaks, turn off the test port valve, and slowly apply pressure while observing the airspeed indicator. If the valves are properly closed and sealed there will be no indication of airspeed increase. (9) Disconnect the test hose from the pilot's test port and install the dust cap. (10) Repeat Steps (4) thru (9) on the copilot's pitot system.

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4. STATIC SYSTEM A. Test Upon completion of the PITOT SYSTEM TEST, proceed with the following: (1) Apply nonporous tape to the alternate static source buttons located on each side of the lower fuselage at FS 120.00. (2) Tape off the static ports on each pitot/static mast. (3) Remove the dust caps and connect the hoses from the Pitot/Static tester (3, Table 7, Chapter 91-00-00) to the pilot's static and pitot system test ports, which are located on the forward side of the bulkhead at FS 84.00 on the pilot's side. (4) Open the coupler valve of the test equipment. (5) Set the airplane altimeter to 29.92 inches Hg. NOTE: Static air toggle switches for the pilot and copilot are located on the left and right sides of the instrument panel and are placarded NORMAL and ALTERNATE. (6) Separate tests should be performed on the pilot's static system, pilot's alternate air static system, copilot's static system, and the copilot's alternate static system. While performing the pilot's and copilot's static system test, the alternate static air valve must remain in the NORMAL position. While performing the pilot's and copilot's alternate static system test, the alternate static air valve must remain in the ALTERNATE position, but only on the side being tested. (7) Slowly open the suction valve while observing the panel instruments for response. The coupler valve can be adjusted to delay the pressure change in the pitot system, thus causing the airspeed indicator to read upscale to near half scale. (8) When the required altitude reading is attained, the pitot and static system pressures will become equalized and the airspeed indicator will return to zero. (9) Make a preliminary leak check at 1,000 feet above the airplane pressure altitude to determine any loose connections. CAUTION: Excessive leakage in the static system can cause the airspeed indicator to run backwards, which will damage the indicator. (10) Close the suction valve and observe the altimeter for 30 seconds. After verifying there are no leaks, continue to increase altitude to the airplane altimeter reading indicated. (4.8 psi of vacuum) (Ref. Figure 201). (11) Close the suction valve and observe the airplane altimeter for one minute. The acceptable altitude loss may be determined from the graph in Figure 201. NOTE: The acceptable altitude loss over a one minute period is 2% of the difference between the altimeter reading before and after application of 4.8 psi of vacuum to the static system. (12) Reducing the altitude reading on the altimeter is a critical portion of the test. Use the pressure bleed valve to allow ambient air to enter the system.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL CAUTION: Never disconnect any hose or move the alternate static air selector until the altimeter has returned to field elevation (ambient Pressure). (13) Reduce the vacuum very slowly while monitoring the instruments to avoid damage. (14) Repeat Steps (3) thru (13) on the pilot's alternate static system. NOTE: In Step (6) the alternate air selector valve, on the side being tested, must be placed in the ALTERNATE position (the other side must be in the NORMAL position) and remain there. (15) Close the coupler valve. (16) With the system at ambient pressure, close the test valves, and slowly apply suction to the static system while observing the altimeter. A rise in the altimeter reading indicates a leaky valve. Very slowly reduce the vacuum to ambient. Slowly apply a positive pressure to the pitot system test valve and observe the airspeed indicator. A rise in the airspeed indicator indicates a leaky valve. If the test valves are properly closed and sealed there will be no indication on the instrument panel. (17) Repeat Steps (3) thru (15) on the copilot's system. CAUTION: At the completion of this test the alternate air selector valve must be placed in the NORMAL position for both the pilot's and copilot's system. (18) Upon satisfactory completion of these tests, remove the test hoses from the test ports and install the dust caps. (19) Disconnect the shop air from the test equipment and remove the test equipment from the pilot's compartment. (20) Remove the tape from the alternate static source buttons, located on each side of the lower fuselage at FS 120.00. Clean the static buttons. (21) Remove the covers and tape from the pitot/static masts, located on the left and right side of the fuselage nose. (22) Be certain the alternate static air selector valves are in the NORMAL position.

5. PITOT SYSTEM HOSES A. Inspecting Inspect the hoses for signs of deterioration, particularly at bends and at the connection points to the pitot/static mast and airspeed indicator. Hoses that are cracked or hardened should be replaced. NOTE: Any time a hose is replaced, the PITOT SYSTEM TEST and STATIC SYSTEM TEST procedures in this section must be accomplished.

6. PITOT HEADS A. Inspecting The pitot/static masts installed on the Model 1900 Series airplanes have a precision contour machined on their head in order to accurately measure static pressure. The surface finish of a new contour is very smooth, 32 micro inches or finer, and the color ranges from shiny metal to very dark.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Over time the high temperatures the probe heads reach, when the heater is used on the ground, will darken the entire probe heads. This darkening is a natural oxide layer protecting the heads from corrosion. In some cases over time, contaminates can break down this oxide layer, attack the base metal and cause a gradual roughening of the contour surfaces. CAUTION: No smoothing or repair to the roughened surface of the head on the probe should be attempted. Attempts to repair the roughened contour cause damage to static port edges and alter the contour. This damage will cause erroneous pressure reading. (1) Inspect the critical area of the head, i.e. the area from the pitot tip to 1/2 inch aft of the S2 static ports (the static port closest to the strut), for roughening. Replace the pitot/static mast when any of the following conditions are found: (a) Roughening in the critical area exceeds that of lightly sandblasted metal. (b) Roughening in the critical area is generally non-uniform. (c) Evidence of metal flaking in the critical area. (d) Roughening and an indication of faulty static pressure measurement, such as altitude splits between the pilot and copilot altimeters.

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Figure 201 Inspecting Pitot System Hoses and Connections

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Figure 202 Static System Test Graph

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NOTE: WET SHIM AND SURFACE PLATE MUST REMAIN ATTACHED TO MOUNT ASSEMBLY. SEE SHEET 2 OF 2

A

4 8

DETAIL

D D

5

1. PITOT PRESSURE HOSE 2. STATIC PRESSURE HOSE (S2) 3. STATIC PRESSURE HOSE (S1) 4. ELECTRICAL CONNECTOR 5. PITOT/STATIC MAST 6. UPPER HALF DOOR 7. LOWER HALF DOOR 8. MOUNT ASSEMBLY

B

1

2 3

LEFT HAND PITOT STATIC MAST SHOWN

8 DETAIL

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NOTE: ON THE RIGHT HAND PITOT STATIC MAST ITEMS 2 AND 3 ARE REVERSED

6 SEALANT BEAD PLACEMENT 6

SEALANT NOT TO PROTRUDE BEYOND OUTSIDE DOOR CONTOUR

SEALANT BEAD PLACEMENT

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C 5 7 7 DETAIL

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DETAIL

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Figure 203 (Sheet 1 of 2) Pitot/Static Mast Installation

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NOTE: DO NOT DETACH WET SHIM (15) AND SURFACE PLATE (14) FROM MOUNT ASSEMBLY (8)

8

A 9

10

11

6

15 12

14 13 5. PITOT/STATIC MAST 6. UPPER HALF DOOR 7. LOWER HALF DOOR 8. MOUNT ASSEMBLY 9. AIRPLANE SKIN 10. LARGE DOOR SCREWS (8 PLACES) 11. SMALL DOOR SCREWS (2 PLACES) 12. MAST MOUNTING SCREWS (4 PLACES) 13. GASKET 14. SURFACE PLATE 15. WET SHIM

5 7

DETAIL

A

Figure 203 (Sheet 2 of 2) Pitot/Static Mast Installation

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

A

DETAIL

A UC34B 070774AA.AI

Figure 204 Pitot/Static Mast Heat Switches

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7. OUTSIDE AIR TEMPERATURE (OAT) INDICATOR A. Removal (1) Remove the ashtray from the pilot's left sidewall escutcheon (Ref. Figure 205). (2) Reach through the opening for the ashtray and loosen the jam nut on the indicator stem. (3) Unscrew the indicator from the support assembly. Remove the nut, washer, and packing on the stem. NOTE: Be careful not to drop any of the components being removed from the stem. (4) Carefully pull the indicator through the escutcheon.

B. Installation (1) Carefully push the stem portion of the indicator through the grommet in the sidewall escutcheon (Ref. Figure 205). (2) Install the nut, washer, and packing on the stem. (3) Screw the gage down to the desired position. (4) Tighten the nut on the stem. (5) Install the ashtray in the pilot's left sidewall escutcheon.

C. Inspection Inspect the sunshield for dents, plugged openings and any misalignment that would allow contact with the stem. Inspect the nut or sunshield and stem base for stripped or damaged threads. Inspect the washers or spacers for deep scoring and distortion. Check the rubber mounting washer and packing for peeling, cracking, and resiliency. Inspect the index markings and numerals for legibility. Inspect the pointer for chipped or peeling paint. Should the indicator malfunction, it should be repaired by a certified instrument repair shop or replaced.

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Figure 205 Outside Air Temperature Indicator

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NAVIGATION AND PITOT/STATIC MAGNETIC COMPASS MAINTENANCE PRACTICES

34-20-00 200200

1. PROCEDURES A. Calibration Correction of the standby compass card for magnetic error is accomplished by removing the correction card plate and adjusting the compensator with a brass or other non-magnetic screwdriver. Calibrate the standby compass to a compass rose, or by use of a calibrated precision sight compass. Compass calibration is performed under the following conditions. (1) Start both engines. (2) Turn all normal electrical and avionics equipment ON. (3) Turn left and right generators ON. (4) Verify windshield electric heat is OFF. (5) Verify windshield wipers are OFF. (6) Turn power steering OFF. (If Installed) (7) Turn left and right pitot heat ON. (8) Turn stall warning heat ON. (9) Set the adjustment screws of the compensator on zero. Zero position of the adjustment screws is obtained by lining up the dot on the compensator frame. CAUTION: The N-S and E-W adjustment screws make approximately 3/4 turn and hit a stop. DO NOT FORCE BEYOND THE STOPS. This can cause a misalignment of the magnets and make the compass unable to be compensated. (10) Align the aircraft to a magnetic NORTH heading. Adjust the N-S adjustment screw until the compass reads exactly NORTH. (11) Align the aircraft to a magnetic EAST heading. Adjust the E-W adjustment screw until the compass reads exactly EAST. (12) Align the aircraft to a magnetic SOUTH heading. Note the resulting SOUTH error. Adjust the N-S adjustment screw until one half of this error is removed. (13) Align the aircraft to a magnetic WEST heading. Note the resulting WEST error. Adjust the E-W adjustment screw until one half of this error is removed. (14) Align the aircraft in successive magnetic 30° headings and record all errors on the outer ring of the deviation card of the compass. In each position the standby magnetic compass should read within ± 10° of the actual aircraft magnetic heading (Ref. Figure 201). NOTE: Steps (10) thru (14) may be repeated as necessary to obtain an accuracy of ± 10° at each 30° heading.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (15) Turn Left and Right Generators OFF. (16) Repeat Step (14) recording errors on the inner ring of the deviation card of the compass.

COMPASS CORRECTION CALIBRATE WITH RADIO ON

UC34B 023676AA.AI

Figure 201 Magnetic Compass Correction Card Plate

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NAVIGATION AND PITOT/STATIC RADIO MAGNETIC INDICATOR SYSTEM DESCRIPTION AND OPERATION

34-50-00 00

1. GENERAL UA-1 and After are equipped with two identical King KI-229 radio magnetic indicators (Ref. Figure 1). UB-1 and After and UC-1 and After may be equipped with either two identical King KI-229 or two identical Collins RMI-30 radio magnetic indicators, one on the pilot’s instrument panel and one on the copilot’s instrument panel. The RMI is an internally lighted, panel-mounted, dual-needle indicator that displays the airplane gyro-stabilized magnetic-heading information to either VOR or ADF stations. Each pointer is read against the compass card with a fixed lubber (index) line. Each pointer operates independently and can be switched to separate ADF or VOR receivers. A compass flag is used to show loss of power, compass invalid condition and servo error. The RMI’s derive their power from the airplane 28-VDC and 26-VAC electrical power systems. Each indicator is protected by circuit breakers. The pilot’s and copilot’s circuit breakers are located in the right sidewall circuit breaker panel. With the RMI operational, the pilot has the option of displaying heading information from NAV System No. 1, NAV System No. 2, or the ADF System. A choice of ADF or VOR heading information on the single needle is made by placing the single needle pointer button on the front of the indicator in the ADF or VOR position and selecting the desired input information on the RMI VOR/ADF remote switch next to the indicator. Compass information is displayed by the servoed compass card. The pointers will be stored in a horizontal position with VOR selected when the NAV signal is invalid, or when the selected NAV receiver is turned OFF.

Figure 1 Radio Magnetic Indicator

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NAVIGATION AND PITOT/STATIC RADIO MAGNETIC INDICATOR MAINTENANCE PRACTICES

200200

1. PROCEDURES A. Removal (1) Turn the BATTERY switch OFF and disengage the appropriate circuit breakers. Attach a caution tag to the circuit breaker panel indicating maintenance is in progress. (2) Remove the four screws that attach the indicator to the instrument panel. (3) Gently remove the indicator from the instrument panel. (4) Disconnect the electrical connector from the indicator.

B. Installation (1) Turn the BATTERY switch OFF and disengage the appropriate circuit breakers. Attach a caution tag to the circuit breaker panel indicating maintenance is in progress. (2) Connect the electrical connector on the rear of the indicator and secure with the screws attached to the connector. (3) Position the indicator in the instrument panel and attach with four attach screws. (4) Restore electrical power to the system and check for proper operation.

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CHAPTER 35 - OXYGEN TABLE OF CONTENTS SUBJECT

PAGE

OXYGEN SYSTEM 35-00-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Crew Oxygen Masks and Smoke Goggles (All Airplanes with Kit No. 118-5000 Installed) . . . . . . . . . . . 2 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Oxygen Cylinder Positive Pressure Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202 Oxygen System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202 Servicing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202 Purging . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 203 Plumbing Maintenance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 204 Oxygen System Test Procedure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 204 Oxygen System Low Pressure Test - Crew and Auxiliary Section . . . . . . . . . . . . . . . . . . . . . . . . . . . . 204 Oxygen System Low Pressure Test - Cabin Section . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 207 Manual Opening Passenger Mask Container Door . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 209 Crew oxygen mask and container . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 209 Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 209 Passenger Oxygen Mask And Container . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 210 Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 210 Oxygen Mask . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 210 Cleaning . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 210 Oxygen Mask And Container . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 210 Disinfecting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 210 Passenger Oxygen Mask . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 211 Packing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 211 Puritan - Bennett Passenger Mask Packing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 211 AVOX/Scott Passenger Oxygen Mask Packing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 211 Container Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 212 Container Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 212

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List of Effective Pages CH-SE-SU

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1 thru 6 201 thru 213

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OXYGEN OXYGEN SYSTEM DESCRIPTION AND OPERATION

35-00-00 00

1. GENERAL The oxygen system is designed to provide an adequate flow of oxygen for an airplane pressure altitude of 25,000 feet. The standard version of the airplane incorporates two oxygen cylinders with a capacity of 76.6 liters each, mounted one on either side, under the floor of the nose baggage compartment (Ref. Figure 1). The oxygen cylinders are equipped with pressure regulator/shutoff valves. The regulators are set to deliver oxygen at a constant regulated pressure. Two cylinder pressure gages are mounted on the copilot's right subpanel. A gage indicating pressure to the cabin masks is mounted at the right side of the forward pressure bulkhead (Ref. Figure 3). The optional executive liner version of the airplane is equipped with a single 22 cubic-foot oxygen cylinder which is mounted beneath the left nose baggage compartment floor. A cylinder-mounted regulator provides a reduced constant output pressure to the four phase dilution type crew and auxiliary masks. A cylinder pressure gage is mounted on the copilot's right subpanel. Earlier airplanes and executive liner versions of the airplane equipped with the 22 cubic-foot oxygen system provide two supplemental oxygen masks mounted in the ceiling aft of the forward cabin entrance door. Access to the masks is achieved by manually opening the compartment cover and then removing the mask. Oxygen flow to these masks is controlled by the removal and insertion of a lanyard pin located on the masks. The oxygen system is manually activated by the pilot using push-pull cable controls. The control on the top left side of the pilot's subpanel (OXYGEN PULL ON) actuates both oxygen cylinder valves to pressurize the crew system and the plumbing up to the passenger system control valve (CABIN PULL ON). The passenger control valve is located on the lower left side of the pilot's subpanel. To provide oxygen to the passenger oxygen masks, both controls must be pulled on. The oxygen system for the crew utilizes cylinder-mounted regulators of the constant flow type with a reduced controlled output pressure. The regulator provides for a flow rate to the crew masks of 3.8 liters per minute. The crew masks are located behind the overhead light control panel. Access to the crew masks is accomplished by manually opening the mask compartment cover and removing the mask. Oxygen flow to these masks is controlled by removing and inserting a lanyard pin on the mask. The passenger oxygen system on the standard version of the airplane is equipped with an altitude compensated, constant flow regulator mounted on the left side of the airplane just aft of the forward pressure bulkhead. This regulator varies the flow rate to the oxygen masks from a minimum of 0.1 liters per minute per mask at 1,000 feet to a maximum of 3.8 liters per minute per mask at 25,000 feet. The 19 passenger oxygen masks are located on the outboard sidewalls of the airplane close to the elbow of a seated passenger. When the CABIN OXYGEN control knob is pulled on, a surge valve momentarily allows high pressure to reach the passenger mask container assemblies. The high pressure extends a plunger which allows the dispenser door to open, allowing access to the oxygen mask. The surge valve then closes, reducing the high pressure in the line. The high pressure upstream of the surge valve is released by the vented shutoff valve. The altitude compensated regulator then controls pressure to the passenger oxygen masks. When the CABIN OXYGEN control knob is pushed in (passenger oxygen turned OFF), the surge valve resets. In order to allow oxygen flow from the mask, a lanyard valve pin must be pulled out. The pin must be inserted to stop the flow of oxygen.

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A. Crew Oxygen Masks and Smoke Goggles (All Airplanes with Kit No. 118-5000 Installed) The crew oxygen mask is a diluter/demand type with a mask vent valve (Ref. Figure 2). The vent valve is used in conjunction with smoke protection goggles. The mask has a diluter/demand type regulator and is equipped with a dynamic microphone. The crew oxygen mask are stowed in an overhead container. To begin oxygen flow, place the regulator manual select switch to the desired operating position. An oxygen flow indicator on the mask hose indicates CLEAR with proper flow and RED with no flow. The manually operated push-pull vent valve is used in conjunction with the regulator EMERGENCY pressure feature to divert a small flow of oxygen from the mask cavity into the smoke goggle cavity to vent smoke or fumes which may be present in the smoke goggle cavity. The smoke goggles are in a pouch stowed behind the crew members seats.

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CONTROL CABLE

LOW PRESSURE OXYGEN LINES

HI PRESSURE OXYGEN LINE LEGEND

E

ALTITUDE COMPENSATED REGULATOR

SHUTOFF VALVE (VENTED)

PRESSURE REGULATOR (WITH KIT NO. 118-5000 INSTALLED)

SHUTOFF VALVE

SURGE VALVE

HIGH PRESSURE OVERBOARD RELIEF

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D

FILL GAGE

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FILL VALVE

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D E FULL GAGE AND FILL VALVE DETAIL

CONTAINER ASSEMBLIES

A

OXYGEN INLET

LANYARD PIN

ON/OFF VALVE

B FORWARD PRESSURE BULKHEAD

CYLINDER PRESSURE GAGES

OXYGEN OUTLET GAGE PRESSURE

OXYGEN MASK OXYGEN INLET

CONTROL CABLE

HIGH PRESSURE OVERBOARD RELIEF LINE CABIN LOW PRESSURE

CABIN LOW PRESSURE

CONTROL CABLE

LANYARD PIN

OXYGEN MASK

OXYGEN BOTTLES

OXYGEN INLET

GAGE LINE

LANYARD GAGE LINE OXYGEN MASK AUXILIARY OXYGEN MASK CONTAINER ASSEMBLIES WITH INTERNAL REGULATORS (IF INSTALLED) DETAIL

D

PASSENGER MASKS (19 PLACES) DETAIL

E

CREW OXYGEN MASK CONTAINER ASSEMBLIES WITH INTERNAL REGULATORS (WITHOUT KIT NO. 118-5000 INSTALLED)

FILL LINE

REGULATOR ASSEMBLIES DETAIL

DETAIL

C

B UC35B 042387AB.AI

Figure 1 Oxygen System

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Figure 2 Crew Oxygen Mask and Smoke Goggles (All Airplanes with Kit No. 118-5000 Installed)

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Figure 3 Duel Bottle Schematic Regulated Oxygen System

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OXYGEN OXYGEN SYSTEM MAINTENANCE PRACTICES

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1. PROCEDURES A. Oxygen Cylinder Positive Pressure Check WARNING: Avoid making sparks and keep all burning cigarettes or fire away from the vicinity of the airplane. Make sure that your hands, tools, and clothing are clean, particularly with respect to oil or grease, for these contaminants will ignite upon contact with pure oxygen under pressure. As a further precaution against fire, open and close all oxygen valves slowly. The oxygen cylinders are located along both outboard sides of the nose wheel well (Ref. Figure 1, in the Description and Operation section). Access is gained by removing the nose baggage compartment floor. NOTE: If the pressure gauge on an oxygen cylinder reads zero, there may still be positive pressure in the cylinder. An oxygen system cylinder without any internal positive pressure, is empty. A slight positive pressure, would not be considered empty. If a cylinder is empty it must be removed, inspected and cleaned per SAE-ARP1176. (1) Loosen the cabin low pressure line at the oxygen cylinder regulator. (2) Disconnect the oxygen regulator control cable from the regulator. (3) Slowly open the cylinder valve. (4) If positive pressure is detected perform the following Steps: (a) Close the cylinder valve. (b) Connect the oxygen regulator control cable to the oxygen regulator. (c) Tighten the cabin low pressure line at the oxygen cylinder regulator. (d) Perform the OXYGEN SYSTEM SERVICING procedure. (5)

If positive pressure is not detected, remove the oxygen system cylinder for inspection by performing the OXYGEN CYLINDER REMOVAL procedure in this section. Inspect and clean per SAE-ARP1176.

B. Removal WARNING: Avoid making sparks and keep all burning cigarettes or fire away from the vicinity of the oxygen cylinder. Make sure that your hands, tools, and clothing are clean, particularly with respect to oil or grease, for these contaminants will ignite upon contact with pure oxygen under pressure. The oxygen cylinders are located along both outboard sides of the nose wheel well (Ref. Figure 1, in the Description and Operation section). Access is gained by removing the nose baggage compartment floor. (1) Make sure the OXYGEN control knob is pushed completely in (oxygen shut off).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (2) Disconnect all lines and the control cable from the regulator. (3) Immediately cap each open line with a clean fitting. (4) Immediately cap each open line with a clean fitting. NOTE: Observe the special handling precautions on the tag attached to the oxygen cylinder.

C. Installation (1) Position the cylinder in the mounting brackets and install the bracket clamp wing nuts. (2) Tighten and safety the wing nuts. (3) Carefully inspect the fittings on both the cylinder and the line for cleanliness and the presence of foreign matter, since such matter may contaminate the oxygen until it is unfit for breathing. (4) Connect the oxygen lines to the regulator. Refer to OXYGEN SYSTEM PLUMBING MAINTENANCE procedure in this section. (5) Connect the OXYGEN control cable to the on off lever on the regulator. (6) Test the connections for leaks. Refer to OXYGEN SYSTEM TEST PROCEDURE procedure in this section.

2. OXYGEN SYSTEM A. Servicing WARNING: Avoid making sparks and keep all burning cigarettes or fire away from the vicinity of the airplane. Make sure that the oxygen shutoff valve control (placarded OXYGEN PULL ON) located on the subpanel to the left of the copilot's seat is in the OFF position. Inspect the filler connection for cleanliness before attaching it to the filler valve. Make sure that your hands, tools, and clothing are clean, particularly of grease or oil, for these contaminants will ignite upon contact with pure oxygen under pressure. As a further precaution against fire, open and close all oxygen valves slowly. NOTE: Refer to Advisory Circular 43.13-1B for the additional servicing precautions recommended by the FAA on the oxygen systems. Determine Cause of oxygen system pressure loss, i.e. test, leak or crew write up due to oxygen use. The system should be purged any time system bottle pressure drops to zero psi. Refer to OXYGEN CYLINDER POSITIVE PRESSURE CHECK, and/or OXYGEN SYSTEM PURGING instructions in this section. (1) Access to the pressure gage and filler valve of the oxygen system is gained through an access door located on the left side of the nose section below the nose baggage compartment. WARNING: Ensure both the airplane oxygen system and the servicing equipment are properly grounded before servicing the system. (2) Ground the airplane oxygen system and the servicing equipment to a common ground in the ramp.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL WARNING: When filling the oxygen system, use only Aviator's Breathing Oxygen (28, Table 1, Chapter 91-00-00). Do not use oxygen intended for medical purposes, or industrial purposes such as welding. Such oxygen may contain excessive moisture that could freeze in the valves and lines of the oxygen system. (3) To recharge the oxygen system, remove the protective cap from the filler valve and attach the hose from an oxygen recharging cart to the filler valve. (4) To prevent overheating, fill the oxygen system slowly by adjusting the recharging rate with the pressure regulating valve on the cart. (5) At a temperature of 70°F the cylinders should be filled to 1,850 psi. This pressure may be increased an additional 3.5 psi for each degree of increase in temperature. For each degree of drop in temperature, reduce the pressure for the cylinders by 3.5 psi (Ref. Chapter 12-10-00, Table 3). (6) When the oxygen system is properly charged, disconnect the filler hose from the filler valve and replace the protective cap on the filler valve. (7) If at any time, in the process of servicing and purging the system or replacing the oxygen cylinder, it becomes necessary to disconnect a fitting, refer to OXYGEN SYSTEM PLUMBING MAINTENANCE in this section.

B. Purging Offensive odors may be removed from the oxygen system by purging. The system should also be purged any time the lines are left open. Purging is accomplished simply by connecting a recharging cart into the system and permitting oxygen to flow through the lines and outlets until any offensive odors have been carried away. The following Steps outline the procedures recommended for purging the oxygen system. WARNING: Avoid making sparks and keep all burning cigarettes or fire away from the vicinity of the airplane when the outlets are in use. Inspect the filler connection for cleanliness before attaching it to the filler valve. Make sure that your hands, tools, and clothing are clean, particularly of grease or oil stains, for these contaminants will ignite upon contact with pure oxygen. As a further precaution against fire, open and close all oxygen valves slowly during filling. When filling the oxygen system, use only Aviator's Breathing Oxygen (28, Table 1, Chapter 91-00-00). Do not use oxygen intended for medical purposes, or such industrial uses as welding. Such oxygen may contain excessive moisture that could freeze in the valves and lines of the oxygen system. Ensure both the airplane oxygen system and the servicing equipment are properly grounded before servicing the system. (1) Ground the airplane oxygen system and the servicing equipment to a common ground in the ramp. (2) Open the access panel for the filler valve, remove the protective cap, and attach the hose from an oxygen recharging cart to the filler valve. (3) Open the cabin door, then activate the system by pulling out the OXYGEN and CABIN OXYGEN control knobs.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL NOTE: The packing in the control cable may be lubricated with lubricant (46, Table 1, Chapter 91-00-00) to reduce friction during operation of the cable. (4) Activate the crew and passenger oxygen outlets by removing the lanyard pins. Activate the auxiliary outlets by turning on the valve located inside the container assembly. (5) Set the cart pressure regulator to deliver 200 psi of pressure to the system and turn the valve on slowly. (6) Allow the system to purge. (7) If any offensive odors remain, continue to purge. (8) If such odors still remain, replace the airplane’s oxygen cylinders. (9) After the system has been adequately purged, return the cylinder valve to its normal operation position and service the system.

C. Plumbing Maintenance Apply antiseize tape (168, Table 1, Chapter 91-00-00) to tapered threads of male fittings. Do not allow the tape to enter the inside of the fitting. Do not apply to female fittings. Apply a thin film of grease (169, Table 1, Chapter 91-00-00) to the straight threads of male fittings. Do not apply to the first two threads. Do not apply to the female fittings. Do not allow grease to get inside the fittings. When the oxygen system plumbing has been connected after maintenance, the new connections should be checked for leakage by applying leak detector fluid (47, Table 1, Chapter 91-00-00), to the connections and pressurized. Wipe dry immediately after testing. When connections leak, check that they are tightened to the proper torque value for that fitting. If this does not stop the leakage, disassemble the connection and check all mating surfaces for damage. Smooth rough mating surfaces if possible to provide a tight connection or install new fittings.

3. OXYGEN SYSTEM TEST PROCEDURE A. Oxygen System Low Pressure Test - Crew and Auxiliary Section (Auxiliary masks not used in airplane serials UB-23 and After, and UC-1 and After) The test equipment required to perform this test consists of: Supply Valve and Test Harness (5, Table 7, Chapter 91-00-00) and Pressure Gauge (6, Table 7, Chapter 91-00-00) (Ref. Figure 201, Sheet 2 of 2). WARNING: Avoid making sparks and keep all burning cigarettes or fire away from the vicinity of the oxygen cylinder. Make sure that your hands, tools, and clothing are clean, particularly with respect to oil or grease, for these contaminants will ignite upon contact with pure oxygen under pressure. (1) Be sure that the OXYGEN control cable is pushed in completely (oxygen shut off). Connect a jumper hose and a test gage between the low pressure fitting on the regulator and the low pressure line (Ref. Figure 201, Sheet 1 of 2). On the two bottle system, attach the jumper to the left side bottle. Disconnect the low pressure line which connects to the right side bottle and plug it. Cap the regulator assembly or deactivate the control cable which connects to the shutoff valve lever.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (2) Open the supply valve on the jumper hose and gage assembly. (3) Confirm that cabin supply control knob is pushed in (oxygen supply to cabin off). (4) Pull the oxygen supply control knob on and allow pressure in the lines to stabilize and note pressure for Step (9). (5) Crew masks do not auto deploy. The crew opens the lids to oxygen masks manually. (6) Pull lanyard pins to initiate flow and check flow indicators and masks for proper operation at this time. (7) Install lanyard pins to stop flow and allow two minutes for pressure in the system to equalize, then close the supply valve on the jumper hose to trap pressure in the system. (8) Close the valve on the regulator. Push in the oxygen control cable (oxygen shut off). (9) After 15 minutes, the pressure should not have dropped more than 5 psig. (10) If the leakage is excessive, apply leak detector fluid or equivalent (47, Table 1, Chapter 91-00-00). Use the fluid sparingly on the fittings and wipe dry immediately after testing. NOTE: Refer to OXYGEN SYSTEM PLUMBING MAINTENANCE procedure in this section when connecting fittings for final assembly. (11) Make any necessary repairs and retest. (12) If the cabin low pressure test is going to be performed at this time, refer to Step (2), OXYGEN SYSTEM LOW PRESSURE TEST - CABIN SECTION procedure in this section. (13) Remove the test equipment and connect the low pressure lines to the regulators. Connect the control cable on the RH cylinder if it has been disconnected. (14) Pull out the OXYGEN PULL ON control knob. (15) Carefully apply leak detector fluid to the fittings to check for leaks. Wipe dry immediately after checking. (16) Push the OXYGEN CONTROL knob in. (17) Replace the crew oxygen masks in the containers.

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Figure 201 (Sheet 1 of 2) Oxygen System Test Equipment

Figure 201 (Sheet 2 of 2) Oxygen System Test Equipment

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B. Oxygen System Low Pressure Test - Cabin Section The test equipment required to perform this test consists of: Shutoff Valve and Test Harness (5, Table 7, Chapter 91-00-00) and Pressure Gauge (6, Table 7, Chapter 91-00-00) (Ref. Figure 201, Sheet 2 of 2). WARNING: Avoid making sparks and keep all burning cigarettes or fire away from the vicinity of the oxygen cylinder. Make sure that your hands, tools, and clothing are clean, particularly with respect to oil or grease, for these contaminants will ignite upon contact with pure oxygen under pressure. (1) Be sure that the OXYGEN control cable is pushed in completely (oxygen shut off). Connect a jumper hose and a test gage between the low pressure fitting on the regulator and the low pressure line. On the two bottle system, attach the jumper to the left side bottle. Disconnect the low pressure line which connects to the right side bottle and plug it. Cap the regulator assembly or deactivate the control cable which connects to the shutoff valve lever. (2) Disconnect the Altitude Compensator from the oxygen system by removing the left oxygen bottle plumbing line and the cabin plumbing line from the compensator. Be certain that the RH bottle is capped. Connect a jumper hose (7, Table 7, Chapter 91-00-00) from the LH bottle plumbing to the cabin plumbing and cap both fittings on the Altitude Compensator. Refer to Figure 202 for installation and location of test equipment. (3) Open the supply valve on the jumper hose and gage assembly. (4) Pull the CABIN OXYGEN cable on. (5) Pull the CABIN OXYGEN PULL ON control knob and allow pressure in the lines to stabilize. NOTE: Oxygen pressure is required to open the oxygen inlet valve when the lanyard pin is pulled. However, the relatively low oxygen pressure supplied by the altitude compensated regulator at field elevations may not be adequate to initiate flow at all masks. Masks not showing flow at field elevation may be rechecked at higher cabin altitudes (up to 10,000 feet maximum) for acceptance. (6) All cabin masks should deploy at this time. (7) Pull lanyard pins of all cabin masks to initiate flow and check flow indicators and masks for proper operation at this time. Do not pull crew masks lanyards. (8) Install all lanyard pins to stop flow and allow two minutes for pressure in the system to equalize (note pressure for Step (10)), then close the supply valve on the jumper hose to trap pressure in the system. (9) Close the valve on the regulator. Push in the OXYGEN PULL ON control knob (oxygen off). (10) After 15 minutes, the pressure should not have dropped more than 5 psig. (11) If the leakage is excessive, apply leak detector fluid or equivalent (47, Table 1, Chapter 91-00-00). Use the fluid sparingly on the fittings and wipe dry immediately after testing. NOTE: Refer to OXYGEN SYSTEM PLUMBING MAINTENANCE procedure when connecting fittings for final assembly. (12) Make any necessary repairs and retest.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (13) Remove the test equipment and connect the low pressure lines to the regulators. Connect the control cable on the RH cylinder if it has been disconnected. (14) Remove the jumper test harness and caps from the altitude compensator and connect the lines. (15) Pull the CABIN OXYGEN PULL ON control knob. (16) Carefully apply leak detector fluid to the fittings to check for leaks. Wipe dry immediately after checking. (17) Push in the OXYGEN CONTROL knob. (18) Perform the appropriate oxygen mask packing procedure. (19) Observe the oxygen gauge. Service oxygen system if required.

Figure 202 Altitude Compensator Test Equipment

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4. MANUAL OPENING PASSENGER MASK CONTAINER DOOR CAUTION: Never hold or push in on the lid while attempting to open the container: this can cause excessive force to be applied to the plunger and may result in damage to the lid retaining mechanism. To manually open the door to the passenger oxygen mask container, apply a light probing force with a small rod through the hole in the door until the door falls open (Ref. Figure 203).

Figure 203 Passenger Masks

5. CREW OXYGEN MASK AND CONTAINER A. Inspection It is recommended that the oxygen masks be deployed and inspected once a year. Check that none of the following exist: (1) The oxygen mask sticks to the container or to itself. (2) Contamination of the oxygen mask, hose or the container. (3) On constant flow masks, excessive force to remove lanyard pin. A force of over four pounds should be considered excessive. (4) On constant flow masks, proper installation of lanyard pin in valve actuator. (5) Tears, cracks or deterioration of the mask or, on constant flow mask, reservoir bag (unfold bag if necessary). (6) Hose kinking.

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(7) Proper connection of oxygen hose to oxygen outlet. (8) Microphone cable connection secure. (9) Legibility of the instructions label (usually found in the container door).

6. PASSENGER OXYGEN MASK AND CONTAINER A. Inspection Check that none of the following exist: (1) The oxygen mask sticks to the container or to itself. (2) Contamination of the oxygen mask, hose or the container. (3) Excessive force to remove lanyard pin. A force of over four pounds should be considered excessive. (4) Proper installation of lanyard pin in valve actuator. (5) Tears, cracks or deterioration of the mask or reservoir bag (unfold bag if necessary). (6) Hose kinking. (7) Legibility and presence of the donning instructions label (usually found in the container door). (8) Proper connection of oxygen hose to oxygen outlet.

7. OXYGEN MASK A. Cleaning Should the oxygen masks need cleaning, wipe the surface to be cleaned with a clean, soft, lint free cloth that has been moistened with a mild detergent and warm water solution (not to exceed 110°F or 43°C). Rinse thoroughly with clean water and allow to completely air dry. NOTE: Isopropyl alcohol (30, Table 1, Chapter 91-00-00) can also be used for cleaning as well as for disinfecting. Refer to the OXYGEN MASK AND CONTAINER DISINFECTING procedure in this section.

8. OXYGEN MASK AND CONTAINER A. Disinfecting (1) Clean the mask and container. Refer to the OXYGEN MASK CLEANING procedure in this section. (2) Disinfect the mask and container with an aqueous solution of zephiran chloride (152, Table 1, Chapter 91-00-00) or isopropyl alcohol (30, Table 1, Chapter 91-00-00). (3) Use a clean, lint free cloth moistened with a solution per Step (2). Wipe quickly and lightly over the entire area. (4) After disinfecting, thoroughly air dry the mask or container.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (5) After drying, lightly dust the outside of the face piece with neo-novacite (153, Table 1, Chapter 91-00-00). (6) Install mask in container. Refer to the PASSENGER OXYGEN MASK PACKING procedure in this section.

9. PASSENGER OXYGEN MASK A. Packing WARNING: Packing and installation of the passenger masks shall be performed by personnel familiar with the procedure and warnings presented in these instructions. Failure to properly pack and install the passenger masks can result in damage to the mask or failure of the mask to deploy properly. All procedures described in these instructions shall be performed in an area free of oil, grease, flammable solvents or other contaminates.

B. Puritan - Bennett Passenger Mask Packing Pack the mask as shown on the door of the mask container.

C. AVOX/Scott Passenger Oxygen Mask Packing (1) Perform the OXYGEN MASK AND CONTAINER INSPECTION procedure in this section. (2) Insert head strap into the mask cup (Ref. Figure 204, Detail A). (3) Fold the outside thirds of the rebreather bag over the center third (Figure 204, Detail A and B). (4) Fold half of the rebreather bag over the other half (Figure 204, Detail C). (5) Collapse the face piece (Figure 204, Detail D). (6) Fold the rebreather bag under the face piece (Figure 204, Detail E). (7) Fold face piece over folded rebreather bag (Figure 204, Detail F). (8) Coil the oxygen hose under the mask. (9) If disconnected, connect the end of the hose to the valve outlet. (10) Install the lanyard pin in the valve actuator. (11) Place mask in the container, making sure that the hose and the lanyard cord are free of obstructions and are not caught on the container door. (12) Close the door to secure the mask in place.

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Figure 204 Packing Avox/Scott Passenger Masks

D. Container Removal (1) Push the OXYGEN control knob to the OFF position. (2) Push the CABIN OXYGEN control knob to the OFF position. (3) Perform the MANUAL OPENING PASSENGER MASK CONTAINER DOOR procedure in this section. (4) Remove the four screws inside the container assembly and lower the container slightly. (5) Disconnect fitting from oxygen line. (6) If the container is being replaced remove the elbow from used container.

E. Container Installation (1) If the container was replaced refer to the OXYGEN SYSTEM PLUMBING MAINTENANCE procedure in this section. (2) Install elbow into new container.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (3) Connect the fitting of the oxygen line to the elbow. (4) Perform the MANUAL OPENING PASSENGER MASK CONTAINER DOOR procedure in this section. (5) Place container inside overhead panel and install four screws into the panel. (6) Perform the PASSENGER OXYGEN MASK PACKING procedure in this section. (7) Perform the OXYGEN SYSTEM LOW PRESSURE TEST - CABIN SECTION procedure in this section.

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CHAPTER 36 - PNEUMATIC TABLE OF CONTENTS SUBJECT

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PNEUMATIC SYSTEM 36-00-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Pneumatic Bleed Air Shutoff Valve . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Regulator/Relief Valve . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202 Adjustment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202

PNEUMATIC DISTRIBUTION SYSTEM 36-10-00 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Bleed Air Precooler Duct Repair . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201

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PNEUMATIC PNEUMATIC SYSTEM DESCRIPTION AND OPERATION

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1. GENERAL Air is bled from the third stage of each engine compressor at a flow rate sufficient to produce 18 psi of regulated pressure required to operate the bleed air warning system, surface deicer, and vacuum system. Bleed air is routed aft from each engine through a shutoff valve and check valve to a single pressure regulator/relief valve, located just to the right of the airplane centerline and forward of Fuselage Station 258.25 in line with the second cabin window (Ref. Figure 1). The regulator valve is set at approximately 18 psi of pressure and incorporates a safety valve that is set at approximately 21 psi. The relief valve is set higher than the regulator pressure as a safety feature to protect against regulator failure. CAUTION: Do not exceed the recommended 18 ± 1 psi regulator setting because higher settings will result in excessive wear on the relief valve. From the pressure regulator/relief valve, lines are routed to the various airplane systems that utilize pneumatic pressure (Ref. Figure 2). The pneumatic system also incorporates a solenoid-operated shutoff valve that is energized when the bleed air control switch located on the copilot's subpanel is placed to the OFF position. This shutoff valve closes off engine bleed air at the nacelle firewall on UA and UB airplanes and at the inboard leading edge on UC airplanes and airplanes modified by Kit 114-9020-1 in the event of bleed air line breakage downstream of the valve itself. The valve is “fail/safe” in that it opens upon a loss of electrical power.

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Figure 1 Pressure Regulator

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Figure 2 Pneumatic System Diagram

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PNEUMATIC PNEUMATIC SYSTEM MAINTENANCE PRACTICES

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1. PNEUMATIC BLEED AIR SHUTOFF VALVE A. Removal The pneumatic bleed air shutoff valves are located as specified in the preceding paragraph. The appropriate panels must be removed to gain access to the valves. (1) Remove electrical power from the airplane. (2) Remove the appropriate panels. (3) Disconnect the electrical connector from the valve (Ref. Figure 201). (4) Cut safety wire on outboard B-nut. Disconnect the lines from both sides of the valve and remove the valve from the airplane. (5) Install plugs in the open lines to prevent the entrance of contaminants.

B. Installation (1) Remove electrical power from the airplane. (2) Remove the plugs from the lines (Ref. Figure 201). (3) Position the valve in the airplane with the flow arrow pointing inboard and connect the bleed air lines. The valve must be installed as near vertical as possible to promote drainage. Make sure that plumbing, electrical wiring or control cables are not directly beneath the valve vent as hot air from the vent could possibly cause damage. (4) Safety wire the outboard B-nut. (5) Connect the electrical connector to the valve. (6) Cover the valve body with insulation. DO NOT install insulation over the vent or around the solenoid of the valve. (7) Install the panels removed earlier. (8) Restore electrical power to the airplane.

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2. REGULATOR/RELIEF VALVE A. Adjustment The unit incorporates a safety valve that is set at approximately 21 psi. Ordinarily, the regulator/relief valve will not require adjustment since each unit is preset at the factory; however, the pressure regulator valve is equipped with an adjusting screw should it become necessary to reset the valve in the field (Ref. Figure 202). Adjustment may be accomplished as follows: (1) Run up the engine until a reading between 70% to 80% is registered by the gas generator tachometer. (2) Check the deicer pressure gage for a reading of 18 ± 1 psi. If necessary, unsafety the lock retaining washer and loosen the adjusting-screw locknut located on top of the valve. (3) Turn the adjusting screw clockwise to increase and counterclockwise to decrease the pressure setting until the pressure gage indicates 18 ± 1 psi. CAUTION: Do not exceed the recommended setting of 18 ± 1 psi because higher settings will result in excessive wear on the relief valve. (4) Tighten and safety the locknut, then shut down the engine. NOTE: Check the system suction gage for a reading of 4.9 to 5.9 in. Hg. Adjustment of the pressure regulator valve may affect system suction enough to necessitate readjusting the vacuum regulator valve. Ascertain that the air filter for the vacuum-operated instruments is clean and free of lint or foreign material prior to making any adjustment of the vacuum regulator valve (Ref. Chapter 37-00-00, VACUUM REGULATOR ADJUSTMENT).

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Figure 201 Pneumatic Bleed Air Shutoff Valve

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Figure 202 Pressure Regulator

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PNEUMATIC PNEUMATIC DISTRIBUTION SYSTEM MAINTENANCE PRACTICES

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1. BLEED AIR PRECOOLER DUCT A. Repair This is a general purpose procedure for the repair of the bleed air ducts attached to the bleed air precooler. Refer to Figure 201 for the left engine or to Figure 202 for the right engine for an illustration of the bleed air ducts (shaded items) covered by this procedure. (1) Remove the duct requiring repair from the airplane (Ref. Figure 201 or Figure 202). (2) Remove from the duct any material or parts that may be affected by the cleaning and welding of the duct. (3) Clean the exterior of the duct with hot, soapy water. Remove soot from the interior of the duct with hot, soapy water and a bristle brush. Rinse the duct with clean water. Dry the duct using clean shop air. (4) Clean the exterior of the duct in the area of the crack with a clean cloth soaked with acetone or alcohol. Allow the area to air dry. CAUTION: This welding repairs in this procedure should be performed only by personnel familiar with using an argon backing purge while welding stainless steel. (5) If desired, a patch may be welded over the repaired area. Grind the repair weld down before applying the patch. (6) It is essential that the edges of the crack and the surrounding area be kept clean during the welding process. Purge the interior of the duct with argon gas while welding the crack. Use AWS A5.9 Class ER347 weld rod. Tube material is 0.020 wall, 321 CRES. (7) Pressure test the repaired duct by applying a pressure source of 225 psi. There may be no loss of pressure during a one minute period. (8) Install the duct in the airplane.

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Figure 201 Bleed Air Precooler Duct Repair (Left Engine)

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Figure 202 Bleed Air Precooler Duct Repair (Right Engine)

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CHAPTER 37 - VACUUM TABLE OF CONTENTS SUBJECT

PAGE

VACUUM SYSTEM 37-00-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Troubleshooting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 101 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 101 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Instrument Air Filter Servicing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Vacuum Regulator Valve Filter Servicing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Vacuum Regulator Adjustment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201

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List of Effective Pages CH-SE-SU

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1 101 201 thru 203

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VACUUM VACUUM SYSTEM DESCRIPTION AND OPERATION

37-00-00 00

1. GENERAL Bleed air from the engines is routed through the venturi of an ejector to provide the vacuum necessary for operation of the instruments and deflation of the deicer boots. The vacuum is regulated by a suction relief valve designed to admit into the system the amount of air required to maintain sufficient vacuum (4.3 to 5.9 in. Hg) for proper operation of the vacuum operated instruments. As part of the vacuum system, the turn and slip indicator operates on a reduced vacuum setting. The vacuum ports of the vacuum operated instruments are plumbed to a vacuum manifold which is located to the right of the airplane centerline and aft of forward pressure bulkhead at Fuselage Station 97.00. The instrument air inlet ports are plumbed to an air intake manifold, located forward and inboard from the vacuum manifold between Fuselage Stations 84.00 and 94.00. Filtered air is supplied to the intake manifold through the instrument air filter. The port on the end of each manifold is plumbed to the gyro suction gage. The second port of each manifold is plumbed to the turn and slip indicator. The third port of each manifold is plumbed to the directional gyro indicator. The fourth port of each manifold is plumbed to the gyro horizon indicator (Ref. Figure 1). CAUTION: Permit no oil, grease, pipe compound, or any foreign material to enter the gyros. Be certain that all air lines are clean and free of foreign particles and/or residue before connecting lines to the gyros. Do not use thread lube on gyros with plastic housings and threads. Do not exceed 20 inch-pounds of torque on the fittings.

Figure 1 Instrument Vacuum System

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VACUUM VACUUM SYSTEM TROUBLESHOOTING

100100

1. PROCEDURES WARNING: When an airplane has experienced abnormal landing gear procedures of any type, as a safety precaution, place the airplane on jacks prior to performing any inspection or maintenance. Ensure that all three landing gears are down and locked prior to removing the airplane from jacks. CAUTION: Jacking of an airplane for the purpose of landing gear operation, inspection, servicing or maintenance, should be accomplished within an enclosed building or hangar. In the interest of safety, should it become necessary to jack the airplane in the open, wind velocity in any direction and terrain variations, must be compensated for prior to jacking the airplane. Table 101 Troubleshooting - Instrument Vacuum System PROBLEM 1. Lack of adequate vacuum to operate instruments properly.

PROBABLE CAUSE a. Clogged lines.

CORRECTIVE ACTION a. Clean or remove obstruction from lines.

b. Leaking vacuum lines. c. Dirty filters in vacuum regulator valve, pneumatic relay or cabin pressure controller. d. Dirty instrument air filter. e. Vacuum regulator valve out of adjustment. f. Vacuum regulator valve malfunctioning. g. Insufficient bleed air pressure through ejector.

b. Pressure test vacuum lines. c. Clean filters or replace.

h. Kinked (pinched) hose. 2. Vacuum drop registered on gage. a. Gage malfunctioning. b. Leakage in lines. c. Vacuum regulator valve out of adjustment. d. Vacuum regulator valve malfunctioning. e. Insufficient bleed air pressure through ejector.

3. Excessive vacuum registered on gage

f. Dirty instrument air filter. a. Vacuum regulator valve out of adjustment or malfunctioning. b. Gage malfunctioning. c. Kinked (pinched) hose.

d. Replace instrument air filter. e. Adjust vacuum regulator valve. f. Replace vacuum regulator valve. g. Check pneumatic system air pressure. Make adjustments as necessary to system components. h. Reposition or replace hose. a. Replace gage b. Pressure test vacuum lines. c. Adjust vacuum regulator. d. Replace vacuum regulator valve. e. Check pneumatic system air pressure. Make adjustments as necessary to system components. f. Replace instrument air filter. a. Check vacuum regulator valve for adjustment or malfunctioning. If vacuum regulator valve is satisfactory and problem still exist, replace gage. b. Replace gage. c. Reposition or replace hose.

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VACUUM VACUUM SYSTEM MAINTENANCE PRACTICES

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1. PROCEDURES A. Instrument Air Filter Servicing The instrument air filter is located just forward of the instrument intake manifold on the forward side of the pressure bulkhead at Fuselage Station 84.00. The instrument air filter should be checked and replaced at intervals specified in Chapter 5 or sooner if conditions warrant, such as operation in heavy smoke or dust (Ref. Figure 201).

B. Vacuum Regulator Valve Filter Servicing The vacuum regulator valve is located on the forward right side of the forward pressure bulkhead at Fuselage Station 84.00. The valve is protected by a foam-type filter mounted on top of the vacuum regulator. The filter should be cleaned with cleaning solvent (2, Table 1, Chapter 91-00-00) and blown dry with compressed air or replaced at each 200 hour interval detailed inspection. If the vacuum in the system exceeds 5.9 inches Hg, clean or replace the filter and check vacuum pressure before attempting to adjust the regulator valve (Ref. Figure 202).

C. Vacuum Regulator Adjustment The vacuum regulator valve is located on the forward right side of the forward pressure bulkhead at Fuselage Station 84.00. Adjust the valve to provide sufficient vacuum for operation of the instruments as follows (Ref. Figure 202): NOTE: Prior to making any adjustments to the vacuum regulator valve, make certain that the instrument air filter and vacuum regulator filter are clean, free of lint or other foreign materials. (1) Start one engine and run it up until a reading of 70% to 80% is registered by the gas generator tachometer. (2) Check the vacuum gage for a reading of approximately 5.9 inches Hg. If the indication is incorrect, adjust the vacuum regulator valve by removing the safety wire from the lock retaining washer and loosening the adjusting screw locknut. A calibrated vacuum gage should be used to ensure the correct setting. (3) Turn the adjusting screw IN (clockwise) to increase or OUT (counterclockwise) to decrease the vacuum setting. (4) When the vacuum gage registers a reading of 5.9 inches Hg with the engine operating within the range designated in Step (1), tighten the locknut and safety wire, then shut down the engine.

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Figure 201 Instrument Vacuum System

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Figure 202 Vacuum Regulator Valve

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CHAPTER 39 - ELECTRIC PANELS, PARTS AND INSTRUMENTS TABLE OF CONTENTS SUBJECT

PAGE

ELECTRICAL PANELS AND COMPONENTS 39-00-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1

INSTRUMENT AND CONTROL PANELS 39-10-00 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Instrument Panel (Zone 248) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Instrument Panel Components . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Instrument Subpanel (Zones 244 And 245) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Instrument Subpanel Section And Component . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Annunciator Caution/Advisory Light Panel . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Pedestal Panels And Components (Zone 243) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Upper Pedestal Panel Component . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Lower Pedestal Panel Component . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Fuel Control Panels (Zone 247) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Fuel Control Panel Section and Component . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Circuit Breaker Panel (Zone 246) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Circuit Breaker . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Overhead Panels (Zone 253) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Overhead Instrument Panel Instrument . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Overhead Instrument Panel Switch . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Overhead Light Control Panel Component . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Overhead Light Dimming Transistor . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Glareshield Annunciator Warning Light Panel (Zone 249) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

39-CONTENTS

201 201 201 202 202 202 202 202 202 202 203 203 203 204 204 204 204 205 205 205 207 207 207 207 212 212 212 212 214 214 214 214 214 214 214 215 215 215 215 215 215 218 218 218

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CHAPTER 39 - ELECTRIC PANELS, PARTS AND INSTRUMENTS TABLE OF CONTENTS (CONTINUED) SUBJECT

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ELECTRICAL AND ELECTRONIC EQUIPMENT RACKS 39-20-00 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Radar Equipment (Zone 100) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Avionic Instruments Forward Compartment (Zone 812) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Bellcrank Potentiometer (Zone 122) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Spare Fuse and Limiter Container (Zone 241) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Electrical Equipment, Forward Cabin Bulkhead Panels (Zones 220 And 221) . . . . . . . . . . . . . . . . . . . . . Electric Antiskid Control Box (Zone 121) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Antiskid Control Box . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Power Steering Amplifier Assembly (Zone 121) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Electrical Equipment, Forward Lower Cabin (Zones 133, 143 And 153) . . . . . . . . . . . . . . . . . . . . . . . . . . Electrical Equipment, Center Lower Cabin (Zone 163) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Electrical Equipment, Aft Lower Cabin (Zone 173) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Emergency Locator Transmitter (Zone 312) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Aft Autopilot Servos and Electric Elevator Trim Tab Assemblies (Zones 311 And 312) . . . . . . . . . . . . . . Autopilot Elevator Servo And Autotrim Clutch Relay Panel . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Autopilot Rudder Servo, Electric Trim Control Box and Electric Actuator . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Battery Compartment (Zone 611) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Battery Compartment Components . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Battery Compartment Fuses . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Nacelle Electrical Power Distribution Panels (Zones 521 and 621) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Inverter Panels (Zones 522 and 622) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . External Power Plug and External Power Control Circuit Breaker (Zone 740) . . . . . . . . . . . . . . . . . . . . . Stall Warning Panel and External Power Overvoltage Sensor and Advisory Light P.C. Board (Zone 522). .

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39-CONTENTS

201 201 201 201 201 201 201 202 202 202 203 203 203 203 203 204 204 204 205 205 205 205 205 205 205 205 205 206 206 206 206 206 206 207 207 207 207 207

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List of Effective Pages CH-SE-SU

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ELECTRICAL PANELS, PARTS AND INSTRUMENTS ELECTRICAL PANELS AND COMPONENTS DESCRIPTION AND OPERATION

39-00-00 00

1. GENERAL The information contained in this chapter is effective for the 1900 Series Airliners. The wiring diagrams pertinent to these airplanes are found in the following Wiring Diagram Manuals: Wiring Diagram Manual P/N 114-590032-3B for 1900 Airliner (UA Serials). Wiring Diagram Manual P/N 114-590021-13E for 1900C Airliner (UB Serials). Wiring Diagram Manual P/N 114-590021-61C for 1900C Airliner (UC Serials). This chapter is divided into two sections. Section 39-10-00 contains information necessary for the location and basic maintenance of panels containing the various indicating instruments and primary controls. Airplane system performance and total airplane system control is achieved through the use of these various panels and instruments; therefore, all the electrical panels covered in this section are located in the flight compartment and are within easy access of the flight crew. Section 39-20-00 deals with the location and maintenance of panels which contain most of the secondary controls and electrical circuitry components. Access to these panels and components is necessary only during maintenance and is gained through access openings located throughout the airplane. Panels and equipment found within this chapter can be located in the airplane by referring to the zone diagrams. Refer to Figure 1 and 2 for UA/UB/UC Serials. NOTE: Pictorial depictions in this Chapter are representative of a typical airplane and do not necessarily represent the configuration of every airplane. The appropriate wiring diagram manual should be used in conjunction with this manual.

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Figure 1 Top and LH Side View of Airplane

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Figure 2 Bottom Zone View and Flight Crew Compartment

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ELECTRICAL PANELS, PARTS AND INSTRUMENTS INSTRUMENT AND CONTROL PANELS MAINTENANCE PRACTICES 39-10-00

200200 39-10-00

CAUTION: Any time the Pitot/Static System is opened (i.e. Differential Pressure Gage or Pneumatic Pressure Indicator removed or Installed) the effected System Test procedure (PITOT SYSTEM TEST, STATIC SYSTEM TEST) must be accomplished (Ref. Chapter 34-10-00, 201).

1. INSTRUMENT PANEL (ZONE 248) The instrument panel is divided into three sections: the pilot's flight instruments (left side), the copilot's flight instruments (right side) and the avionic instruments (center section) (Ref. Figure 201). All of the instruments and avionic equipment, contained within the instrument panel, mount from the front. Sufficient excess plumbing and wiring lengths have been provided so that working behind the instrument panel is unnecessary. The UA/UB and UC serials are indicated where applicable. The instrument panel components may be removed for service or replacement by the use of the following procedure:

Figure 201 Instrument and Control Panel

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2. INSTRUMENT PANEL COMPONENTS A. Removal (1) Free the component from the front of the instrument panel by releasing or removing the fasteners on the face of the component. (2) Pull the component out of the instrument panel opening and carefully remove all plumbing and/or electrical connections from the back of the component. NOTE: Modular type avionic components have quick disconnect connectors which are released when the retaining allen screw on the face of the component is loosened.

B. Installation (1) Carefully attach the plumbing and/or electrical connectors to the component. (2) Slide the component into its opening in the panel and secure with the fasteners previously loosened or removed.

3. INSTRUMENT SUBPANEL (ZONES 244 AND 245) The instrument subpanel is located just below the instrument panel and consists of five sections: the pilot's left and right subpanels (Zone 245), the copilot's left and right subpanels (Zone 244), and the annunciator caution/advisory light panel (located between Zones 244 and 245) (Ref. Figure 202 and 203).

4. INSTRUMENT SUBPANEL SECTION AND COMPONENT A. Removal (1) For the pilot's and copilot's subpanels, remove the screws that fasten the edgelit placard panel to the subpanel section and carefully remove the placard panel. (2) Remove the screws that fasten the subpanel section to the subpanel framework and carefully lift out the section. (3) Remove all plumbing and/or electrical connections from the component. Remove the jam nut or fasteners from the front of the panel that holds the component in place and remove the component from the backside of the panel.

B. Installation (1) Position the component in the subpanel and secure with its jam nut or fasteners. Carefully connect all plumbing and/or electrical connections to the component. (2) Insert the subpanel section into the subpanel framework and secure with the previously removed screws. (3) Position the edgelit placard panel on the subpanel and secure.

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5. ANNUNCIATOR CAUTION/ADVISORY LIGHT PANEL A. Removal (1) Remove the four screws from around the face of the annunciator panel. (2) Carefully push the annunciator panel into the opening, turn horizontally and pull out of the same opening. (3) Remove the electrical connections from the annunciator panel.

B. Installation (1) Connect the electrical connections to the annunciator panel. (2) With the annunciator panel turned, slide it into its opening. Square it with the subpanel and pull forward. (3) Fasten with the four screws.

Figure 202 Pilot’s Subpanel

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Figure 203 Copilot’s Subpanel

6. PEDESTAL PANELS AND COMPONENTS (ZONE 243) The pedestal panels and components are located below the annunciator/caution advisory light panel between the pilot's and copilot's seats (Ref. Figure 204). They are sectioned into upper and lower panels. Remove them by the following procedures:

7. UPPER PEDESTAL PANEL COMPONENT A. Removal (1) Remove the screws that secure the upper panel assembly to the pedestal and lift the panel out (Ref. Figure 204). (2) Remove the screws that secure the instrument to the panel and carefully remove it from the back side of the panel. (3) Remove all plumbing and/or electrical connections from the back of the instrument.

B. Installation (1) Attach all plumbing and/or electrical connections to the back of the instrument. (2) Carefully position the instrument to the back side of the panel and secure it with its screws (Ref. Figure 204). (3) Install the panel to the pedestal using the screws previously removed to secure it. Page 204 Jan 1/17

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8. LOWER PEDESTAL PANEL COMPONENT A. Removal (1) Remove the edgelit panel by removing the screws that secure it (Ref. Figure 204). (2) Remove the panels on the sides of the pedestal by prying them out. (3) Remove the jam nut or screws from the face of the lower pedestal that secures the component to it. (4) By reaching through the side, carefully remove the component from the underside of the pedestal panel and disconnect the electrical connections from the back of the instrument.

B. Installation (1) Connect the electrical connections to the back of the instrument. (2) By reaching through the side, carefully position the component to the underside of the pedestal panel and secure with its jam nut or screws (Ref. Figure 204). (3) Place the panels into the sides of the pedestal. (4) Position the edgelit panel on the pedestal and secure with the attaching screws.

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Figure 204 Pedestal and Console Assembly

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9. FUEL CONTROL PANELS (ZONE 247) The Fuel Control panel is located in the left interior sidewall of the airplane next to the pilot. The panel is hinged at the bottom and divided into two sections (Ref. Figures 205 and 206 for UA/UB Serials) (Ref. Figure 207 and 208 for UC Serials). Component replacement is accomplished using the following procedures:

10. FUEL CONTROL PANEL SECTION AND COMPONENT A. Removal (1) Remove the appropriate edgelit panel by removing the screws that secure it. (2) Remove the three screws at the top of the fuel control panel and lean the top of the panel away from the sidewall. (3) Disconnect all of the electrical connections from the back of the component. (4) Remove the jam nut, fasteners or screws that secure the component to the panel and remove the component from the back of the panel.

B. Installation (1) Position the component to the back of the panel and secure from the front with the jam nut, fastener or screws. (2) Connect the electrical connections to the back of the component. (3) Position the fuel control panel in the sidewall and secure it with its three attaching screws. (4) Install the edgelit panel and secure with the attaching screws.

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Figure 205 Fuel Control Panel (UA-1 and After, UB-1 and After)

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Figure 206 Rear View of Fuel Control Panel (UA-1 and After, UB-1 and After)

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Figure 207 Fuel Control Panel (UC-1 and After) Page 210 Jan 1/17

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Figure 208 Rear View of Fuel Control Panel (UC-1 and After)

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11. CIRCUIT BREAKER PANEL (ZONE 246) The circuit breaker panel is located in the right interior sidewall of the airplane next to the copilot and is hinged at the bottom (Ref. Figure 209). Component replacement is accomplished using the following procedure:

12. CIRCUIT BREAKER A. Removal (1) Remove the edgelit panel by removing the screws that secure it. (2) Lean the top of the panel away from the sidewall. (3) Disconnect the wires at the circuit breaker to be removed. (4) Remove the jam nut on the front side of the panel which secures the circuit breaker to the panel and remove the circuit breaker from the back of the panel.

B. Installation (1) Insert the circuit breaker into the back side of the panel and secure from the front side with the jam nut. (2) Connect the wires to the circuit breaker. (3) Position the circuit breaker panel in the sidewall. (4) Install the edgelit panel and secure with the attaching screws.

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Figure 209 Circuit Breaker Panel (UA-1 and After, UB-1 and After)

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13. OVERHEAD PANELS (ZONE 253) The overhead panels are located in the ceiling above the pedestal controls. They are separated into three sections; the overhead instrument panel, the overhead light control panel and the two transistor panels which are exposed when the overhead light control panel is hinged open (Ref. Figures 210 and 211).

14. OVERHEAD INSTRUMENT PANEL INSTRUMENT A. Removal (1) Free the instrument from the front of the overhead instrument panel by removing the screws on the face of the instrument. (2) Pull the instrument out of the panel opening and carefully remove all electrical connections from the back of the instrument.

B. Installation (1) Carefully attach the electrical connections to the instrument. (2) Slide the instrument into the panel and secure with the screws previously removed.

15. OVERHEAD INSTRUMENT PANEL SWITCH A. Removal (1) Remove the knob from the Voltmeter Select Switch by loosening the knobs set screw and pulling it off. (2) Remove the edgelit panel by removing the six screws that secure it and lifting it off. (3) Remove the screws from the face of the overhead instrument panel instrument immediately to the left of the switch to be removed. (4) Remove the jam nut (Voltmeter Select Switch) or two screws (Emergency Light Control Switch) from the front of the panel that secures the switch and carefully pull the switch out of the back side of the panel and through the instrument opening. (5) Disconnect the electrical wires from the switch.

B. Installation (1) Connect the electrical wires to the switch. (2) Position the switch in the panel from the back side and secure with its jam nut or screws. (3) Connect the electrical connection to the back of the instrument and carefully slide the instrument into the panel. (4) Secure the instrument to the panel with the screws previously removed. (5) Position the edgelit panel to the panel and secure. (6) Position the selector knob on the Voltage Select Switch and tighten its set screw.

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16. OVERHEAD LIGHT CONTROL PANEL COMPONENT A. Removal (1) Remove all the rheostat selector knobs from the front of the panel by loosening their set screws and pulling them off. (2) Remove the edgelit panel by removing the six screws that secure it. NOTE: Remove the screws at the bottom of the panel first. The screws at the top of the panel allow the panel to drop away from the headliner. (3) Remove the wires from the back of the component. (4) Remove the jam nut or screws from the front of the panel that secures the component and then remove the component from the back of the panel.

B. Installation (1) Position the component to the back of the panel and secure from the front with the jam nut or screws. (2) Connect the wires to the component. (3) Position the edgelit panel on the front of the overhead light control panel and secure both panels to the headliner with the screws previously removed. NOTE: Install the screws at the top of the edgelit panel first. The screws at the top of the panel are used to secure the panel to the headliner. (4) Install all rheostat selector knobs on the appropriate post and secure by tightening the knob set screws.

17. OVERHEAD LIGHT DIMMING TRANSISTOR A. Removal (1) Remove the two upper corner screws from the overhead light control panel and allow the panel to hinge open. (2) Remove the two screws securing the transistor to the bracket and carefully remove the transistor from the bracket. (3) Disconnect all wiring to the transistor.

B. Installation (1) Connect all wiring to the transistor. (2) Carefully position the transistor into the bracket and secure it with its two screws. (3) Position the overhead light control panel into the headliner and secure with the two upper corner screws.

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Figure 210 Overhead Panels Page 216 Jan 1/17

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Figure 211 Overhead Light Control Panel

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18. GLARESHIELD ANNUNCIATOR WARNING LIGHT PANEL (ZONE 249) The annunciator warning light panel is located in the middle of the instrument panel glareshield and is removed using the following procedure (Ref. Figures 201):

A. Removal (1) Remove the cover plate on top of the light panel by removing the four screws that attach it to the glareshield. (2) Remove the two screws from the front panel that secure the annunciator panel to it. (3) Carefully lift the annunciator panel from the backside of the face panel and disconnect the electrical connection from the back of the annunciator panel.

B. Installation (1) Connect the electrical connection to the back of the annunciator panel and carefully position it into the front panel from the back side. (2) Secure the annunciator panel to the front panel with the two screws previously removed. (3) Install the cover plate onto the glareshield and secure with its four screws.

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ELECTRICAL PANELS, PARTS AND INSTRUMENTS ELECTRICAL AND ELECTRONIC EQUIPMENT RACKS MAINTENANCE PRACTICES

39-20-00 200200

1. PROCEDURES 2. RADAR EQUIPMENT (ZONE 100) The radar equipment is located in the forward compartment of the fuselage under the nose cone. Access to it is gained by removing the nose cone section.

3. AVIONIC INSTRUMENTS FORWARD COMPARTMENT (ZONE 812) The forward avionics instruments are found in the avionics instrument compartment. It is located in the right side of the nose section opposite the forward baggage compartment and is accessible through the hinged avionics compartment door found on the right side.

4. BELLCRANK POTENTIOMETER (ZONE 122) The bellcrank potentiometer is located under the flight compartment floor on the pilot's side and can be removed for replacement or service by the following procedure (Ref. Figures 201 and 205):

A. Removal (1) Remove flight compartment access panel (1) that extends itself between the pilot's rudder pedals (Ref. Figures 205). (2) Disconnect all electrical connections and the tie rod from the potentiometer (Ref. Figures 201). (3) Remove the screws that secure the potentiometer to the bracket and carefully remove the potentiometer.

B. Installation (1) Position the potentiometer to the bracket and secure with the previously removed screws (Ref. Figure 201). (2) Connect the tie rod and all the previously removed electrical connections to the potentiometer. (3) Place the access panel (1) over the opening and secure (Ref. Figures 205).

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Figure 201 Bellcrank Potentiometer

5. SPARE FUSE AND LIMITER CONTAINER (ZONE 241) The spare fuse and limiter container is found in the storage pouch located on the interior sidewall, behind the pilot's seat.

6. ELECTRICAL EQUIPMENT, FORWARD CABIN BULKHEAD PANELS (ZONES 220 AND 221) The two forward cabin bulkhead electrical equipment panels are mounted on the upper left and right sections of the forward cabin pressure bulkhead. The components on these panels are accessible by reaching under the instrument panel (Ref. Figures 206 and 207).

7. ELECTRIC ANTISKID CONTROL BOX (ZONE 121) The electric antiskid control box is located under the flight compartment on the copilot's side and can be removed for repair or replacement by the following procedure (Ref. Figures 202 and 205):

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8. ANTISKID CONTROL BOX A. Removal (1) Remove access panel (5) which is directly in front of the copilot's seat (Ref. Figure 205). (2) Disconnect the electrical connection from the antiskid box (Ref. Figure 202). (3) Remove the screws that secure the antiskid box to its mountings and carefully lift it out of the opening.

B. Installation (1) Place the antiskid box onto the mountings and secure with its screws (Ref. Figure 202). (2) Connect the electrical connection to the antiskid box. (3) Place the access panel (5) over the opening and secure (Ref. Figure 205).

Figure 202 Antiskid Control Box

9. POWER STEERING AMPLIFIER ASSEMBLY (ZONE 121) The power steering amplifier is located under the flight compartment floor on the copilot's side and can be removed for repair or replacement by the following procedure (Ref. Figures 203 and 205):

A. Removal (1) Remove the copilot's seat (Ref. Chapter 25-00-00, FLIGHT COMPARTMENT SEAT REMOVAL).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (2) Remove access panel (7) located directly under the copilot's seat between the seat rails (Ref. Figure 205). (3) Remove the electrical connection from the power steering amplifier box (Ref. Figure 203). (4) Remove the screws that secure the power steering amplifier and carefully lift it out of the access opening.

B. Installation (1) Place the power steering amplifier into the access opening and secure to the brackets with its screws (Ref. Figure 203). (2) Connect the electrical connection to the power steering amplifier box. (3) Place the access panel (7) over the opening and secure (Ref. Figure 205). (4) Install the copilot's INSTALLATION).

seat

(Ref.

Chapter

25-10-00,

FLIGHT

COMPARTMENT

SEAT

Figure 203 Power Steering Amplifier

10. ELECTRICAL EQUIPMENT, FORWARD LOWER CABIN (ZONES 133, 143 AND 153) The forward lower cabin electrical equipment components are located in the space below the center floor panels between the flight compartment and the main wing spar. To gain access to these components remove the applicable center floor panel (Ref. Figures 208, 209, 210 and 211).

11. ELECTRICAL EQUIPMENT, CENTER LOWER CABIN (ZONE 163) The center lower cabin electrical equipment components are located in the space below the first center floor panel aft of the main wing spar. Remove this floor panel to gain access to the components (Ref. Figures 208, 211 and 212).

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12. ELECTRICAL EQUIPMENT, AFT LOWER CABIN (ZONE 173) The aft lower cabin electrical equipment components are located in the space below the second and third center floor panel aft of the main wing spar. To gain access to these components remove the applicable center floor panel (Ref. Figures 208, 213 and 214).

13. EMERGENCY LOCATOR TRANSMITTER (ZONE 312) The emergency locator transmitter is located in the right aft section of the fuselage just forward of FS 598. Access to it is gained by removing the access panel just below the right stabilon.

14. AFT AUTOPILOT SERVOS AND ELECTRIC ELEVATOR TRIM TAB ASSEMBLIES (ZONES 311 AND 312) The elevator and rudder autopilot servos and electric elevator trim tab assemblies are located in the aft section of the fuselage between the aft pressure bulkhead and Canted Fuselage Station 605.984. Use the following procedures to remove the applicable components for repair or replacement:

15. AUTOPILOT ELEVATOR SERVO AND AUTOTRIM CLUTCH RELAY PANEL A. Removal (1) Remove the fuselage access panel located immediately below the left stabilon. (2) Reaching through the access opening, remove all electrical wires from the applicable component. (3) Remove the screws that secures the component to the panel and carefully remove the component through the access opening.

B. Installation (1) Insert the component through the left access opening, position on the panel and secure with the screws previously removed. (2) Connect all electrical wires to the component. (3) Place the access panel over the opening and secure.

16. AUTOPILOT RUDDER SERVO, ELECTRIC TRIM CONTROL BOX AND ELECTRIC ACTUATOR A. Removal (1) Remove the fuselage access panel located immediately below the right stabilon. (2) Reaching through the access opening, remove all electrical wires from the applicable component. (3) Remove the screws that secures the component on the panel and carefully remove the component through the access opening.

B. Installation (1) Insert the component through the right access opening, position on the panel and secure with the screws previously removed.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (2) Connect all electrical wires to the component. (3) Place the access panel over the opening and secure.

17. BATTERY COMPARTMENT (ZONE 611) The battery compartment is located in the right wing center section. With the exception of the fuses, use the following procedure to remove the electrical components for repair or replacement (Ref. Figure 216):

18. BATTERY COMPARTMENT COMPONENTS A. Removal (1) Remove the battery compartment access panel located on the top surface of the right wing center section. (2) Cut the safety wire and remove the battery disconnect. (3) Cut the safety wires and loosen the hold-down wing nuts. Push the wing nuts aside and remove the battery hold-down bar. (4) Lift the battery out of the battery compartment. (5) Remove all electrical wires and connections from the component to be removed. (6) Unfasten the component from the battery compartment walls or floor and then remove the component.

B. Installation (1) Position the component in the battery compartment and secure. (2) Connect the electrical wires to the component. (3) Place the battery in the battery compartment. (4) Install the hold-down bar and secure it in place with the wing nuts, then safety the wing nuts. (5) Install the battery disconnect and safety wire. (6) Place the access panel over the battery compartment and secure. Be sure that the battery cover vent opening aligns with the access panel vent and gasket.

19. BATTERY COMPARTMENT FUSES A. Removal (1) Remove the battery compartment access panel located on the top surface of the right wing center section. (2) Remove the three fasteners that secure the fuse holder cover and lift the cover off. (3) Remove the fuse to be replaced.

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B. Installation (1) Place the fuse into the fuse box. (2) Position the cover over the fuse box and secure with the three fasteners. (3) Place the access panel over the battery compartment and secure. Be sure that the battery cover vent opening aligns with the access panel vent and gasket.

20. NACELLE ELECTRICAL POWER DISTRIBUTION PANELS (ZONES 521 AND 621) The nacelle electrical power distribution panels are located in the upper section of the left and right nacelles just aft of the engine firewall. Remove the access panels located on top of these sections when needing to gain access to repair or replace any of the components found on these panels (Ref. Figures 204, 217 and 218).

21. INVERTER PANELS (ZONES 522 AND 622) The nacelle electrical power distribution panels are located in the upper section of the left and right nacelles just aft of the engine firewall. Remove the access panels located on top of these sections when needing to gain access to repair or replace any of the components found on these panels (Ref. Figures 204 and 219).

Figure 204 Upper Aft Nacelle Covers

22. EXTERNAL POWER PLUG AND EXTERNAL POWER CONTROL CIRCUIT BREAKER (ZONE 740) The external power plug and external power control circuit breaker are located in the lower aft section of the left nacelle. Access is gained by unfastening the hinged access panel found there (Ref. Figure 220).

23. STALL WARNING PANEL AND EXTERNAL POWER OVERVOLTAGE SENSOR AND ADVISORY LIGHT P.C. BOARD (ZONE 522) The stall warning panel and external power P.C. board are located just aft of the inverter panel under the upper aft cover of the left nacelle. Remove this cover to gain access to the stall warning panel and external power overvoltage sensor and advisory light P.C. board (Ref. Figures 204 and 220).

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Figure 205 Flight Compartment Access Panels

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Figure 206 Left Electrical Equipment Forward Bulkhead Panel

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Figure 207 Right Electrical Equipment Forward Bulkhead Panel

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Figure 208 Fuselage Floor Panel Layout

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Figure 209 Forward Lower Cabin Electrical Equipment (ZONE 133)

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Figure 210 Forward Lower Cabin Electrical Equipment (ZONE 143)

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Figure 211 Forward Lower Cabin Electrical Equipment Relay

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Figure 212 Center Lower Cabin Electrical Equipment

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Figure 213 Autopilot Aileron Servo Actuator

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Figure 214 Aft Lower Cabin Electrical Equipment and Yaw Damper Computer

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Figure 215 Aft Autopilot Servos and Electric Trim Control Assemblies

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Figure 216 Battery Compartment

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Figure 217 LH Nacelle Electrical Power Distribution Panel

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Figure 218 RH Nacelle Electrical Power Distribution Panel

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Figure 219 Inverter Panels

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Figure 220 External Power and Stall Warning Panels

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CHAPTER 51 - STANDARD PRACTICES AND STRUCTURES TABLE OF CONTENTS SUBJECT

PAGE

STRUCTURES 51-00-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1

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STANDARD PRACTICES AND STRUCTURES STRUCTURES DESCRIPTION AND OPERATION

51-00-00 00

1. GENERAL The Model 1900 Airliner is semimonocoque in construction and is pressurized between bulkheads 84.00 and 557.50. When making repairs or modifications that create a break in the pressure vessel, the mating surface must be sealed with the proper sealer described in Table 2, Chapter 91. To ensure effective bonding of the sealers, clean thoroughly all mating surfaces, mating parts and rubber seals. Information pertaining to the cabin airstair doors, cargo door, baggage compartment door and emergency exits is included in Chapter 52. Maintenance practices for the horizontal stabilizer and vertical stabilizer are included in Chapter 55. Information pertaining to the windshield and the cabin windows is included in Chapter 56. Information pertaining to the wing sections and the one piece spar is included in Chapter 57. If repair to a structural member is necessary, refer to the MODEL 1900 SERIES AIRLINER STRUCTURAL REPAIR MANUAL P/N 114-590021-9.

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CHAPTER 52 - DOORS TABLE OF CONTENTS SUBJECT

PAGE

AIRSTAIR DOOR 52-10-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Airstair Door Damper (Hydraulic Snubber) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Service . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Airstair Door . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Mounting and Fitting (New) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202 Airstair Door Latching Mechanism . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 204 Adjustment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 204 Airstair Door Forward and Aft Handrail . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 212 Upper Forward And Upper Aft Handrail Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 212 Upper Forward And Upper Aft Handrail Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 212 Lower Forward And Lower Aft Handrail Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 212 Lower Forward And Lower Aft Handrail Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 213 Airstair Door Forward and Aft post . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 213 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 213 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 214 Airstair Door Handrail . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 215 Functional Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 215 Airstair Door Damper . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 217 Replacement . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 217 Airstair Door Seal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 217 Removal and Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 217 Repair . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .218

EMERGENCY EXIT 52-20-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Emergency Exit Door . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Latching Mechanism Adjustment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201

EMERGENCY EXIT LATCH MECHANISM 52-20-01 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Latch Mechanism . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

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CARGO/NOSE BAGGAGE COMPARTMENT DOORS 52-30-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Cargo Door . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Nose Baggage Compartment Door . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Cargo Door . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Prerigging . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 204 Cargo Door Latching Mechanism . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 204 Adjustment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 204 Gas Spring Assembly . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 206 Replacement . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 206 Cargo Door Seal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 206 Replacement . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 206 Nose Baggage Compartment Door . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 207 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 207 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 207 Seal Replacement . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 207

CARGO AND AIRSTAIR DOOR WARNING 52-70-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Cargo and Airstair Door Warning Switches . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Airstair Door Doorsill Switch . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Adjustment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Airstair Door Forward and Aft Latch Switches . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Adjustment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Airstair Door Latch Handle Switch . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202 Adjustment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202 Cargo Door Doorsill Switch . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 204 Adjustment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 204 Cargo Door Latch Handle Switch . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 204 Adjustment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 204 Cargo Door Latch Pin Switch . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 204 Adjustment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 204 Cabin/cargo “FWD Cabin Door/AFT Cabin Door” Annunciator Circuitry Check . . . . . . . . . . . . . . . . . . . . . 207 Cabin Door . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 207 Cargo Door (If Installed) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 207 Annunciator Circuitry . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 207

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DOORS AIRSTAIR DOOR DESCRIPTION AND OPERATION

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1. GENERAL A swing down door, hinged at the bottom, provides cabin security for flight and a convenient stairway for entry and exit. A plastic encased cable provides support for the door in the open position, a hand hold for passengers, and a convenience for closing the door from the inside. A rubber seal around the cabin door positively seals the pressure vessel while the airplane is in flight. A hydraulic damper permits the door to lower gradually during opening (Ref. Figure 1). The latching mechanism for the airstair door is operated by two handles interconnected inside the door. The rotary handle on each side of the door is attached to cable drums. When the handle is rotated, cables attached to the drums wind or unwind in response to the direction of handle movement. The cables from the handle are routed around splined drums at each of the camlock latches on the forward and aft sides of the door. The splines in the drum engage splines on the camlock shafts of the latch so that rotation of the drum, when actuated by the cable from the handle, rotates the face of the latch to engage the latch post in the frame of the door. The latch cables are adjusted by turnbuckles to a tension sufficient to ensure proper latching and unlatching of the door in response to handle movement. A splined cam on the ends of the camlock shaft opposite the latch face actuates switches mounted on brackets at the lower latches to indicate when the door is closed and latched. Whether unlocking the door from the outside or inside, the release button adjacent to the door handle must be held depressed before the handle can be rotated to unlock the door. The release button acts as a safety device to help prevent accidental opening of the door. As an additional safety measure, a differential pressure sensitive diaphragm is incorporated into the release button mechanism. The outboard side of the diaphragm is open to atmospheric pressure, the inboard side to cabin air pressure. As the cabin to atmospheric pressure increases, it becomes increasingly difficult to press the release button, because the diaphragm moves inboard when either the inside or outside release button is pressed.

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Figure 1 Airstair Door Damper

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DOORS AIRSTAIR DOOR MAINTENANCE PRACTICES

200200

1. PROCEDURES 2. AIRSTAIR DOOR DAMPER (HYDRAULIC SNUBBER) A. Service (1) Attach the barrel rod end to the cabin door supports and the movable piston rod to the supports on the door frame when installing the hydraulic snubber (Ref. Figure 201). Shim as necessary with AN960-14 washers to eliminate side play between the rod ends and their supports. (2) Service the snubber with hydraulic fluid (39, Table 1, Chapter 91-00-00). To service, loosen the cap on the end of the cylinder barrel with the movable rod by removing the three set screws securing the cap to the barrel. Push down on the rod end until the piston is bottomed in the cylinder barrel. Slide the cap up on the rod end and fill the barrel with 215 cubic centimeters (7.2 ounces) hydraulic fluid. Secure the cap to the cylinder barrel with the three set screws.

3. AIRSTAIR DOOR A. Removal (1) Disconnect all electrical power. (2) Disconnect the electrical connector adjacent to the door hinge. (3) Support the door and remove the screw, washer, and bushing attaching the handrail cable eye to the fuselage mount assembly. (4) Remove the bolt and washers attaching the door damper to the airstair door support bracket. (5) Remove the threshold plate. (6) Remove the screw and hinge retainer plate at each end of the hinge pin. (7) Support the door and remove the hinge pin. Carefully remove the door.

B. Installation (1) If necessary, lubricate the airstair door (Ref. Chapter 12-20-00, CABIN DOOR LUBRICATION). (2) Carefully move the door into position and support the door. Mate the hinge halves, and install the hinge pin. (3) Install the hinge pin retainers on each end of the hinge. (4) Install the bushing, spacer, and screw attaching the handrail cable eye to the fuselage mount assembly. (5) Install the bolt and washer attaching the door damper to the airstair door support bracket. (6) Connect the electrical connector adjacent to the door hinge.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (7) Connect electrical power to the airplane and engage the cabin door circuit breaker. (8) Close and lock the airstair door. With electrical power ON, observe that the FWD CABIN DOOR light is out. (9) Unlock the door and check the time required for the door to free-fall from full-closed to full-open position. The time required shall not be less than five seconds. CAUTION: Check that the nut on the bolt securing the base of the door cable post is installed toward the center of the door to prevent the handrail from hanging up under the bolt.

Figure 201 Airstair Door Damper

C. Mounting and Fitting (New) A new airstair door is supplied with the door hinge half, the seal retainer, and the seal separate from the door. Do not rivet the hinge half or seal retainer to the new door until after the door has been fitted. NOTE: To determine if the door hinge requires replacement, a minimum wall thickness of 0.109 inch is permissible. If only the hinge is being replaced, the factory installed hinge used NAS1739B5 rivets in 0.176 to 0.180 inch diameter holes. If an oversized or elongated hole is encountered, use NAS9303B6 rivets in 0.192 to 0.196 inch diameter holes. If additional oversized (larger than above) rivets are required on subsequent replacement, use NAS9305B6 or NAS1739B6 rivets in 0.205 to 0.209 inch diameter holes. Countersink the holes to 105° X 0.353 inch diameter. When measuring the diameter of the hinge pin, if the diameter is less than 0.126 inch replace the pin. (1) Remove the old door from the airplane. Perform the AIRSTAIR DOOR REMOVAL procedure.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (2) Place a shim, approximately 0.038 to 0.050 inch thick and 8 to 10 inches long, into the recessed airstair door frame on the fuselage. NOTE: A shim will hold the skin of the new airstair door flush with the outside skin of the fuselage once the airstair door is properly trimmed. (3) Position the airstair door in the fuselage frame. The excess skin on the door will not allow it to fit into the recessed frame; however, center the camlock latches on the posts as loosely as possible. (4) Place the handle of the airstair door to the fully open position. This will allow the six camlock latches to align with the posts in the fuselage frame. CAUTION: Extreme caution must be used during skin removal to prevent removing excess material and creating an enlarged gap between the fuselage and the airstair doors. The gap between the airstair door and the fuselage is to be a maximum of 0.060 inch. (5) Starting at the bottom of the airstair door, determine where the excess skin needs to be removed to allow the bottom of the door to fit into the recessed frame. Remove the airstair door from the fuselage frame. Using a 40 grit sanding disc attached to an electric motor, remove excess skin from the bottom of the door as required. (6) Position the airstair door in the fuselage frame and center the forward and aft camlock latches on the post. Determine where excess skin needs to be removed from the sides of the door so it will fit into the recessed frame. Remove excess skin from the sides as required. NOTE: Repeat the above procedure as many times as required until the sides of the door fit into the recessed frame. (7) As the door begins to fit into the opening, place four 3/8 to 1/2 inch shims, two on each side, into the recessed opening to support the door flush with the outside skin of the fuselage. (8) Mark and begin removing excess skin material from the top of the door. This is best accomplished by leaving the door in the opening and raising the top up for sanding. Repeat this process, as required, until the top of the door will fit into the recessed opening. Place three 3/8 to 1/2 inch shims into the recessed opening to support the top of the door flush with the outside skin of the fuselage. (9) Once the door fits completely into the opening, check that the camlock latches are centered on the post and latch the door closed. Mate the new hinge to the existing hinge half on the fuselage frame and install the hinge pin. Use shims as required all around the door to keep it from moving while drilling hinge holes. (10) Drill twenty-three 0.128 to 0.133 inch diameter holes through the door, using the first row of pilot holes adjacent to the hinge pin as guides. Countersink the holes 100° X 0.230 inch diameter on the hinge half. (11) Install #30-pull cleco fasteners (minimum of seven) equally spaced to hold the door to the hinge half. Drill thirty-five 0.176 to 0.180 inch diameter holes, using the remainder of the holes in the hinge half as guides. Countersink the holes 100° X 0.286 inch diameter on the hinge half. CAUTION: Do not exceed the maximum gap of 0.060 inch between the door and the fuselage frame.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (12) The door can now be opened and closed freely on the hinge. Remove the remaining excess skin material until the desired gap of 0.040 inch is obtained between the airstair door and the fuselage. (13) Remove the hinge pin and the cleco fasteners. Remove the airstair door from the fuselage frame. Apply sealer (38, Table 1, Chapter 91-00-00) between the hinge half and the door frame. (14) Position the new hinge half and the lower seal retainer on the door and align the pilot holes in the retainer with the twenty-three holes drilled in Step (10). Seal between the lower seal retainer and the door with sealer. (15) Install twenty-three rivets through the hinge half, door skin, and the retainer. Install thirty-five rivets through the remaining holes in the hinge half. (16) Position the two side retainers and the top retainer along the inside edge of the door. The outside edge of the retainers should be 0.025 inch from the outside edge of the door. Drill 0.096 inch diameter holes through the door, using the holes in the retainers as guides. Countersink the holes 100° X 0.230 inch in diameter on the outside skin. (17) Seal between the retainers and the door with sealer. Install rivets to secure the retainers to the door. (18) Carefully place the airstair door into position, mate the hinge halves, and install the hinge pin. Install the hinge pin retainers on each end of the hinge with screws. (19) Install the bushing, spacer, washer, and screw attaching the handrail cable eye to the fuselage mount bracket. (20) Perform the ADJUSTMENT OF AIRSTAIR DOOR LATCHING MECHANISM procedure. (21) Connect the electrical connector for the airstair door warning switches located under the doorsill step adjacent to the door hinge. Perform the ADJUSTMENT OF CARGO AND AIRSTAIR DOOR WARNING SWITCHES procedure in Chapter 52-70-00. (22) Apply soapstone powder to the new seal and retainer. Starting with the seal at the marked location on the door, work the seal into the retainer. A thin flat strip of plexiglas or phenolic with a rounded end will facilitate installation. Clean off the soapstone powder as necessary. NOTE: To ensure retention of the seal in the track, the seal may be cemented in place with adhesive (43, 145, 167 or 193, Table 1, Chapter 91-00-00). (23) Remove upholstery from the old door and install on the new door, or install new upholstery as required. (24) Perform the AIRSTAIR DOOR INSTALLATION procedure. (25) After the door has been installed and painted to match the fuselage, lubricate the door (Ref. Chapter 12-20-00, LUBRICATION SCHEDULE).

4. AIRSTAIR DOOR LATCHING MECHANISM A. Adjustment (1) Remove the upholstery panels as necessary to gain access to the latching mechanism. (2) Remove the seal from its retainer around the periphery of the airstair door. Page 204 May 1/10

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (3) Shim the airstair door with spacer blocks taped to the door at the six latch points so that the outside contour of the door matches that of the fuselage. NOTE: An undercontour of 0.06 inch between the airstair door and the fuselage is acceptable, but the airstair door contour must NEVER be outside that of the fuselage. (4) With the outside handles in the closed position and the pressure lock engaged by the arm assembly, insert a rig pin through the handle support into the cable drum, then adjust the handle rod to fit the arm of the outside handle as shown (Ref. Figure 202, Detail B). NOTE: Only the rod end that attaches to the inside handle is adjustable. After the handle rod is adjusted to the proper length, check that the threads of the adjustable end are visible in the inspection hole of the rod before tightening the jam nut and attaching the rod to the handle arm. If it is necessary to replace the door pressure lock or the handle actuator assembly, the 0.015 to 0.045 inch dimension shown must be maintained (Ref. Figure 208). (5) Remove the rig pin from the support of the inside arm. (6) Remove upholstery panels and Step(s) on the airstair door as necessary to inspect the latch mechanism cable. Check the latch mechanism cables for clearances. NOTE: If a door cable requires replacement, wrap the cable around each latch drum and the handle drum so that the retaining pin slot in the drum is aligned with the drum guide retaining screw on the inboard side of the latch housing (Ref. Figure 203). This will ensure adequate rotation of the camlock shaft to complete the latching and unlatching movements of the latching mechanism. At each change of season, or if difficulty is experienced with airstair door operation, it is recommended that the bearings on the cam and drum assemblies be lubricated with lubricant (111, Table 1, Chapter 91-00-00) or equivalent. (7) Check latch mechanism cables for clearances, damage, broken strands, corrosion, condition and security of attachment. Move door handle from open to closed to check entire length of cables. Replace cables with broken strands. (8) Rig the latch cables to the tension designated by the graph (Ref. Figure 207). Cable tension for the camlock latches is adjusted with the turnbuckles on each side of the door. (9) After the cable tension has been set, safety the turnbuckle as indicated (Ref. Figure 204). (10) With the door handles in the open position, insert the splines of each camlock shaft into the splines of its respective cable drum so that the cutout in the latch face of the shaft is centered with respect to the arc made by the latch when the door is closing as indicated (Ref. Figure 205). (11) Engage the splines of each latch cam with the splines on the small end of the camlock shaft so that the cam just contacts its stop pin (Ref. Figure 203). On the latch on each side of the door where the electrical switches are located, back the cam off one spline from the point where it contacts the stop pin, then install the washer and nut securing it in place. On the other four latches, back the cam off two splines from the point where it contacts the stop pin, then install the washer and nut securing the cam in place.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL NOTE: Any time the camlock shafts are repositioned with respect to the cable drums, the latch cams must be reindexed to the camlock shaft to ensure full travel (90°) of the latch handles. (12) Adjust each latch post on the fuselage frame adjacent to the latches of the airstair door until the post's point of contact with the latch face of the camlock shaft aligns with the closing arc made by the latch when the door is fully closed (Ref. Figure 206). To make this adjustment, rotate the eccentric bushing, in conjunction with the eccentric latch post, until the post is properly aligned with the latch, then tighten and safety the nut securing the post in place. A special latch post wrench (P/N 101-590052-1) is required to adjust the eccentric bushing. (13) Perform the ADJUSTMENT OF CARGO AND AIRSTAIR DOOR WARNING SWITCHES procedure in Chapter 52-70-00. NOTE: Coating the latch post with Blue Dykem, or its equivalent, will facilitate pinpointing its area of contact with the face of the latch. (14) Remove the spacer blocks taped to the door at the six latch points. (15) Work the door seal into the retainer channel around the periphery of the door. NOTE: To ensure retention of the seal in the track, the seal may be cemented in place with adhesive (43, 145, 167 or 193, Table 1, Chapter 91-00-00). (16) Attach a pull gage to the outside handle and check the torque required to operate the latching mechanism while closing the door. The torque required to start rotation of the handle should not exceed 250 inch-pounds, and the torque required to continue the latching movement should not exceed 100 inch-pounds. (17) Check the torque required to operate the latching mechanism while opening the door. The torque required to start rotation of the handle should not exceed 250 inch-pounds, and the torque required to continue the latching mechanism movement should not exceed 100 inch-pounds. NOTE: Multiply the distance in inches from the handle shaft to the attachment point of the pull gage times the reading in pounds on the gage to measure the inch-pounds of torque required for actuation of the latch mechanism. If the actuating torque does not fall within the prescribed limits, adjust the door posts and cable rigging until the proper torque is acquired. (18) Install all upholstery that was removed to facilitate latch adjustment.

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Figure 202 Airstair Door Latching Mechanism Adjustment

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Figure 203 Cable Drum and Camlock Assembly

Figure 204 Turnbuckle Safety Clip Installation

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Figure 205 Camlock Latch Adjustment

Figure 206 Camlock Latch and Latch Post Adjustment

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Figure 207 Airstair Door Cable Tension Graph

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Figure 208 Airstair Door Pressure Lock

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5. AIRSTAIR DOOR FORWARD AND AFT HANDRAIL A. Upper Forward And Upper Aft Handrail Removal NOTE: This procedure is typical for both the forward and aft upper handrails. CAUTION: To prevent damage to the airplane, a support shall be placed beneath the airstair door whenever any handrail(s) are not fully installed. (1) Open airstair door (37) and place a support under the door to remove all tension from the handrail assemblies (Ref. Figure 209). (2) Remove nut (3), washer (2), screw (4) and spacer (5) attaching clevis (6) to anchor (1). (3) Remove nut (10), washer (9), screw (12) and bushing (11) attaching clevis (8) to link (13). (4) Remove handrail (7) from the airplane. (5) Inspect handrail (7) for damage and wear; replace as necessary.

B. Upper Forward And Upper Aft Handrail Installation NOTE: This procedure is typical for both the forward and aft upper handrails. CAUTION: To prevent damage to the airplane, a support shall be placed beneath the airstair door whenever any handrail(s) are not fully installed. (1) Attach clevis (6) to anchor (1) by installing spacer (5), screw (4), washer (2) and nut (3) (Ref. Figure 209). NOTE: When the term “clockwise” or “counterclockwise” is used, the following assumptions are made: 1) the mechanic installing the handrails is standing at the foot of the open airstair door facing inboard; 2) the handrail base (Clevis 6 or 25) has been installed; 3) The mechanic will hold the handrail end (Clevis 8 or 18) towards him, looks towards the installed base of the handrail, and twist the handrail clockwise or counterclockwise based on this orientation. (2) Using half turn increments, preload handrail (7) by facing inboard and twist handrail clevis (8) clockwise for the upper forward handrail or counterclockwise for the upper aft handrail no more than 1.5 total twists. (3) Attach clevis (8) to link (13) by installing bushing (11), screw (12), washer (9) and nut (10). (4) Perform the AIRSTAIR DOOR HANDRAIL FUNCTIONAL CHECK procedure in this section.

C. Lower Forward And Lower Aft Handrail Removal NOTE: This procedure is typical for both the forward and aft lower handrails. CAUTION: To prevent damage to the airplane, a support shall be placed beneath the airstair door whenever any handrail(s) are not fully installed. (1) Open airstair door (37) and place a support under the door to remove all tension from the handrail assemblies (Ref. Figure 209). (2) Remove nut (20), washer (21), screw (24) and washer (23) from clevis (25). Page 212 May 1/10

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (3) Remove spacers (26), spring (27) and bushing (22) attaching clevis (25) to anchor (28). (4) Remove nut (14), washer (15), screw (17) and bushing (16) attaching clevis (18) to link (13) and remove handrail (19) from the airplane. (5) Inspect handrail (19) for damage and wear; replace as necessary.

D. Lower Forward And Lower Aft Handrail Installation NOTE: This procedure is typical for both the forward and aft lower handrails. CAUTION: To prevent damage to the airplane, a support shall be placed beneath the airstair door whenever any handrail(s) are not fully installed. (1) Position clevis (25) over anchor (28) and install bushing (22) through the clevis and bracket (Ref. Figure 209). (2) Position spring (27) and spacers (26) over clevis (25) so that pressure is exerted toward the airstair door (37). (3) Attach clevis (25) to anchor (28) by installing screw (24), washers (21 and 23) and nut (20). NOTE: When the term “clockwise” or “counterclockwise” is used, the following assumptions are made: 1) the mechanic installing the handrails is standing at the foot of the open airstair door facing inboard; 2) the handrail base (Clevis 6 or 25) has been installed; 3) The mechanic will hold the handrail end (Clevis 8 or 18) towards him, looks towards the installed base of the handrail, and twist the handrail clockwise or counterclockwise based on this orientation. (4) Using half turn increments, preload handrail (19) by facing inboard and twist handrail clevis (18) clockwise for the lower forward handrail or counterclockwise for the lower aft handrail no more than 1.5 total twists. (5) Attach clevis (18) to link (13) by installing bushing (16), screw (17), washer (15) and nut (14). (6) Perform the AIRSTAIR DOOR HANDRAIL FUNCTIONAL CHECK procedure in this section.

6. AIRSTAIR DOOR FORWARD AND AFT POST A. Removal NOTE: This procedure is typical for both the forward and aft airstair door post removal. CAUTION: To prevent damage to the airplane, a support shall be placed beneath the airstair door whenever any handrail(s) are not fully installed. (1) Open airstair door (37) and place a support under the door to remove all tension from the handrail assemblies (Ref. Figure 209). (2) Remove nut (10), washer (9), screw (12) and bushing (11) attaching clevis (8) to link (13). (3) Remove nut (14), washer (15), screw (17) and bushing (16) attaching clevis (18) to link (13). (4) Remove nut (30), washer (31) and bolt (32) from airstair door post assembly (29). (5) Remove bushing (35) and airstair door post assembly (29) from door brackets (36).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (6) Remove spring (33) and bushing (34) from door bracket (36).

B. Installation CAUTION: To prevent damage to the airplane, a support shall be placed beneath the airstair door whenever any handrail(s) are not fully installed. NOTE: Preloading the airstair door handrails helps to avoid damaging the handrails when closing the airstair door. Preloading is accomplished by twisting the handrails during installation. Because the condition of airstair door handrails vary, there are no specific preloading requirements regarding the total number of twists to be used. It is required that the handrail preload not exceed 1.5 total twists and not interfere in any way during the opening or closing of the airstair door. (1) Position bushings (34) in spring (33) and position between door brackets (36) (Ref. Figure 209). (2) Position post assembly (29) onto spring (33) arm and door bracket (36). (3) Install bushing (35) through door brackets (36), bushing (34) and spring (33). NOTE: Install bolt (32) with the head of the bolt facing outward. (4) Install bolt (32) through bushing (35) and door brackets (36). Install washer (31) and nut (30). NOTE: When the term “clockwise” or “counterclockwise” is used, the following assumptions are made: 1) the mechanic installing the handrails is standing at the foot of the open airstair door facing inboard; 2) the handrail base (Clevis 6 or 25) has been installed; 3) The mechanic will hold the handrail end (Clevis 8 or 18) towards him, looks towards the installed base of the handrail, and twist the handrail clockwise or counterclockwise based on this orientation. (5) Using half turn increments, preload handrail (7) by facing inboard and twist handrail clevis (8) clockwise for the upper forward handrail or counterclockwise for the upper aft handrail no more than 1.5 total twist. (6) Attach clevis (8) to link (13) by installing bushing (11), screw (12), washer (9) and nut (10). (7) Using half turn increments, preload handrail (19) by facing inboard and twist handrail clevis (18) clockwise for the lower forward handrail or counterclockwise for the lower aft handrail no more than 1.5 total twists. (8) Attach clevis (18) to link (13) by installing bushing (16), screw (17), washer (15) and nut (14). (9) Perform the AIRSTAIR DOOR HANDRAIL FUNCTIONAL CHECK procedure in this section.

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7. AIRSTAIR DOOR HANDRAIL A. Functional Check The purpose of this functional check is to ensure that the airstair door handrails do not become damaged and do not interfere with the operation of the door. The handrails should be installed so that as the door closes the handrails twist away from the edges of the door to avoid being pinched and damaged. Also, with the airstair door in the closed position, the handrails must not be situated in a way that may interferer with the door being opened. (1) Perform a functional check of the airstair door handrails by performing the following Steps: NOTE: While performing this functional check, do not pull on the handrails from the inside of the airplane to close the airstair door. The handrails should be allowed to move on their own as the door is closed. Observe the handrail behavior from inside the airplane as the door is slowly closed by someone outside the airplane. (a) Inspect the airstair door decals for installation, condition and legibility on the interior and exterior of the airstair door assembly (Ref. Chapter 11). (b) With weight on the bottom airstair door step, adjust the turnbuckle on the lower forward handrail (19) so the stairs are level and the load is shared equally by the forward and aft handrails (Ref. Figure 209). (c) With assistance, observe from inside the airplane the behavior of the airstair door handrails as the door is closed. The handrails should not interfere in any way with the operation of the airstair door during closing. If the handrails cause any interference, disconnect the handrail at clevis (8 or 18) and return to UPPER FORWARD AND AFT HANDRAIL INSTALLATION Step (2) and LOWER FORWARD AND AFT HANDRAIL INSTALLATION Step (4). (d) With assistance, observe from inside the airplane the behavior of the airstair door handrails as the door is opened. The handrails should not interfere in any way with the operation of the airstair door during opening of the door. If the handrails cause any interference, disconnect the handrail at clevis (8 or 18) and return to UPPER FORWARD AND AFT HANDRAIL INSTALLATION Step (2) and LOWER FORWARD AND AFT HANDRAIL INSTALLATION Step (4). (e) When no interference is observed, repeat Steps (1) (c) and (1) (d) at least three times to ensure proper operation of the airstair door handrails. If proper operation cannot be obtained, it may be necessary to replace one or more of the handrails.

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A

2

1 37

3 6

5 4

35

7

36 34 32

33 31 30

29 10

8 9

11

12

13 14 15 18

16

28

17 19

25

20 22

26

24 23 DETAIL

A

Figure 209 Airstair Door Aft Handrails

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26

27

26

1. ANCHOR 2. WASHER 3. NUT 4. SCREW 5. SPACER 6. CLEVIS 7. HANDRAIL 8.CLEVIS 9.WASHER 10. NUT 11. BUSHING 12. SCREW 13. LINK 14. NUT 15. WASHER 16. BUSHING 17. SCREW 18. CLEVIS 19. HANDRAIL 20. NUT 21. WASHER 22. BUSHING 23. WASHER 24. SCREW 25. CLEVIS 26. SPACERS 27. SPRING 28. ANCHOR 29. POST 30. NUT 31.WASHER 32. BOLT 33. SPRING 34. BUSHING 35. BUSHING 36. BRACKET 37. DOOR

UC52B 022737AB.AI

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

8. AIRSTAIR DOOR DAMPER A. Replacement (1) Remove the bolts and washers attaching the door damper to the fuselage frame support brackets. (2) Service the replacement damper, if required. Refer to the applicable Steps of the AIRSTAIR DOOR DAMPER (HYDRAULIC SNUBBER) procedures. (3) Attach the barrel end of the damper to the fuselage frame and the piston rod end to the airstair door support bracket with bolts and washers. Shim as necessary with AN960-416 washers to align the damper ends and eliminate side play. NOTE: If the piston rod end requires adjustment, lock the threads with thread locking compound (59, Table 1, Chapter 91-00-00).

9. AIRSTAIR DOOR SEAL A. Removal and Installation (1) Disconnect all electrical power. (2) Open the airstair door. (3) Disconnect the electrical connector adjacent to the hinge door. (4) Remove the threshold plate. (5) Remove the bolt and washers attaching the door damper to the fuselage frame. (6) Support the door and remove the screw, washer, spacer, and bushing attaching the handrail cable eye to the fuselage frame. (7) Index mark the seal splice at the top of the door frame. (8) Remove the old seal from the door seal retainer. (9) Apply soapstone powder to the new seal and the retainer. NOTE: To ensure retention of the seal in the track, the seal may be cemented in place with adhesive (43, 145, 167 or 193, Table 1, Chapter 91-00-00). Use adhesive in accordance with manufacturer’s instructions. (10) Starting with the seal splice at the marked location on the door, work the seal into the retainer. A thin flat strip of phenolic or Plexiglas with a rounded end will facilitate installation. (11) Clean off soapstone powder as necessary. (12) Install the bushing, spacer, washer, and screw attaching the handrail cable eye to the fuselage frame.

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(13) Install the bolt and washer attaching the door damper to the fuselage frame. (14) Connect the electrical connector adjacent to the door hinge. (15) Engage the cabin door circuit breaker, with power ON; observe that the FWD CABIN DOOR light is illuminated.

B. Repair NOTE: This repair procedure allows repair of the bottom portion of the airstair door seal. (1) Open the airstair door. (2) Remove the screws securing the threshold hinged plate and fold plate forward. (3) Measure 12 inches up from the bottom of the door seal and place a mark on the seal on both sides of the door opening. (4) Cut the seal at the mark on both sides and carefully remove the damaged bottom portion of the seal from the seal retainer. (5) Using the removed seal as a guide for length, cut the new seal to fit. NOTE: Use adhesive (43, 145, 167 or 193, Table 1, Chapter 91-00-00) in accordance with manufacturer’s instructions. (6) Apply a bead of adhesive (43, 145, 167 or 193, Table 1, Chapter 91-00-00) to the inside of the seal retainer to bond the seal in the retainer. (7) Starting with the seal splice on one side, work the new door seal into the retainer. A tongue depressor or a thin flat strip of phenolic or plexiglas with a rounded end will help facilitate installation. (8) Clean off excess adhesive as necessary. (9) Clean joint splice areas and splice tube material with cleaner (14, Table 1, Chapter 91-00-00) and allow to dry thoroughly. (10) Apply a very thin coat of adhesive (43, 145, 167 or 193, Table 1, Chapter 91-00-00) to one splice tube and carefully insert one end of the splice tube into the old seal and the other end of the splice tube into the end of the new seal. Repeat for the other joint. NOTE: Do not close the airstair door. Allow adhesives to dry. Refer to adhesive specification. (11) Fold the hinged threshold plate into position and secure with screws. (12) Close the airstair door after adhesive has cured.

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DOORS EMERGENCY EXIT DESCRIPTION AND OPERATION

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1. GENERAL The Model 1900 Airliner is semimonocoque in construction and is pressurized between bulkheads 84.00 and 557.50. When making repairs or modifications that create a break in the pressure vessel, the mating surface must be sealed with the proper sealer described in Table 2, Chapter 91. To ensure effective bonding of the sealers, clean thoroughly all mating surfaces, mating parts and rubber seals. Information pertaining to the cabin airstair doors, cargo door, baggage compartment door and emergency exits is included in Chapter 52. Maintenance practices for the horizontal stabilizer and vertical stabilizer are included in Chapter 55. Information pertaining to the windshield and the cabin windows is included in Chapter 56. Information pertaining to the wing sections and the one piece spar is included in Chapter 57. If repair to a structural member is necessary, refer to the MODEL 1900 SERIES AIRLINER SERIES STRUCTURAL REPAIR MANUAL P/N 114-590021-9.

Figure 1 Emergency Exit Door

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DOORS EMERGENCY EXIT MAINTENANCE PRACTICES

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1. PROCEDURES 2. EMERGENCY EXIT DOOR A. Removal To remove the emergency exit door, pull the release handle out and pull down on the door. CAUTION: Never open the escape hatch when the cabin is pressurized.

B. Installation (1) To install the door, place the lower edge of the door to the outboard side of the two adjustable lugs on the lower exit frame. (2) Connect the electrical plug and push the top of the door upward and inward. (3) Push the handle in until the latches have closed completely. NOTE: When installing a new emergency exit door, the adjustable side lugs will have to be adjusted with the cabin pressurization on. This must be done to achieve proper matching of the door with the airplane contour. Adjust the latching mechanism. Refer to LATCHING MECHANISM ADJUSTMENT in this section.

C. Latching Mechanism Adjustment Adjust the turnbuckle to a length required to securely engage the latches and obtain a proper seal between the door seal and frame, then install a locking clip on the turnbuckle. The door is properly adjusted when 15 ± 5 pounds of pull is required to release the hatch. The preferred pull unpressurized release is 15 to 20 pounds.

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DOORS EMERGENCY EXIT LATCH MECHANISM MAINTENANCE PRACTICES

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1. LATCH MECHANISM A. Removal (1) Perform the EMERGENCY EXIT DOOR REMOVAL procedure (Ref. Chapter 52-20-00). (2) Perform the ESCAPE HATCH UPHOLSTERY REMOVAL procedure (Ref. Chapter 25-20-03). (3) Remove the screws (30) and washers (29) attaching the EXIT-PULL marker (28) to the door handle assembly (27) and remove the marker from the door handle (Ref. Figure 201). (4) Remove cotter pin (23), washer (24), spacer (25), spring (41), spacer (40) washer (39) and clevis pin (38) and disconnect clevis end (26) from support (37). (5) Remove nut (51), washer (50), bolt (49), handle assembly (27) with bushing (77). (6) Remove cotter pin (31), washer (32) and clevis pin (35). (7) Remove nut (45), washer (46), bolt (48), handle lock (33) with bushing (34). (8) Remove stop bolt (47). (9) Remove cotter pin (20), washers (19 and 43) and clevis pin (44) from clevis end (21) and remove turnbuckle (22). (10) Remove forward stop assembly: (screw (55), spacer (54), washer (53) and nut (52)). Remove aft stop assembly (opposite of forward, not shown). (11) Remove nut (56), washer (57), bolt (17), hook assembly (16) with bushing (78). (12) Remove aft stop assembly: (screw (13), washer (14) and bushing (15)). Remove forward stop assembly (opposite of aft, not shown). (13) Remove three screws (67) and washers (68) and remove the exterior handle assembly: (items 1 thru 7 and 72 thru 76). Disassemble as necessary to inspect or replace parts. (14) Remove spring (66). NOTE: Items (9 thru 12) are attached to door (8) by rivets and will stay with the door. (15) Remove cotter pin (69), washer (70) and clevis pin (71) and remove the lock linkage assembly: (items 36 and 58 thru 65). Disassemble as necessary to inspect or replace parts.

B. Installation NOTE: Items (9 thru 12) are attached to door (8) by rivets and should be with the door (Ref. Figure 201). Assemble lock linkage assembly as necessary if disassembled for inspection or parts replacement.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (1) Position the lock linkage assembly: (items 36 and 58 thru 65) and install clevis pin (71), washer (70) and cotter pin (69). (2) Install spring (66). NOTE: Assemble exterior handle assembly as necessary if disassembled for inspection or parts replacement. (3) Position the exterior handle assembly: (items 1 thru 7 and 72 thru 76) and install three screws (67) and washers (68). (4) Install aft stop assembly: (screw (13), washer (14) and bushing (15)). Install forward stop assembly (opposite of aft, not shown). (5) Position hook assembly (16) with bushing (78) and install bolt (17), washer (57) and nut (56). (6) Install forward stop assembly: (screw (55), spacer (54), washer (53) and nut (52)). Install aft stop assembly (opposite of forward, not shown). NOTE: Ensure right-hand thread clevis end (21) is at the top position (Ref. Mandatory Service Bulletin No. 2740). Ensure both clevis ends (21 and 26) are installed and safetied with lock clip (42). (7) Position turnbuckle (22) with clevis end (21), install clevis pin (44), washers (19 and 43) and cotter pin (20). (8) Position handle assembly (27) with bushing (77) and install bolt (49), washer (50) and nut (51). (9) Position clevis end (26) with spring (41) and install clevis pin (38), washer (39), spacer (40), spacer (25), washer (24), and cotter pin (23). (10) Install stop bolt (47). (11) Position handle lock (33) with bushing (34) and install bolt (48), washer (46) and nut (45). (12) Install clevis pin (35), washer (32) and cotter pin (31). (13) Install the screws (30) and washers (29) attaching the EXIT-PULL marker (28) to the door handle assembly (27). (14) Perform the EMERGENCY EXIT DOOR LATCHING MECHANISM ADJUSTMENT procedure (Ref. 52-20-00). (15) Cycle the emergency exit door open and closed at least three times with the exterior and interior handles. Push and pull on the door to confirm security. (16) Perform the ESCAPE HATCH UPHOLSTERY INSTALLATION procedure (Ref. Chapter 25-20-03). (17) Perform the EMERGENCY EXIT DOOR INSTALLATION procedure (Ref. Chapter 52-20-00).

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Figure 201 Emergency Exit Door Latch Mechanism

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DOORS CARGO/NOSE BAGGAGE COMPARTMENT DOORS DESCRIPTION AND OPERATION

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1. GENERAL A. Cargo Door The cargo door is hinged at the top and incorporates a bottom latching mechanism. Gas springs charged with nitrogen are attached to the door frame and to brackets on each side of the door for automatically opening the door after the initial opening force is applied. Two small gas springs, attached to brackets on the door and on the fuselage structure above the door, cushion the closing force of the large gas springs on the sides of the door. The door is counterbalanced to remain in the open position; however, a support rod is provided for additional support and to hold the door open during gusty conditions. The gas springs also apply a closing force to assist in closing and latching the door. A rubber seal is held in place by retainers around the periphery of the door to effectively seal the pressure vessel when the cabin is pressurized. The latching mechanism for the cargo door is operated by two interconnected handles, one on the outside and one on the inside of the door. The handles are attached to cable drums. When the handles are rotated, cables attached to the drums wind or unwind in response to the direction of handle movement. The cables from the handles are routed around splined drums at each of the camlock latches on the forward and aft sides of the door. The splines in the drums engage splines on the camlock shafts of the latch so that the rotation of the drum, when actuated by the cable from the handle, rotates the face of the latch to engage the latch post in the door frame. The latch cables are adjusted by turnbuckles to a tension sufficient to ensure proper latching and unlatching of the latches in response to handle movement. A splined cam on the end of the camlock shaft opposite the latch face actuates switches mounted on brackets at the upper latches to indicate when the door is closed and latched. The movement of the inside latch handle fore and aft is transferred by a tube running the width of the cargo door to four latch pins that engage latch brackets on the door with lugs attached to the doorsill of the fuselage. To open the cargo door after it is unlatched, push on the bottom of the door. After the cargo door is manually opened a few feet, gas springs take over and raise the door to the fully open position. To close the cargo door, pull it down and inboard. The gas springs will resist the closing effort until the door is only open a few feet. Then, as the springs move overcenter, they begin applying a closing force to the door. The door closes against an inflatable rubber seal around the opening in the cargo door frame. When the cabin is pressurized, air seeps into the rubber seal through small holes drilled in the side of the seal. The higher the cabin differential pressure, the more the seal inflates. This is a passive seal system and has no mechanical connection to a bleed air source.

B. Nose Baggage Compartment Door The nose baggage compartment door has two hinges located at the top which permits the door to be opened upward. Three latch pins, one located on each side of the door and one at the bottom, are connected by couplings to a bellcrank, located at the center of the door. A microswitch, actuated by the forward latching pin, will extinguish the DOOR OPEN warning light when the door is closed and locked.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL To open the nose baggage door, use the thumb to press in on the forward portion of the door latch handle. As the forward portion is pressed in, the aft portion of the latch will protrude past the recess in the door. While holding in on the forward portion of the handle, grasp the aft portion and rotate the handle downward until the door is completely unlocked. Raise the door up to the open position. The door is equipped with a gas spring to assist opening and to steady the door in gusty conditions. To close the door, exert downward pressure to overcome the gas spring. While holding the door in the closed position, rotate the handle upward; the handle will automatically return to the locked position.

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DOORS CARGO/NOSE BAGGAGE COMPARTMENT DOORS MAINTENANCE PRACTICES

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1. CARGO DOOR A. Removal (1) To open the cargo door, proceed as follows: (a) Pull the safety latch release pin and move the handle to the open position. (b) Attach one end of the door stabilizer assembly (door support rod) to the ball stud located on the forward side of the door. Ensure the detent pin is installed. CAUTION: Avoid side loading of the gas spring assemblies to prevent damage to the internal mechanisms. (c) Push out on the cargo doorsill and allow the door to swing open. (d) Attach the free end of the door support rod to the ball stud on the forward fuselage door frame. (2) Disconnect the electrical connector at the upper fuselage doorframe. (3) Remove the retaining clips and detach the four gas cylinder assemblies at the bracket mounted ball studs on the upper fuselage doorframe (2 places), and on the lower forward and aft door surface (2 places). (4) Remove the screws and the hinge pin retainers on each end of the hinge. (5) Support the forward and aft portions of the door. Remove the door support rod. (6) Remove the hinge pin from the hinge halves and carefully remove the door.

B. Installation (1) If necessary, lubricate the cargo door. Refer to Lubrication Table, Chapter 12-20-00. (2) Carefully move the door into position, mate the hinge halves, and install the hinge pin (Ref. Figure 201). (3) Attach the door support rod to the door and fuselage ball studs. Ensure the detent pin is installed. (4) Install the hinge pin retainers on each end of the hinge with screws. (5) Attach the gas spring assemblies to the bracket mounted ball studs on the upper fuselage door frame and on the lower forward and aft sides of the door. Secure with the retaining clips. (6) Connect the door electrical connector at the upper fuselage doorframe. (7) With electrical power on, engage the cabin door circuit breaker and observe that the CABIN DOOR light is ON.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (8) Station two maintenance personnel outside the airplane at each end of the door hinge to observe for a binding condition between the hinge pin retainers and the edge of the door skin as the door is closed. (9) To close the cargo door, proceed as follows: (a) Detach the lower end of the door support rod. Ensure the detent pin is installed. (b) Firmly grasp the free end of the door support rod while exerting a downward force to overcome the pressure of the gas spring assemblies. Remove the door support rod after the gas spring assemblies pass the overcenter position. (c) If necessary, apply an inward force against the lower portion of the door to compress the door seal, permitting the latching mechanism to engage. (d) With electrical power ON, observe that the cabin door light is out.

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Figure 201 Cargo Door Installation

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C. Prerigging The following rigging instructions may be accomplished with the door on a work bench or installed on the airplane in the open position: (1) Place the cargo door latching handle in the closed and locked position. (2) Adjust the latching mechanism pushrod so that the untapered shoulder of each latch pin extends a minimum of 0.15 inch beyond its respective bracket (Ref. Figure 202). Shim the pin with AN960-616 or AN960-616L washers as necessary to obtain this dimension. (3) Place the cargo door latching handle in the unlocked position and adjust the latch pins until each pin clears the bracket by 0.18 inch (Ref. Figure 202). (4) Adjust the cargo door latch handle stop bolt to maintain the 0.18 inch clearance with a handle rotation of 85 1/2° from open to lock. If the 85 1/2° rotation cannot achieved, readjust the cargo door latch handle stop bolt until 85 1/2° of latch handle travel is assured (Ref. Figure 203). (5) Install the door latch cable drum on the door with the cable slot in the drum (where the cable originates) on the inboard side of the latch housing. This is a starting point only and is not absolute, as this position may change with further rigging. (6) Rig the cables as indicated in the Cable Tension Graph (Ref. Figure 208). (7) Safety the turnbuckle (Ref. Figure 204). NOTE: If the preceding Steps were done on a work bench, the cargo door must now be installed on the airplane prior to proceeding. (8) With the latch handle in the fully open position, insert the splines of each camlock shaft in the splines of its respective cable drum so that the cutout in the latch face is centered with respect to the arc made by the latch when the door is closing (Ref. Figure 205). (9) Engage the splines of each latch cam with the splines on the small end of the camlock shaft so that the cam just contacts its stop pin (Ref. Figure 206). On the top latch on each side of the door where the electrical switches are located, back the cam off one spline from the point where it contacts the stop pin, then install the washer and nut securing it in place. CAUTION: Anytime the camlock shafts are repositioned with respect to the cable drums, the latch cams must be reindexed to the camlock shaft to ensure full travel of the latch handle.

2. CARGO DOOR LATCHING MECHANISM A. Adjustment NOTE: Ensure that all other cargo door rigging requirements are complied with before performing this procedure. The gas springs must be disconnected and the door seal removed before these instructions are performed. (1) Disconnect the door latch mechanism pushrod from the latch pin guide tube, and tie the rod back out of the way. (2) Manually engage the latch pin to hold the door in the closed position. Page 204 May 1/10

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (3) With the door closed and held in place by the bottom latch pin, adjust each camlock latch post on the side of the doorframe so that the post maintains a clearance of 0.062-inch with the face of the camlock latch. When the camlock latch is in the latched position shim the post as necessary with P/N 101-514033-1 laminated washers to obtain this clearance (Ref. Figure 207). (4) With the door closed and held in place by the bottom latch pins, adjust each latch post on the side of the door frame until the post's point of contact with the latch face of the camlock shaft aligns with the closing arc made by the latch when the door is fully closed (Ref. Figure 209). To make this adjustment, rotate the eccentric bushing, in conjunction with the eccentric latch post, until the post is properly aligned with the latch, then tighten and torque the nut 290 to 410 inch-pounds. Safety the nut securing the post in place. A special latch post wrench (P/N 101-590052-1) is required to adjust the eccentric bushing. NOTE: Coating the latch post with “Blue Dykem”, or equivalent, will facilitate pinpointing its area of contact with the face of the latch (Ref. Figure 209). (5) Paint the grooves on the camlock assemblies with Calypso Orange paint or equivalent. (6) With the door latch handle in the closed and latched position, align the orange grooves on the camlock latch assemblies with the slots on the indicator assemblies mounted on the door frame (Ref. Figure 210). (7) Reconnect the door latching mechanism pushrod. (8) With the door in the closed and latched position, align the pointer on the latching pin guide rod with the target assembly mounted below the lower pulley (Ref. Figure 211). (9) Adjust the eccentric bushing of the roller guide assembly in the doorframe on each side of the cargo door adjacent to the upper latches until the roller just contacts the guide bracket on the door when the door is fully latched (Ref. Figure 212). After the roller is properly positioned on the door frame with respect to the guide bracket on the door, tighten the nut securing the roller in place. (10) Close the door and check that the camlock latches on the sides of the cargo door and the bottom latches operate properly when actuated by the latching handles. (11) Adjust the doorsill, lower latch, lower latch handle, rotary handle, and forward and aft latch switches as necessary to ensure that the cabin door warning system functions properly. Adjust the switches in accordance with the procedures described under CARGO AND AIRSTAIR DOOR WARNING SWITCHES in 52-70-00. (12) Work the door seal into the retainer channel around the perimeter of the door with a tongue depressor or its equivalent. (13) Engage the ends of the gas spring assemblies with the ball studs securing them to the brackets on the sides and over the cargo door, then lock each in place with a retaining clip. (14) Check the latch mechanism for proper operation in the open and closed positions. CAUTION: Adjustment of the two cargo door alignment guides are not to be made until all other cargo door mechanism rigging requirements are completed. (15) Center the cargo door lower latch mechanism with the lower doorsill lugs. Adjust the striker plates on the forward and aft fuselage door frame, using shims (P/N 101-514190-11) as required to maintain a 0.06 to 0.03 inch clearance between the roller and striker plate.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (16) Install all upholstery that was removed during the adjustment procedure. (17) Install the gas springs.

3. GAS SPRING ASSEMBLY A. Replacement Under no circumstances shall any attempt be made to service the gas spring assemblies. The following criterion may be used to determine replacement of the gas spring assemblies. (1) Open the cargo door and install the door support rod. NOTE: It should be noted that sluggish operation of the gas springs is normal during exceptionally cold weather and does not indicate a need for replacement of the part. (2) During door closing, after the gas spring assemblies pass overcenter position, the downward force is insufficient to close the door. (3) Gas spring rods are nicked or bent.

4. CARGO DOOR SEAL A. Replacement NOTE: This procedure is typical for both the forward and aft airstair door post removal. CAUTION: To prevent damage to the airplane, a support shall be placed beneath the airstair door whenever any handrail(s) are not fully installed. (1) The door opens sluggishly or fails to fully open. CAUTION: The door support rod will have to be detached temporarily during seal replacement. During this time the cargo door must be supported. (2) Disengage the cabin door circuit breaker. (3) Disconnect the electrical connector at the upper fuselage door frame. (4) Detach the gas spring assemblies at the door. (5) Index mark the seal splice at the top of the door frame. (6) Remove the old seal from the door seal retainer. (7) Apply soapstone powder to the new seal and the seal retainer. (8) Starting with the seal splice at the marked location on the door frame, work the seal into the retainer. A thin flat strip of phenolic or Plexiglas with a rounded end will facilitate installation. (9) Clean off the soapstone powder as necessary.

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(10) Attach the gas spring assemblies to the door. (11) Connect the electrical connector at the top of the door frame. (12) Engage the cabin lights circuit breaker, with power ON; observe that the CABIN LIGHT is on.

5. NOSE BAGGAGE COMPARTMENT DOOR A. Removal (1) Raise the door to the fully open position (Ref. Figure 213). (2) Disconnect the gas spring assembly. (3) While holding the door, remove the hinge pin and lift the door free.

B. Installation (1) Align the two hinge halves and install the hinge pin in the hinge halves. (2) Reconnect the gas spring assembly. (3) Lubricate all working parts of the latching system (Ref. Chapter 12-20-00). (4) Close and latch door.

C. Seal Replacement (1) Remove the old seal and wash all traces of old adhesive from the door frame with solvent (14, Table 1, Chapter 91-00-00). (2) Wash the new seal with solvent (14, Table 1, Chapter 91-00-00) to remove the soapstone powder preservative. (3) Apply adhesive (16, Table 1, Chapter 91-00-00) to the edge of the door frame on the fuselage. Work the seal around the door frame edge in each direction from the center of the seal. (4) Cut the outer seal as required to allow the seal to compress around the corners. Hand form the seal to the corners to achieve a smooth fit.

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Figure 202 Latch Pin Adjustment

Figure 203 Latch Handle Adjustment Page 208 May 1/10

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Figure 204 Safetying Cable Turnbuckles

Figure 205 Camlock Latch Adjustment

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Figure 206 Cable Drum and Camlock Assembly

Figure 207 Latchpost Clearance Adjustment

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Figure 208 Airstair Door Pressure Lock

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Figure 209 Camlock Latch and Latchpost Adjustment

Figure 210 Camlock Alignment

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Figure 211 Door Latch Target Indicator Alignment

Figure 212 Roller Guide Adjustment

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Figure 213 Nose Baggage Compartment Door

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Figure 213 Nose Baggage Compartment Door

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DOORS CARGO AND AIRSTAIR DOOR WARNING DESCRIPTION AND OPERATION

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1. GENERAL The warning system for the cargo and cabin entrance (airstair) doors utilize nine switches, five on the cargo door and four on the airstair door, to indicate that all door handles and latches are safety positioned when the doors are closed and latched. Two separate switch circuits are used to illuminate the FWD CABIN DOOR and AFT CABIN DOOR lights in the warning annunciator panel to indicate that the cargo and airstair doors are not properly secured. The circuit for the cargo door is wired in parallel through the doorsill, forward and aft camlock latch, latch handle and the latch pin switches to send a signal to the annunciator fault detection card assembly, which in turn activates the AFT CABIN DOOR light in the warning annunciator panel to indicate that any one of the switches is not actuated. The doorsill, forward and aft camlock latch and the latch handle switches on the airstair door are wired in the same manner as the switches on the cargo door to illuminate the FWD CABIN DOOR light on the warning annunciator panel. The camlock switches utilized in both circuits are actuated by a cam attached to the shaft of the camlock latch that is rotated by the cable and handle drums; consequently, these switches sense the rotary motion of the handle and subsequent movement of the cables when the door is being locked or unlocked. If the cable breaks or the switch actuator or cam fails, the switch will return to the normally closed position to indicate an unlocked condition.

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1. PROCEDURES 2. CARGO AND AIRSTAIR DOOR WARNING SWITCHES Adjustment of the airstair door switch, the forward and aft latch switches and the latch handle switch may be accomplished by fabrication and using the test box in lieu of using the electrical system on the airplane (Ref. Figure 203). The test box is connected to the airstair door switch circuit through the electrical connector located just forward of the lower forward corner of the door. The test procedures in the following switch adjustment Steps will remain basically the same, but indication of switch actuation will be observed on the test box instead of through the airplane annunciator system (Ref. Figure 201 and 202).

3. AIRSTAIR DOOR DOORSILL SWITCH A. Adjustment The airstair doorsill switch is mounted on the left side of the airstair door frame where it can be actuated by the fuselage frame when the door is closed. (1) Close and latch the airstair door. (2) Turn ON the battery master switch, and the circuit breaker for the reading lights. (3) Remove the upholstery over the doorsill switch, to gain access to the switch mounting bracket. (4) Adjust the switch in its mounting bracket until the FWD CABIN DOOR light just goes out. Tighten the locknuts securing the switch in place. (5) Unlatch the door and ensure that the FWD CABIN DOOR light on the warning annunciator panel illuminates with any outward movement of the door.

4. AIRSTAIR DOOR FORWARD AND AFT LATCH SWITCHES A. Adjustment A switch is mounted on a bracket adjacent to the lower camlock latch on each side of the airstair door. As each latch rotates, its respective switch is actuated by a cam attached to the end of the camlock shaft opposite the latch face. NOTE: The airstair door must be closed and latched so that the doorsill switch is actuated during the following adjustments. (1) Turn ON the battery master switch and engage the circuit breaker for the annunciator indicator. (2) Remove the upholstery from the airstair door as necessary to gain access to the pressure lock plunger and to the switches mounted adjacent to the lower forward and aft camlock latches.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (3) While rotating the handle of the airstair door toward the locked position, move each latch switch in its slotted mounting bracket until the FWD CABIN DOOR light is extinguished when the pressure lock plunger rides in the keyway at a point 0.01- to 0.40-inch from where the plunger drops in the keyhole (Ref. Figure 201, Detail C). Tighten the mounting screws securing each switch in place at its mounting bracket. (4) Install all upholstery removed to facilitate switch adjustment. (5) Turn OFF the battery master switch and disengage the annunciator indicator circuit breaker.

5. AIRSTAIR DOOR LATCH HANDLE SWITCH A. Adjustment The latch handle switch is mounted on a bracket located adjacent to the pressure lock of the airstair door. The airstair door should be closed and latched so that the doorsill switch is actuated during the following adjustment: (1) Remove upholstery from the airstair door as necessary to gain access to the switch adjacent to the pressure lock. (2) Turn ON the battery master switch and engage reading lights circuit breaker. (3) While rotating the airstair door handle toward the locked position, move the switch in its slotted mounting holes until the FWD CABIN DOOR light on the warning annunciator panel extinguishes when the pressure lock plunger rides in the keyway at a point 0.04-inch or less from where the plunger drops into the keyhole (Ref. Figure 201, Detail D). Tighten the mounting screws securing the switch in place on its mounting bracket. (4) Install all upholstery removed to facilitate adjustment. (5) Turn OFF the battery master switch and disengage the cabin door and reading light circuit breakers.

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Figure 201 Airstair Door Switch Adjustment

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6. CARGO DOOR DOORSILL SWITCH A. Adjustment The cargo door doorsill switch is mounted on the left side of the door structure where it can be actuated by the doorsill when the cargo door is closed. (1) Remove the lower upholstery panel from the cargo door to gain access to the switch. (2) Turn ON the battery master switch and engage the circuit breaker for the reading lights. (3) Loosen the locknuts and adjust the switch in its mount until the CABIN DOOR WARNING light on the annunciator panel goes out when the cargo door is closed and latched, then tighten the locknut securing the switch in place. (4) Install the lower upholstery panel on the cargo door. (5) Turn OFF the battery master switch and the circuit breaker for the reading light.

7. CARGO DOOR LATCH HANDLE SWITCH A. Adjustment The latch handle switch is mounted on the handle support immediately aft of the handle. (1) Turn ON the battery master switch and the circuit breaker for the reading lights. (2) Remove upholstery as necessary to gain access to the latch handle switch. (3) Move the switch in its slotted mounting bracket until the CABIN DOOR warning light on the warning annunciator panel goes out as the lower handle approaches the point where the latch handle detent begins to engage the safety latch. (4) Install all upholstery removed to facilitate switch adjustment. (5) Turn OFF the battery master switch and the circuit breaker for the reading lights.

8. CARGO DOOR LATCH PIN SWITCH A. Adjustment The latch pin switch is mounted on the door structure immediately forward of the aft latch in the lower latch mechanism. The movement of the latch pin forward through the latch pin brackets and the lug in the doorsill actuates this threaded bushing type switch when the mechanism is fully latched. (1) Turn ON the battery master switch and the circuit breaker for the reading lights. (2) Remove the upholstery in the lower aft corner of the cargo door to gain access to the switch. (3) Place the door latch handle in the closed position. (4) Loosen the switch mounting screws and adjust the switch until the CABIN DOOR light on the warning annunciator panel goes out as the switch plunger is depressed against the latch pin. (5) Tighten the screws securing the switch to its mounting bracket. Page 204 Nov 1/09

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (6) Install the upholstery panels removed to facilitate switch adjustment. (7) Turn OFF the battery master switch and the circuit breaker for the reading lights.

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Figure 202 Cargo Door Switch Adjustment

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9. CABIN/CARGO “FWD CABIN DOOR/AFT CABIN DOOR” ANNUNCIATOR CIRCUITRY CHECK A. Cabin Door The following inspection and test should be performed after any of the door latches or switches are replaced or adjusted: (1) Ensure that the forward and aft (if installed) cabin door is closed and locked by the following: (a) Check the position of the safety arm and pressure lock plunger. (2) Check that the orange index marks on each of the six rotary camlocks align within the indicator windows.

B. Cargo Door (If Installed) The following inspection and test should be performed after any of the door latches or switches are replaced or adjusted: (1) Ensure that the cargo door (if installed) is closed and locked by the following: (a) Observe through the access cover window that the upper handle is in the locked position. (b) Check that the orange index marks on each of the four rotary camlocks align within the indicator windows. (c) Observe through the handle access cover window that the latch handle is in the locked position. (d) Observe through the window in the aft lower corner of the door that the orange colored indicator aligns with the orange stripe on the carrier rod. NOTE: The untapered shoulder of the latching pins must extend past each attachment lug.

C. Annunciator Circuitry (1) Turn the battery switch ON. With the cabin and cargo doors open and mechanisms in the locked position, ensure FWD CABIN DOOR and AFT CABIN DOOR annunciators are illuminated. (2) Check that, with doors closed but not locked, the FWD CABIN DOOR and AFT CABIN DOOR annunciators are illuminated. (3) Close and lock the doors. Check that the FWD CABIN DOOR and AFT CABIN DOOR annunciators extinguish.

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Figure 203 Airstair Door Switch Adjustment Test Box

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CHAPTER 53 - FUSELAGE TABLE OF CONTENTS SUBJECT

PAGE

FUSELAGE AND FLOOR ACCESS OPENINGS 53-00-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1

MAIN FRAME 53-10-00 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201

ATTACH FITTINGS 53-40-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Flight Compartment Seat Tracks . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Passenger Compartment Seat Tracks . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Flight Compartment Seat Track . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Wear Limits . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Passenger Compartment Seat Track . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Wear Limits . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Corrosion on the Seat Track . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201

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FUSELAGE FUSELAGE AND FLOOR ACCESS OPENINGS DESCRIPTION AND OPERATION

53-00-00 00

1. GENERAL Being of semimonocoque construction, the Model 1900 Airliner fuselage is pressurized to the skin between pressure bulk-heads at stations 84.00 and 557.50. All skin, bulkhead and structure points, plumbing and wiring connections passing through a pressure wall, access doors, windows, control cables, and torque shafts are sealed to minimize air leakage. The Model 1900 Airliner is equipped with forward and aft, left side swing down cabin entrance doors. The Model 1900C Airliner is provided with a forward left side swing down cabin entrance door and a left rear swing up cargo door. The swing down cabin entrance door provides a convenient stairway for boarding the airplane. The swing up cargo door provides a means of loading and unloading baggage as well as cargo. Three emergency exits are installed for quick egress should the need arise. Two of these exits are located on the right side of the airplane at the third and fifth cabin windows and the other exit on the left side at the fifth cabin window. Other fuselage openings include a swing up nose baggage door on the left side of the airplane, an avionics compartment door opposite the nose baggage compartment door, an Emergency Locator Transmitter (ELT) access door on the right side of the airplane below the stabilon, and access doors attached to each main landing gear and the nose landing gear (Ref. Chapter 6-50-00). For fuselage skin thickness (Ref. MODEL 1900 AIRLINER SERIES STRUCTURAL REPAIR MANUAL, P/N 114-590021-9). The subfloor area of the fuselage contains various electrical and mechanical systems components. Access to these components is gained by removing the floorboard panels (Ref. Chapter 6-50-00).

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FUSELAGE MAIN FRAME MAINTENANCE PRACTICES

53-10-00 200200

1. PROCEDURES A. Inspection For all inspection requirements, refer to the MODEL 1900/1900C AIRLINER STRUCTURAL INSPECTION MANUAL P/N 98-30937F or subsequent revision.

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FUSELAGE ATTACH FITTINGS DESCRIPTION AND OPERATION

53-40-00 00

1. GENERAL A. Flight Compartment Seat Tracks The flight compartment seats (pilot and copilot) are moveable forward and aft on tracks attached to the floor. Lock pins are controlled by the seat occupant to adjust and secure the seat in a selected position on the tracks. The pilot and copilot seats are equipped with support channels that are designed to move on the tracks while preventing seat and track separation. Four crew compartment seat tracks extend from FS 118.00 to FS 142.00 along left and right BL 11.00 and BL 21.00. Each seat track is secured to a track support with screws into nut plates on the track support structure.

B. Passenger Compartment Seat Tracks The passenger seats are mounted on one floor mounted seat track and one sidewall mounted seat track in a fixed position. Movement of the seat on the tracks is necessary only during installation and removal of the seat (Ref. Chapter 25-20-00).

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FUSELAGE ATTACH FITTINGS MAINTENANCE PRACTICES

200200

1. PROCEDURES 2. FLIGHT COMPARTMENT SEAT TRACK A. Wear Limits Excessive wear of the seat attach components and seat tracks will cause the crew seat to become unstable. Wear limits of the components have been established to determine when component replacement is necessary. These components include the seat channel on the seat base, seat lock pins and the seat tracks. The maximum wear limits are illustrated in Figure 201. Replacement of worn parts must be accomplished if the limits are exceeded. (1) The seat base must be replaced if measurements of the seat channel are less than the minimum dimensions in Detail A. (2) Replace the lock pins if the diameter is less than 0.220 inch or pin engagement into the seat track is less than 0.11 inch as shown in Detail D. (3) The seat tracks must be replaced if measurements of the tracks are less than the minimum dimensions shown in Figure 201. A slight gouge in the top surface (Detail B), caused by the lock pin, is not detrimental to the operation of the seat. The lock pin hole may be elongated in a forward and aft direction to a maximum of 0.315 inch long, measured at the 0.06 inch depth, before track replacement is necessary. The diameter of the holes must not exceed 0.270 inch below 0.06 inch depth.

3. PASSENGER COMPARTMENT SEAT TRACK A. Wear Limits As the seat tracks wear their strength is reduced. Refer to Figure 202 and Figure 203 for acceptable wear limitations which have been established for the passenger compartment seat tracks. If gouging or chafing is present and the depth does not exceed the values shown in Figure 202 and Figure 203, proceed with the following repair procedure. •

Remove any sharp edges using aluminum oxide sandpaper. Do not sand any deeper than the initial gouge or chafe.

•

Apply Alodine 1200 or 600 on any exposed surface.

If the minimum and maximum limitations shown in Figure 202 and Figure 203 cannot be met, contact Hawker Beechcraft Technical Support.

4. CORROSION ON THE SEAT TRACK If corrosion is found on the seat tracks, refer to Chapter 20-09-00 for information on corrosion removal. Inspect each seat track. Refer to Figure 202 and Figure 203 for acceptable corrosion limitations, which have been established for the passenger compartment seat tracks. If the minimum and maximum limitations shown in Figure 202 and Figure 203 can not be met, contact Hawker Beechcraft Technical Support.

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Figure 201 Flight Compartment Seat Track Wear Limits

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Y

SEAT TRACK LUG REF

0.45 OVER LUG REF

Y 0.03 WIDE BY 0.06 DEEP MAX IN CUTOUT AREA NOT OVER LUG

MAX 0.050 WEAR COMBINED ALLOWED ON UPPER AND LOWER SURFACE e.g. AS SHOWN BELOW: 0.02 + 0.03 = 0.050 MAX

0.050 MAX WEAR ALLOWED

0.065 MIN (0.020 IN FROM EDGE) 0.03 REF 0.297 MIN

0.02 REF 0.02 REF (IN FROM EDGE)

0.03 MAX

VIEW

Y-Y

UC53B 050248AB.AI

Figure 202 Passenger Compartment Floor Mounted Seat Track Wear Limits

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MAX 0.050 WEAR COMBINED ALLOWED ON UPPER AND LOWER SURFACE. E. G. AS SHOWN BELOW 0.02 + 0.03 = 0.050 MAX 0.065 MIN (0.020 IN FROM EDGE)

0.020 MAX (IN FROM EDGE)

0.03 MAX

0.03 REF

0.020 MAX (IN FROM EDGE) 0.02 REF UC53B 050249AA.AI

Figure 203 Passenger Compartment Sidewall Mounted Seat Track Wear Limits

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CHAPTER 54 - NACELLES TABLE OF CONTENTS SUBJECT

PAGE

GENERAL 54-00-00 Operation and Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Recommended Materials . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1

NACELLE INNER FENDER 54-10-01 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Nacelle Inner Fender (Without Kit No. 114-9801) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Nacelle Inner Fender (With Kit No. 114-9801) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 201 201 202 202 202

NACELLE PLATES/SKINS 54-30-00 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Upper-Aft Nacelle Fairing (UC-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

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NACELLE/PYLONS GENERAL OPERATION AND DESCRIPTION

54-00-00 00

1. GENERAL Structural units and associated components and members which furnish a means of mounting and housing the power plant or rotor assembly. Includes skins, longerons, belt frames, stringers, clamshells, scuppers, doors, nacelle fillets, attach/attached fittings etc. Also includes the structure of power plant cowling inclusive of the structural portion of the inlet whether or not integral with the airplane. Structural portions of the exhaust system are excluded where they are not integral with the airframe.

A. Recommended Materials The recommended materials listed in Table 1 as meeting federal, military or supplier specifications are provided for reference only and are not specifically required by Hawker Beechcraft Corporation. Any product conforming to the specification listed may be used. The products included in these Tables have been tested and approved for aviation usage by Hawker Beechcraft Corporation, by the supplier, or by compliance with the applicable specification. Generic or locally manufactured products which conform to the requirements of the specification may be used even though not included in the Tables. Only the basic number of each specification is listed. No attempt has been made to update the listing to the latest revision. It is the responsibility of the technician or mechanic to determine the current revision of the applicable specification prior to usage of the product listed. This can be done by contacting the supplier of the product to be used. Table 1 Recommended Materials MATERIALS 1. Sealant

SPECIFICATION

PRODUCT

SUPPLIER

AC-236

Advanced Chemistry & Technology Garden Grove, CA 92841

AC-240

Advanced Chemistry & Technology Garden Grove, CA 92841

PR-1440

PRC/Products Research Co. Burbank, CA 91504

WS-8020h

Royal Adhesives South Bend, IN 46628

CS3204

Chem Seal Division of Flamemaster Corp. Sun Valley, CA 91353

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NACELLE/PYLONS NACELLE INNER FENDER MAINTENANCE PRACTICES

54-10-01 200200

1. NACELLE INNER FENDER (WITHOUT KIT NO. 114-9801) A. Removal (1) Perform the REMOVING GROUND POWER procedure (Ref. Chapter 24-40-00). NOTE: Record the dimension of the up-position indicator switch (3) protruding below the mount bracket (5) for use during installation (Ref. Figure 201). (2) Remove the safety wire (7), remove jam nut (4) and remove up-position indicator switch (3). (3) Remove seven bolts (6) and remove mount bracket (5). (4) Remove all attach screws (2) from the inner fender. (5) If brake deice system is installed, remove screw (11), washer (10), clamps (9 and 13) and standoff (14). If no brake deice system installed, remove the patch plates (1). (6) Remove the grommet (8) and feed the up-position indicator switch (3) and harness through the opening. (7) Remove the inner fender from the airplane.

B. Installation (1) Perform the REMOVING GROUND POWER procedure (Ref. Chapter 24-40-00). (2) Feed the up-position indicator switch (3) and harness through the opening and install the grommet (8) (Ref. Figure 201). (3) Measure the distance between the fuel lines and the main landing gear up-position switch wire harness. If the distance measured is less than 1/2 inch, complete Steps (a) and (b). (a) Using tie straps (114, Table 1, Chapter 91-00-00), remove excess slack from the main landing gear up-position indicator switch wire harnesses by securing the harnesses to existing wire harnesses which route through the top of the main landing gear wheel well side panels. (b) Ensure a minimum of 1/2 inch clearance between the main landing gear up-position switch wire harnesses and all tubing located in the main landing gear wheel wells. (4) Position the inner fender and install all attach screws (2). (5) If brake deice system is installed, install standoff (14), clamps (9 and 13), washer (10) and screw (11). If no brake deice system installed, install patch plates (1). (6) Install the mount bracket (5) using seven bolts (6). NOTE: Use the dimensions noted during removal when installing the up-position indicator switch (3).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (7) Install up-position indicator switch (3) in the mount bracket (5). Adjust the up-position indicator switch (3) to the dimensions noted during removal, install the jam nut (4) and safety wire (7). (8) Perform the MAIN OR NOSE LANDING GEAR UP-POSITION INDICATOR SWITCH ADJUSTMENT procedure (Ref. Chapter 32-60-00).

2. NACELLE INNER FENDER (WITH KIT NO. 114-9801) A. Removal (1) Perform the REMOVING GROUND POWER procedure (Ref. Chapter 24-40-00). (2) Remove the sealant along the split line of the upper and lower half. (3) Remove the upper half as follows: (a) Remove all attach screws (2) (Ref. Figure 202). (b) If brake deice system is installed, remove screw (11), washer (10), clamps (9 and 13) and standoff (14). If no brake deice system installed, remove the patch plates (1). (c) Remove the upper half of the inner fender from the airplane. (4) Remove the lower half as follows: NOTE: Record the dimension of the up-position indicator switch (3) protruding below the mount bracket (5) for use during installation (Ref. Figure 202). (a) Remove the safety wire (7), remove jam nut (4) and remove up-position indicator switch (3). (b) Remove seven bolts (6) and remove mount bracket (5). (c) Remove all attach screws (15). (d) Remove the grommet (8) and feed the up-position indicator switch (3) and harness through the opening. (e) Remove the lower half of the inner fender from the airplane.

B. Installation (1) Perform the REMOVING GROUND POWER procedure (Ref. Chapter 24-40-00). (2) Install the lower half as follows: (a) Feed the up-position indicator switch (3) and harness through the opening and install the grommet (8) (Ref. Figure 202). (b) Measure the distance between the fuel lines and the main landing gear up-position switch wire harness. If the distance measured is less than 1/2 inch, complete Steps 1 and 2. 1 Using tie straps (114, Table 1, Chapter 91-00-00), remove excess slack from the main landing gear up-position indicator switch wire harnesses by securing the harnesses to existing wire harnesses which route through the top of the main landing gear wheel well side panels.

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54-10-01

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL 2 Ensure a minimum of 1/2 inch clearance between the main landing gear up-position indicator switch wire harnesses and all tubing located in the main landing gear wheel wells. (c) Position the lower half and install all attach screws (15). (d) Install the mount bracket (5) using seven bolts (6). NOTE: Use the dimensions noted during removal when installing the up-position indicator switch (3). (e) Install up-position indicator switch (3) in the mount bracket (5). Adjust the up-position indicator switch (3) to the dimensions noted during removal, install the jam nut (4) and safety wire (7). (3) Install the upper half as follows: (a) Position the upper half and install all attach screws (2) (Ref. Figure 202). (b) If brake deice system is installed, install standoff (14), clamps (9 and 13), washer (10) and screw (11). If no brake deice system installed, install patch plates (1). (4) Install sealant (1, Table 1, 54-00-00) along the split line of the upper and lower half. (5) Perform the MAIN OR NOSE LANDING GEAR UP-POSITION INDICATOR SWITCH ADJUSTMENT procedure (Ref. Chapter 32-60-00).

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1. 2. 3. 4. 5. 6. 7. 8. 9. 10. 11. 12. 13. 14.

PATCH PLATE SCREW UP-POSITION INDICATOR SWITCH JAM NUT MOUNT BRACKET BOLT SAFETY WIRE GROMMET CLAMP WASHER SCREW BLEED AIR SHUTOFF VALVE CLAMP STANDOFF

14 13

9 10 11

12

1

A 2

3 8 4 7 5

6

A

DETAIL VIEW LOOKING AFT

Figure 201 Inner Fender (Without Kit No. 114-9801) Page 204 Aug 1/10

54-10-01

UC54B 100458AA.AI

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

1. 2. 3. 4. 5. 6. 7. 8. 9. 10. 11. 12. 13. 14. 15.

PATCH PLATE SCREW UP-POSITION INDICATOR SWITCH JAM NUT MOUNT BRACKET BOLT SAFETY WIRE GROMMET CLAMP WASHER SCREW BLEED AIR SHUTOFF VALVE CLAMP STANDOFF SCREW

14 13

9 10 11

A 12

1

2

3 8 4 7 5

6 15

A

DETAIL VIEW LOOKING AFT

UC54B 100459AA.AI

Figure 202 Inner Fender (With Kit No. 114-9801)

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NACELLES NACELLE PLATES/SKINS MAINTENANCE PRACTICES

54-30-00 200200

1. PROCEDURES 2. UPPER-AFT NACELLE FAIRING (UC-1 AND AFTER) A. Removal NOTE: Note the position of the AN525-10R8 and AN525-10R9 screws removed from the nacelle fairing panel for installation purposes. (1) Remove the screws (1, 2 and 3) attaching the upper-aft nacelle fairing to the wing skin and forward nacelle fairing (Ref. Figure 201). (2) Inspect the nacelle fairing seal (4) for installation and condition. (3) Inspect the wing skin and nacelle fairing contact area for wear and/or damage. Perform the CHAFING ON WING, TOP SKIN UNDER NACELLE ‘TURTLE BACK’ (UC-1 AND AFTER; UD-1 AND AFTER; UE-1 AND AFTER), Chapter 57-90-04 of the Model 1900 Airliner Series Structural Repair Manual, as required.

B. Installation NOTE: Note the position of the AN525-10R8 and AN525-10R9 screws removed from the nacelle fairing panel for installation purposes. (1) Inspect the nacelle fairing nut plates for installation and condition. (2) Position the upper-aft nacelle fairing on the wing skin. CAUTION: Do not install screws of improper length. Installing improper length screws may result in a fuel leak (Ref. Figure 201). (3) Install AN525-10R8 screws (1) in the forward inboard holes, do not tighten at this time. (4) Install AN525-10R9 screws (2) in the remaining nacelle fairing panel screw holes, do not tighten at this time. (5) Install the countersunk screws (3) in the forward leading edge of the aft nacelle fairing panel. (6) Tighten all the loosely installed screws from Steps (3), (4) and (5).

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1. AN525-10R8 (SHORT SCREWS) 2. AN525-10R9 (LONG SCREWS) 3. COUNTERSUNK SCREWS (19 PLACES) 4. SEAL

A A 3

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Figure 201 Aft Nacelle Fairing Installation Left Shown, Right Opposite Page 202 Nov 1/09

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HORIZONTAL STABILIZER 55-10-00 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Attach Bolts Torque Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 201 202 203

STABILON 55-10-01 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 201 201

TAIL-LET 55-10-02 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 201 201

ELEVATOR 55-20-00 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Balancing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Checking Balance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Balancing Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Check Balancing By Force Measurement Method . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Equipment Required . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Force Measurement Method . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Check Balancing By Counterbalancing Method . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Equipment Required . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Counterbalancing Method . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 201 201 201 201 201 202 202 202 203

VERTICAL STABILIZER 55-30-00 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

55-CONTENTS

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CHAPTER 55 - STABILIZERS TABLE OF CONTENTS (CONTINUED) SUBJECT

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RUDDER 55-40-00 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Balancing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Checking Balance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Balancing Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Equipment Required to Perform Check Balancing by Force Measurement Method . . . . . . . . . . . . . . Force Measurement Method . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Equipment Required to Perform Check Balancing by Counterbalancing Method . . . . . . . . . . . . . . . . Counterbalancing Method . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

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STABILIZERS HORIZONTAL STABILIZER MAINTENANCE PRACTICES

55-10-00 200200

1. PROCEDURES A. Removal (1) Remove the rivets and screws which attach the center fairings at the LH and RH sides of the horizontal and vertical stabilizer intersection (Ref. Figure 201). (2) Remove the access plate (20) from the LH side of the vertical stabilizer, just below the horizontal stabilizer. (3) Disconnect the elevator torque tubes from the elevator bellcrank (Ref. Figure 202). Remove the elevators (13) (Ref. Chapter 27-30-00, ELEVATOR REMOVAL). (4) Remove the access plates (17 and 18) from the top center of the horizontal stabilizer and disconnect the wiring to the rotating beacon (16) and the upper tail fairing navigation light (19). Remove the rotating beacon (Ref. Figure 201). (5) Remove the two access plates (each side) from the upper surface of the LH and RH horizontal stabilizers (21 and 22). Remove the fairing and tail cone. (6) Tag to identify and disconnect the LH and RH elevator trim tab cables from the turnbuckles in the horizontal stabilizer assembly. Remove the cables from the pulleys and secure with tape to prevent them from falling into the vertical stabilizer (5) (Ref. Figure 202). (7) Remove the screws which attach the aft tail fairing (15) to the horizontal stabilizer and remove the fairing (Ref. Figure 201). (8) Drill out the rivets which attach the forward fairing (4) to the vertical and horizontal stabilizers and remove the fairing. (9) Drill out the rivets which attach the fairing angles (25) to the horizontal and vertical stabilizer intersection. (10) Disconnect the surface deicer tubes from the horizontal stabilizer. (11) Remove the 8 bolts (3) which attach the horizontal stabilizer to the vertical stabilizer (Ref. Figure 203). NOTE: When removing the bolts, carefully remove the shims from between the horizontal and vertical stabilizer attach points so they will not be inadvertently dropped into the vertical stabilizer. (12) Remove the horizontal stabilizer assembly from the airplane.

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B. Installation (1) Position the horizontal stabilizer (12) on the vertical stabilizer (5) (Ref. Figure 201) and install the 8 bolts (3) which attach the horizontal stabilizer to the airplane (Ref. Figure 203). NOTE: Before the bolts are tightened, measure the gap at the attach points. If the gap exceeds 0.014-inch, install laminated shims, P/N 101-600012-3 at the forward vertical stabilizer spar and P/N 101-600012-5 at the rear spar as necessary. (2) Torque the forward bolts 205 to 245 inch-pounds. After torquing, the bolts must not rotate under 100 to 160 inch-pounds of torque applied to the bolt. (3) Torque the rear bolts 350 to 430 inch-pounds. After torquing, the bolts must not rotate under 175 to 215 inch-pounds of torque applied to the bolt. (4) Connect the surface deicer tubes to the horizontal stabilizer. (5) Install and rivet the fairing angles (25) to the horizontal and vertical stabilizer intersection (Ref. Figure 201). (6) Install and rivet the forward fairing (4) to the horizontal and vertical stabilizers with Cherrylock rivets. (7) Install and attach the rear fairings (15) to the horizontal and vertical stabilizers with screws. Install the rotating beacon (16). (8) Remove the tape used to secure the elevator trim tab cables. Route the elevator trim tab cables over the pulleys and into the LH and RH horizontal stabilizers and connect them to the turnbuckles. Remove the identification tags (Ref. Figure 202). (9) Connect the wiring to the rotating beacon and the upper tail fairing navigation light (19) (Ref. Figure 201). (10) Install the elevators (13) and connect the elevator horns (Ref. Chapter 27-30-00, ELEVATOR INSTALLATION). (11) Rig the elevators and elevator trim tabs (Ref. Chapter 27-30-00, ELEVATOR CONTROL SYSTEM RIGGING). (12) Install all covers, access plates and the tail cone. (13) Install the rivets and screws which attach the center fairings at the LH and RH sides of the horizontal and vertical stabilizer intersection.

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C. Attach Bolts Torque Check (1) Check the horizontal stabilizer attach bolts. (a) The forward bolts must not rotate under 100 to 160 inch-pounds of torque applied to the bolts. (b) The rear bolts must not rotate under 175 to 215 inch-pounds of torque applied to the bolts. (c) If either of these checks are not satisfactory perform the following: 1 Torque the forward bolts 205 to 245 inch-pounds. After torquing, the bolts must not rotate under 100 to 160 inch-pounds of torque applied to the bolt. 2 Torque the rear bolts 350 to 430 inch-pounds. After torquing, the bolts must not rotate under 175 to 215 inch-pounds of torque applied to the bolt. NOTE: HORIZONTAL STABILIZER Attach Bolts Torque Check procedure must be performed again after 600 hours. 3 If found to be loose remove the affected bolts and inspect the bolt(s) and hole(s) for damage, repair or replace as necessary. If satisfactory continue with the normal inspection schedule (Ref. Chapter 5-20-07).

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Figure 201 Empennage

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Figure 202 Elevator Control System

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Figure 203 Horizontal Stabilizer Page 206 Nov 1/12

55-10-00

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

STABILIZERS STABILON MAINTENANCE PRACTICES

55-10-01 200200

1. PROCEDURES A. Removal (1) Remove the aft pressure bulkhead and the access panels on either side of the empennage below the stabilon (Ref. Figure 201). (2) Disconnect the deice tube. (3) Remove the screws attaching the stabilon angles to the fuselage and the bolts securing the stabilon to the splice plates. (4) Slide the stabilon out of the fuselage.

B. Installation (1) If the replacement stabilon has attach angles, proceed to Step (3), if angles are not installed, proceed with Step (2). (2) Locate, drill and install the upper and lower stabilon attach angles on the stabilon using adjustable preload fasteners. Fastener removal and installation instructions can be found in Chapter 51-40-07 of the Model 1900 Series AIRLINER STRUCTURAL REPAIR MANUAL. NOTE: To determine attach angle location, the stabilon must be properly located on the fuselage. Stabilon location is determined by comparing measurements from the stabilon trailing edge tips to the fuselage centerline at the tailcone when the stabilons are inserted into the fuselage until they are in contact with each other and centered. CAUTION: If the holes in the stabilon are misdrilled, the stabilon is not repairable. New angles or splice plates may be matched to an existing stabilon or a new stabilon may be matched to existing angles and splice plates. (3) Slide the replacement stabilon into position in the fuselage (Ref. Figure 201). (4) Locate and mark the hole positions on the attach angles and the stabilon mounting area at the splice plates. (5) Remove the stabilon and drill holes. The attach angle to fuselage mounting holes should be 0.193 to 0.200-inch in diameter. NOTE: Once the holes are drilled, the attach angles, stabilon and splice plates are matched to each other and may not fit another installation.

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(6) Connect the deice tube. (7) Install the access panels and pressure bulkhead. Example: Locknut running torque = 15 inch-pounds, determined by observing the torque meter while turning the nut on the bolt with the locking feature engaged. 15 inch-pounds + 30 inch-pounds = 45 inch-pounds minimum. 15 inch-pounds + 40 inch-pounds = 55 inch-pounds maximum.

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55-10-01

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Figure 201 Stabilon Removal and Installation

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STABILIZERS TAIL-LET MAINTENANCE PRACTICES

55-10-02 200200

1. PROCEDURES A. Removal (1) Remove the fasteners (Ref. Figure 201). Fastener removal and installation instructions can be found in Chapter 51-40-03 of the Model 1900 SERIES AIRLINER STRUCTURAL REPAIR MANUAL. (2) Remove the tail-let by sliding it down and tilting it outboard. (3) Remove the access panel on top of the stabilizer and remove the rivet shanks, metal chips and any other debris.

B. Installation (1) If the replacement tail-let has attach angles, proceed to Step (3), if the angles are not installed, proceed to Step (2). (2) Locate, drill, and install the four attach angles on the tail-let using adjustable preload fasteners. Fastener removal and installation instructions can be found in Chapter 51-40-07 of the Model 1900 SERIES AIRLINER STRUCTURAL REPAIR MANUAL. NOTE: To determine attach angle location, the tail-let must be properly located in the stabilizer. Tail-let location is determined by inserting the tail-let into the stabilizer until it contacts the lower stabilizer skin and aligning the top brackets with the existing rivet holes in the stabilizer. CAUTION: If the holes in the tail-let are misdrilled, the tail-let is not repairable. New angles may be matched to an existing tail-let or a new tail-let may be matched to existing angles. (3) Place the tail-let into position in the stabilizer and drill the attach angles to match the existing stabilizer structure (Ref. Figure 201). NOTE: Once the holes are drilled, the attach angles, tail-let and stabilizer are matched to each other and may not fit another installation. (4) Install new fasteners to replace those removed during tail-let removal (Ref. Figure 201). (5) Install the access plate on the top of the stabilizer. (6) Seal the tail-let to the stabilizer with sealer (38, Table 1, Chapter 91-00-00) or equivalent.

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Figure 201 Tail-Let Removal and Installation

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STABILIZERS ELEVATOR MAINTENANCE PRACTICES

55-20-00 200200

1. PROCEDURES A. Balancing After repainting and/or repair, the finished elevator must be check balanced to ensure that its static moment about the hinge line is within the prescribed limits. The static moment for all completed elevator assemblies must fall within the range of 1.60 inch-pounds nose heavy to 0.80 inch-pounds tail heavy at the measured moment about the hinge line (Ref. Figure 201). The static moment of the elevator is determined by multiplying the unbalanced weight of the elevator assembly times the perpendicular distance from the hinge centerline to the center of gravity when the chord line is horizontally level. The weight is measured in pounds and the distance in inches. The static moment of a 100 percent balanced elevator assembly is 0.0 inch-pounds. Tail heaviness indicates static underbalance while nose heaviness indicates static overbalance.

B. Checking Balance The balance must be checked in a draft-free area with the elevator completely assembled in flying condition. All painting, including stripes and touch-up, must be completed. The tab, tab pushrod, static wicks, and hinge bolts must be attached. The chord line must be horizontally level and the hinge line must be properly supported when the static moment is measured. Although many different methods of check balancing exist, they can be categorized under the following two headings: (1) Counterbalancing - The application of a known force or weight at a measured distance from the hinge line to counter the unbalance moment of the elevator assembly. (2) Actual Force Measurement - Measurement of the force applied by the elevator surface on a single support at a known distance from the centerline of the hinge. NOTE: Counterbalancing is the simplest method of check balancing.

C. Balancing Procedures WARNING: Airplanes using adjustment weight attachment screws must replace them with adjustment weight attachment bolts while performing these procedures (Ref. Figure 201). Refer to Table 201 for appropriate adjustment weight attachment bolts.

2. CHECK BALANCING BY FORCE MEASUREMENT METHOD A. Equipment Required (1) A stand with knife-edge supports (Ref. Figure 201). The knife edges should be in the same horizontal plane. (2) A certified beam balance calibrated in units of 0.01 pounds or less. The balance should have a flat weighing platform and its capacity should equal tare plus 2.0 pounds minimum. (3) A support spindle similar to the illustration and leveling blocks, as required (blocks + spindle = tare).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (4) A straightedge, rule and spirit level.

B. Force Measurement Method (1) Locate the chord line by placing a straightedge at the inboard end of the elevator so that one end is aligned with the center of the torque tube and the other end is centered on the trailing edge. Mark the chord line by grease pencil or other means on the rib. Remove the straightedge. (2) Fit correct size bolts in the outboard and center hinge brackets and mount the elevator on the knife edges. Ensure that it is free to rotate about the hinge line. (3) Place a small platform scale under the trailing edge if the control surface is tail-heavy or under the leading edge if it is nose-heavy. (4) Place the upper end of the spindle under the trailing edge of tail-heavy surfaces or under the leading edge of the nose-heavy surfaces. The spindle must be vertical throughout the balancing procedure. Hold a spirit level against the marked chord line and level it by extending or contracting the spindle, or by using blocks and shims under the spindle. (5) Measure the perpendicular distance from the hinge centerline to the point supported by the spindle. Ensure that the spirit level and rule are removed from the surface and read the reaction on the beam balance. (6) Calculate the static underbalance moment M from the formula: M = D (R-T) inch-pounds where, D = Perpendicular distance from the hinge center line to the spindle point (inches). R = Reaction (pounds) read from the beam balance. T = Tare, i.e. spindle plus leveling blocks or shims on the scale platform (pounds). EXAMPLE D is 10.0 inches, R is 1.100 lbs and T = 1.000 lbs M = 10.0 (1.100-1.000); M = 1.00 inch-pounds. M is within the prescribed range which is satisfactory. If M is not within the prescribed range, refer to Step (9) of the COUNTERBALANCING METHOD procedures, in this section.

3. CHECK BALANCING BY COUNTERBALANCING METHOD A. Equipment Required (1) A stand with knife-edge supports (Ref. Figure 201). The knife edges should be in the same horizontal plane. (2) A paper cup or similar lightweight container. (3) Approximately 0.5 pound of lead shot. (4) A certified beam balance weighing device calibrated in units of 0.01 pound or less. (5) A straightedge, ruler, and spirit level.

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B. Counterbalancing Method (1) Locate the chord line by placing a straightedge at the inboard end of the elevator assembly so that one end is on the hinge center line and the other end is centered on the trailing edge. Mark the chord line with a suitable marker, such as a grease pencil, then remove the straightedge. NOTE: While the hinge center line is not openly visible, it can be located as follows: With the hinge clevis properly aligned and tightened, locate the center of the clevis mounting bolt on the top of the elevator (covered with a plug button) as a reference point. The hinge line is 1.13 inches forward of this reference point. Mark the hinge line with a grease pencil. (2) Secure the trim tab in its neutral position with a small piece of masking tape. (3) Fit the correct sized bolts in the hinge clevises and mount the elevator on the knife-edge supports (Ref. Figure 201). Ascertain that the elevator is free to rotate about the hinge line. (4) To determine if adjustment weights should be added or removed, suspend a paper cup from a point near the center of the elevator trailing edge if the balance is nose-down or near the inboard end of the balance weight assembly on the elevator leading edge if the balance is tail-down. Use a short length of small diameter string secured to the surface with a small piece of masking tape (Ref. Figure 201). The cup must be free to hang vertically. (5) Add small quantities of lead shot to the cup until the elevator balances with the chord line level. Check this by holding the spirit level aligned with the marked chord line. (6) Carefully measure the perpendicular distance D within 0.1-inch from the hinge centerline to the point of suspension of the cup. (7) Remove the cup, contents, and string, then weigh them to within 0.05-pound. NOTE: Since any weighing error is magnified by the distance D weighing is most important and must be done carefully on scales that are certified for accuracy. (8) Calculate the static balance as follows: (a) The weight of the cup and contents is designated by W. (b) The distance between the center line of the hinge and the suspension point of the cup is designated by D. (c) The over- or underbalance moment is designated by M. (d) M = W x D (e) The following is a typical example of a balancing calculation: Assume the elevator is slightly overbalance (nose-heavy) and the paper cup was suspended from the trailing edge. If the elevator balances with the chord line level at W = 0.100-pound and D = 13.8-inches, then... M = 0.100 x 13.8; M = 1.38 inch-pounds (The product of W x D must be accurate to within 0.05 inch-pound). In this instance M is within the required static balance range and is therefore acceptable. (9) If the static balance does not fall within the range of 1.60 inch-pound nose-heavy (overbalance) to 0.80 inch-pound tail-heavy (underbalance), adjustment weights must be added or removed and the balance rechecked.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL NOTE: A maximum of four small adjustment weights (P/N 101-610026-5) and seven large adjustment weights (P/N 101-610026-7) may be added or removed from the balance weight assembly on the outboard leading edge of the elevator. The steel cover plate must be installed over the adjustment weights with NAS6703HUXX bolts (Ref. SB 27-3187). Bolt lengths vary with the number of weights installed. Refer to Table 201 for the proper attachment bolts. Nut plates are located inside the elevator assemblies to anchor the bolts. Bolts are torqued 20 to 30 inch-pounds then safety wired with MS20995C32 wire. Table 201 ADJUSTMENT WEIGHTS ATTACHMENT BOLTS (AIRPLANES THAT HAVE COMPLIED WITH SB 27-3187) No. of Weights Req. for Balancing

Combination Weights Req. P/ N 101-610026-5 (Small Weight)

Combination Weights Req. P/ N 101-610026-7 (Large Weight)

Number of Bolts Req. Per Elevator

Number of Washers Req. Per Bolt

Part Number of Bolts for Each Elevator

1

1

0

2

0

NAS6703HU12

2

2

0

2

0

NAS6703HU13

2

2

0

2

1

NAS6703HU14

3

3

0

2

0

NAS6703HU14

4

4

0

2

0

NAS6703HU15

4

4

0

2

1

NAS6703HU16

5

4

1

2

0

NAS6703HU16

6

4

2

2

0

NAS6703HU17

6

4

2

2

1

NAS6703HU18

7

4

3

2

0

NAS6703HU18

8

4

4

2

0

NAS6703HU19

8

4

4

2

1

NAS6703HU20

9

4

5

2

0

NAS6703HU20

10

4

6

2

0

NAS6703HU21

10

4

6

2

1

NAS6703HU22

11

4

7

2

0

NAS6703HU22

NOTE: Based on availability it is permissible to use the next grip length bolt and add one P/N NAS 1149C0363R washer under the head of the attachment bolts, instead of using the odd part numbered bolts.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 201 Elevator Balancing

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STABILIZERS VERTICAL STABILIZER MAINTENANCE PRACTICES

55-30-00 200200

1. PROCEDURES A. Removal (1) Remove the horizontal stabilizer (14) (Ref. Figure 201) (Ref. HORIZONTAL STABILIZER REMOVAL, 55-10-00). (2) Remove the rudder (15) (Ref. RUDDER REMOVAL, Chapter 27-20-00). (3) Install identification tags on the rudder cables (2) and disconnect the rudder cables from the turnbuckles (1) in the tail section (Ref. Figure 202). (4) Remove the rudder pushrod (4) from the control horn (3) and the torque shaft assembly (5) (Ref. Figure 202). (5) Drill out the rivets (9) which attach the short aft dorsal fin section (10) to the main dorsal fin assembly (16) and the aft fuselage. Remove the short aft dorsal fin (10) from the airplane (Ref. Figure 201). (6) Drill out the rivets (12) which attach the fuselage/vertical stabilizer fillet (11) to the fuselage. (7) Remove the bolts (1) and countersunk washers (17) attaching the vertical stabilizer at the forward and rear attaching points. (8) Carefully lift the vertical stabilizer assembly (13) from the fuselage.

B. Installation (1) Position the vertical stabilizer (13) on the aft fuselage and install the attaching bolts (1), countersunk washers (17), washers (2), and nuts (3) through the main and rear spars and the fuselage structure (Ref. Figure 201). Shim the attach points as necessary with laminated shims, (4, 5, 6 and 7) P/N 101-600012-1 to a maximum gap of 0.020 inch between the spars and their respective attach points. (2) Torque the attaching bolts (1) 100 to 140 inch-pounds. (3) Attach the vertical stabilizer fillet (11) to the fuselage with rivets (12). (4) Position the small dorsal fin section (10) and rivet (9) to the main dorsal fin assembly (16) and the aft fuselage. Be certain the seal between the aft dorsal fin and vertical stabilizer is in position. (5) Connect the rudder pushrod (4) to the control horn (3) and the torque shaft assembly (5) (Ref. Figure 202).

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(6) Connect the rudder cables (2) to the turnbuckles (1) and remove the identification tags (Ref. Figure 202). (7) Install the rudder (15) (Ref. Figure 201) (Ref. RUDDER INSTALLATION, Chapter 27-20-00). (8) Install the horizontal stabilizer (14) (Ref. HORIZONTAL STABILIZER INSTALLATION, 55-10-00).

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Figure 201 Vertical Stabilizer Removal and Installation

55-30-00

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 202 Vertical Stabilizer

Page 204 Nov 1/09

55-30-00

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

STABILIZERS RUDDER MAINTENANCE PRACTICES

55-40-00 200200

1. PROCEDURES A. Balancing When the rudder surface is being repainted it is recommended to position it by utilizing best shop practice so that excess paint will drain toward the leading edge. This will minimize the amount of balance weights required. After repainting and/or repair, the finished rudder must be check balanced to ensure that its static moment about the hinge line is within the prescribed limits. The static moment for all completed rudder assemblies must fall within the range of 3.55 pound-inches to 9.55 pound-inches tail-heavy at the measured moment about the hinge line. The static moment of the rudder is determined by multiplying the unbalanced weight of the rudder assembly times the perpendicular distance from the hinge centerline to the center of gravity when the chord line is horizontally level (Ref. Figure 201). The weight is measured in pounds and the distance in inches. The static moment of a 100 percent balanced rudder assembly is 0.0 pound-inches. Tail heaviness indicates static underbalance while nose heaviness indicates static overbalance.

B. Checking Balance The rudder balance must be checked in a draft free area with the rudder completely assembled in flying condition. All painting, including stripes and touch-up, must be completed. The tab, tab pushrod, static wicks, hinge clevises, bonding jumpers and hardware, must be attached. The chord line must be horizontally level and the hinge line must be properly supported when the static moment is measured. Although many different methods of check balancing exist, they can be categorized under the following two headings: (1) Counterbalancing - The application of a known force or weight at a measured distance from the hinge line to counter the unbalance moment of the rudder assembly. (2) Actual Force Measurement - Measurement of the force applied by the rudder surface on a single support at a known distance from the centerline of the hinge. NOTE: Counterbalancing is the simplest method of check balancing.

2. BALANCING PROCEDURES A. Equipment Required to Perform Check Balancing by Force Measurement Method (1) A stand with knife-edge supports (Ref. Figure 201). The knife edges should be in the same horizontal plane. (2) A certified beam balance calibrated in units of 0.01 pounds or less. The balance should have a flat weighing platform and its capacity should equal tare plus 2.0 pounds minimum. (3) A support spindle similar to the illustration and leveling blocks, as required (blocks + spindles = tare).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (4) A straightedge, rule and spirit level.

B. Force Measurement Method (1) Locate the chord line by placing a straightedge at the inboard end of the rudder so that one end is aligned with the center of the torque tube and the other end is centered on the trailing edge. Mark the chord line by grease pencil or other means on the rib. Remove the straightedge. (2) Fit correct size bolts in the outboard and center hinge brackets and mount the rudder on the knife edges. Ensure that it is free to rotate about the hinge line. (3) Place a small platform scale under the trailing edge if the control surface is tail-heavy or under the leading edge if it is nose-heavy. (4) Place the upper end of the spindle under the trailing edge of tail-heavy surfaces or under the leading edge of the nose-heavy surfaces. The spindle must be vertical throughout the balancing procedure. Hold a spirit level against the marked chord line and level it by extending or contracting the spindle, or by using blocks and shims under the spindle. (5) Measure the perpendicular distance from the hinge centerline to the point supported by the spindle. Ensure that the spirit level and rule are removed from the surface and read the reaction on the beam balance. (6) Calculate the static underbalance moment M from the formula: M = D (R-T) inch-pounds where, D = Perpendicular distance from the hinge center line to the spindle point (inches). R = Reaction (pounds) read from the beam balance. T = Tare, i.e. spindle plus leveling blocks or shims on the scale platform (pounds). EXAMPLE D is 13.8 inches, R is 1.490 lbs and T = 1.000 lbs M = 13.8 (1.490-1.000); M = 6.76 inch-pounds. M is within the prescribed range which is satisfactory. The complete painted rudder assembly with the tab, tab pushrod, static wicks, hinge clevises, bonding jumper and hardware, must fall within the range of 3.55 pound-inches to 9.55 pound-inches tail heavy. If the static balance is not as noted, adjustment weights must be added or removed and the balance rechecked to bring the rudder balance within the required limits.

C. Equipment Required to Perform Check Balancing by Counterbalancing Method (1) A stand with knife-edge supports (Ref. Figure 201). The knife edges should be in the same horizontal plane. (2) A paper cup or similar lightweight container. (3) Approximately 2 pound of lead shot. (4) A certified beam balance weighing device calibrated in units of 0.01 pound or less. (5) A straightedge, ruler, and spirit level.

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D. Counterbalancing Method (1) Locate the chord line by placing a straightedge at the lower closure rib of the rudder so that one end is aligned with the center of the torque tube while the other end is centered on the trailing edge. Mark the chord line with a suitable marker, such as a grease pencil, then remove the straight edge. (2) Secure the trim tab in its neutral position with a small piece of masking tape. (3) Fit the correct sized bolts in the hinge brackets and mount the rudder on the knife-edge supports. Ascertain that the elevator is free to rotate about the hinge line. (4) To determine if adjustment weights should be added or removed, suspend a paper cup from a point near the center of the leading edge. Use a short length of small diameter string secured to the surface with a small piece of masking tape (Ref. Figure 201). The cup must be free to hang vertically. (5) Add small quantities of lead shot to the cup until the rudder balances with the chord line level. Check this by holding the spirit level aligned with the marked chord line. (6) Carefully measure the perpendicular distance D within 0.1 inch from the hinge centerline to the point of suspension of the cup. (7) Remove the cup, contents, and string, then weigh them to within 0.05-pound. NOTE: Since any weighing error is magnified by the distance D, weighing is most important and must be done carefully on scales that are certified for accuracy. (8) Calculate the static balance as follows: (a) The weight of the cup and contents is designated by W. (b) The distance between the center line of the hinge and the suspension point of the cup is designated by D. (c) The over or underbalance moment is designated by M. (d) M = W x D The following is a typical example of a balancing calculation: Assume the rudder was slightly underbalance (tail heavy) and the paper cup was suspended from the leading edge. If the rudder balances with the chord line level at W = 0.650 pounds and D = 12.5 inches, then... M = 0.650 x 12.5 M = 8.12 pound-inches In this instance, M is within the required static balance range and is therefore acceptable. The complete painted rudder assembly with the tab, tab pushrod, static wicks, hinge clevises, bonding jumper and hardware, must fall within the range of 3.55 pound-inches to 9.55 pound-inches tailheavy. If the static balance is not as noted, adjustment weights must be added or removed and the balance rechecked to bring the rudder balance within the required limits.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL NOTE: Install the P/N 101-630017-1 adjustment weights as necessary to obtain the required balance. Use a maximum of 7 at the root rib and a maximum of 7 at Cant Station 34.59. One P/N 114-630020-7 weight must be installed between the head of the bolt and stack of 101-630017-1 adjustment weights at each location. Nut plates are located on the rib to anchor the attaching screws. Refer to Table 201 for screw sizes. Table 201 ADJUSTMENT WEIGHTS ATTACHMENT SCREWS Number of Weights Necessary for Balancing

Number of Screws Required

Part Number of Screws for Each Installation

0

0

----

1 or 2

2

AN3H4A

3 or 4

2

AN3H5A

5 or 6

2

AN3H6A

7 or 8

2

AN3H7A

All weights are identical. Read across from the left column to determine the number and type of screws for the installation.

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55-40-00

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 201 Rudder Balancing

55-40-00

Page 205 Nov 1/09

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

CHAPTER 56 - WINDOWS TABLE OF CONTENTS SUBJECT

PAGE

WINDOWS 56-00-00 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Plastic Windows . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Cleaning . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Windshields . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Cleaning . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Windows . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removing Paint . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Window Inspection and Repair Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Flight Compartment, Cabin, and Baggage Windows . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Acrylic Window Refurbishing Procedure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 201 201 201 201 202 202 202 202 202

FLIGHT COMPARTMENT 56-10-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Storm Window Seals . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Windshield Surface Seal Coating . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Windshield . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Windshield Weather Seal Inspection and Replacement . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 206 Seal Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 206 Silicone Seal Repair . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 206 Temporary Hump Seal Repair (UB-1 thru UB-74 and UC-1 thru UC-174) . . . . . . . . . . . . . . . . . . . . . . 207 Hump Seal Repair/Replacement . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 207 Windshield . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 208 Replacement Criteria . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 208 Antistatic Tab . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 210 Installation (Aluminum Frames) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 210 Windshield Antistatic Coating and Tab . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 213 Inspection (Aluminum Frames) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 213 Coating Inspection (Lightweight Fiberglass Frame) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 214 Bonded Antistatic Tab . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 215 Repair (Aluminum Frames) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 215 Flight Compartment Side Window . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 217 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 217 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 217 Window Attach Frames . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 218 Inspection and Repair . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 218 Storm Window . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 220 Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 220 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 220 Installation and Adjustment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 221 Storm Window Primary Seal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 221 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 221 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 221 Storm Window Secondary Seal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 222

56-CONTENTS

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CHAPTER 56 - WINDOWS TABLE OF CONTENTS (CONTINUED) SUBJECT PAGE Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 222 Installation (Optional) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 222 Storm Window Stop . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 222 Adjustment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 222

CABIN 56-20-00 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Cabin Window . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

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List of Effective Pages CH-SE-SU

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DATE

56-LOEP

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201 thru 208

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56-10-00

1 201 thru 223

Nov 1/09 Nov 1/13

56-20-00

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WINDOWS MAINTENANCE PRACTICES

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1. PLASTIC WINDOWS A. Cleaning The plastic windows should be kept clean and waxed at all times. To prevent scratches and crazing, wash the windows carefully with plenty of soap and running water. CAUTION: When washing the windows, do not use water from a bucket or pail. Sand, dirt particles or other debris may collect in the standing water and cause scratches in the plastic. Use the palm of the hand to feel and dislodge dirt and mud. A soft cloth, chamois or sponge may be used only for the purpose of carrying water to the surface of the window. After washing, rinse the window thoroughly with running water and dry it with a clean, moist chamois. Do not rub the plastic window with a dry cloth, because this will cause an electrostatic charge which attracts dust. Remove oil and grease with a cloth moistened with kerosene (49, Table 1, Chapter 91-00-00), solvent (54, Table 1, Chapter 91-00-00) or hexane (51, Table 1, Chapter 91-00-00), then rinse the window with clear water. CAUTION: Never use gasoline, benzene, alcohol, acetone, carbon tetrachloride, fire extinguisher or anti-ice fluid, lacquer thinner, or glass cleaner with a base of these materials, for such materials will soften the plastic and may cause crazing. Aliphatic naphtha and similar solvents are highly flammable and extreme care must be exercised when using these chemicals. If using a commercial cleaner to clean the plastic windows, use only cleaners that are approved by Hawker Beechcraft Corporation. There are several cleaners available commercially that state that they are approved for use on acrylic surfaces; however, it has been discovered that some of these cleaners cause acrylic plastic to craze. Only the following product is approved as a cleaner for acrylic plastic windows: Federal Specification P-P-560, Tend Plastic Cleaner and Polish, Loctite plastic cleaner P/N 30559 or Clearview plastic cleaner P/N AVL-CV-16 (48, Table 1, Chapter 91-00-00). Follow the directions on the container. After washing plastic windows with soap and water, apply a good grade of commercial wax. The wax will fill in minor scratches and help prevent further scratches. Apply a thin, even coat of wax and bring it to a high polish by rubbing lightly with a clean, dry, soft flannel cloth. Never use a power buffer, as the heat generated by the buffing pad may soften the plastic. If the windows were cleaned with the commercial cleaner mentioned previously, it will not be necessary to apply wax. This cleaner contains wax, as well as cleaning agents.

2. WINDSHIELDS A. Cleaning Glass windshields with antistatic coating should be cleaned as follows: CAUTION: When washing the windshield, do not use water from a bucket or pail. Sand, dirt particles or other debris may collect in the standing water and cause scratches in the glass. (1) Wash excessive dirt and other substances from the glass with clean water.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (2) Clean the windshield with mild soap and water or with a 50/50 solution of solvent (30, Table 1, Chapter 91-00-00) and water. Wipe the glass surface in a straight rubbing motion with a soft cloth or sponge. Never use any abrasive materials or any strong acids or bases to clean the glass. (3) Rinse the glass thoroughly and dry, but do not apply wax.

3. WINDOWS A. Removing Paint (1) Mask the acrylic crew compartment side windows and the paint on the airplane skin to prevent damage from solvent. CAUTION: The acrylic side windows are extremely susceptible to damage from solvents. To prevent damage, use care not to contact the acrylic windows, or the airplane paint with the methyl propyl ketone or other solvent. To avoid damage to the window anti-static coating, never use any rubbing compound or polish, or any substance containing acid, on the windshield. (2) If the paint on the window is dry, contact Hawker Beechcraft Corporation Technical Support for consultation. (3) If the paint on the window is still wet, use a clean, soft cloth dampened with solvent (14, Table 1, Chapter 91-00-00) to remove the paint. Wipe the glass surface in a straight rubbing motion. Never use any abrasive materials or acids to remove the paint. (4) Rinse the glass thoroughly and dry, but do not apply wax.

4. WINDOW INSPECTION AND REPAIR PROCEDURES A. Flight Compartment, Cabin, and Baggage Windows The two most forward cabin windows of the airplane are made of a fail safe, multi-ply stretched acrylic material. Inspect these windows for deep scratches, chips, excessive crazing, and all other evidence of damage (Ref. Figures 201 and 202). Refer to Table 201 for damage types and recommended action. A preflight inspection should be conducted by the pilot to disclose any possible condition that could warrant further inspection as noted above. The remaining cabin windows and the flight compartment side and storm windows are made of a single-ply stretched acrylic material. These windows are life limited. Refer to 1900 Airliner Series Airworthiness Limitations Manual P/N 129-590000-133, Chapter 04-00-00. Inspect the windows for deep scratches, chips, excessive crazing and all other evidence of damage (Ref. Figures 201 and 202). Refer to Table 202 for types of damage and recommended action. A preflight inspection should be conducted by the pilot to disclose any possible condition that could warrant further inspection as previously noted.

B. Acrylic Window Refurbishing Procedure Table 201 and Table 202 outline the types and limits of window damage which may be repaired utilizing this procedure. If any unusual conditions are found that are not covered in Table 201 and Table 202, contact the Technical Support Department of Hawker Beechcraft Corporation for consultation. Use of this procedure is prohibited except as specifically instructed by Table 201 and Table 202 or by the Technical Support Department of Hawker Beechcraft Corporation.

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NOTE: Only experienced personnel should perform repair work on Plexiglas windows. The flight compartment side windows may be reduced by a total of 0.015 inch from the original thickness of 0.312 inch. The single-ply cabin side windows may be reduced by a total of 0.015 inch from the original thickness of 0.188 inch. The multi-ply cabin side windows may be reduced by a total of 0.015 inch from the original thickness 0.195 inch. (1) Mask off any major portions of the window which are not involved in the refurbishing. (2) Use wet abrasives only. (3) Flush the window surface with a liberal amount of water after using each abrasive grade. (4) Use a straight-line sanding motion and alternate the line of sanding at right angles to cross the sanding pattern. (5) Gradually lighten the working pressure on the abrasive as each finer grade is utilized. Do not exert excessive force on the window pane. (6) Work a marginally larger area as each finer abrasive is used. Completely remove each previous sanding pattern before moving to the next finer abrasive. (7) Do not overheat the acrylic window. NOTE: Depending on the depth of the scratch, the repair may start at any of the following Steps. (8) Start the refurbishing procedure with 240 grit wet/dry silicon carbide cloth or equivalent and an orbital sander. Use this combination until all traces of the damage have been removed. (9) Use a 320 grit wet/dry silicon carbide cloth with the orbital sander and work the window until the 240 grit pattern has been removed. (10) Repeat Step (9) using 400 grit, 600 grit, 1800 grit, 2400 grit, 3200 grit, 4000 grit, and 6000 grit silicon carbide cloth in succession. (11) Finish the refurbishing procedure using an 8000 grit cushioned abrasive cloth or cerium oxide. (12) Inspect the finished window to ensure that all damage and grit patterns have been removed (Ref. Figure 201). (13) When satisfied that the window has been properly refurbished, apply a small amount of antistatic cleaner or wax to the window and buff to a high luster. For removal of minor crazing or scratches, the Plexiglas scratch remover (50, Table 1, Chapter 91-00-00) is recommended.

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Figure 201 Viewing Angles and Light Placement for Window Inspection

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Table 201 MULTI-PLY WINDOW DAMAGE DAMAGE TYPE Scratches, abrasions, gouges and chips with a maximum depth of 0.015 inch.

Scratches, abrasions, gouges and chips with a depth in excess of 0.015 inch.

Crazing with a maximum depth of 0.015 inch.

Crazing with a maximum depth in excess of 0.015 inch.

Haziness and cloudiness.

LOCATION

PROBABLE CAUSE

RECOMMENDED ACTION

Inner ply

Improper maintenance or cleaning procedures.

Airplane MUST be operated unpressurized until the window is replaced.

Outer ply

Improper maintenance, cleaning procedures or object impact.

Damage may be worked out. Refer to ACRYLIC WINDOW REFURBISHING PROCEDURE.

Inner ply

Improper maintenance or cleaning procedures.

Immediate replacement of the window is mandatory before further flight.

Outer ply

Improper maintenance, cleaning procedures or object impact.

Airplane MUST be operated unpressurized until the window is replaced.

Inner ply

Contact with unapproved cleaners, solvents or chemical compounds; stress fatigue.

Airplane MUST be operated unpressurized until the window is replaced.

Outer ply

Contact with unapproved cleaners, solvents, or chemical compounds; stress fatigue.

Damage may be worked out. Refer to ACRYLIC WINDOW REFURBISHING PROCEDURE.

Inner ply

Contact with unapproved cleaners, solvents, or chemical compounds; stress fatigue.

Immediate replacement of the window is mandatory before further flight.

Outer ply

Contact with unapproved cleaners, solvents or chemical compounds; stress fatigue.

Airplane MUST be operated unpressurized until the window is replaced.

Inner ply

Contact with unapproved cleaners, solvents, or chemical compounds; stress fatigue.

Airplane MUST be operated unpressurized until the window is replaced.

Outer ply

Contact with unapproved cleaners, solvents or chemical compounds; stress fatigue.

Airplane MUST be operated unpressurized until the window is replaced.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Table 201 MULTI-PLY WINDOW DAMAGE (Continued) DAMAGE TYPE

LOCATION

PROBABLE CAUSE

RECOMMENDED ACTION

Pane delamination.

Stress; extreme heat at high altitudes.

Airplane MUST be operated unpressurized until the window is replaced.

Out of contour.

Excessive heat or pressurization, high humidity, water absorption, chemical or solvent absorption.

If window is out of contour beyond limits specified by detail A-A or detail B-B in Figure 202, immediate replacement is mandatory before further flight.

Distortion.

Improper restoration when removing damage from outer ply; chemical or solvent damage.

If distortion is caused by poor restoration techniques, the window may be replaced at will with no flight limitations. However, if the distortion is the result of chemical or solvent damage, the airplane MUST be operated unpressurized until the window is replaced.

Inner ply

Improper maintenance; stress fatigue.

Airplane MUST be operated unpressurized until the window is replaced.

Outer ply

Improper maintenance; stress fatigue; object impact.

Airplane may be operated for a maximum of 20 hours at a reduced pressure of 4.6 psid with a maximum altitude of 25,000 feet.

Inner ply

Improper maintenance; stress fatigue.

Airplane MUST be operated unpressurized until the window is replaced.

Outer ply

Improper maintenance; stress fatigue; object impact.

Airplane MUST be operated unpressurized until the window is replaced.

Circumferential cracks, at least 2 inches from the window frame.

Circumferential cracks within 2 inches of the window frame.

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Figure 202 Window Contour Check

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Table 202 SINGLE-PLY WINDOW DAMAGE DAMAGE TYPE

PROBABLE CAUSE

RECOMMENDED ACTION

Scratches, abrasions, gouges and chips with a maximum depth of 0.015 inch.

Improper maintenance, cleaning procedures or object impact.

Damage may be worked out. Refer to ACRYLIC WINDOW REFURBISHING PROCEDURES. (Caution: Repair may cause optical distortion.)

Scratches, gouges and chips with a maximum depth in excess of 0.015 inch.

Improper maintenance, cleaning procedures or object impact.

Airplane MUST be operated unpressurized until the window is replaced.

Crazing with a maximum depth of 0.015 inch.

Contact with unapproved cleaners, solvents or chemical compounds; stress fatigue.

Damage may be worked out. Refer to ACRYLIC WINDOW REFURBISHING PROCEDURES.

Crazing with a maximum depth in excess of 0.015 inch.

Contact with unapproved cleaners, solvent or chemical compounds; stress fatigue.

Immediate replacement of the window is mandatory before further flight.

Haziness and cloudiness.

Improper maintenance or cleaning procedures; contact with unapproved cleaners, solvents or chemical compounds; stress fatigue.

Properly clean window; inspect and check for scratches, abrasion or crazing. If present, follow the preceding action. If the result of a chemical attack, the airplane MUST be operated unpressurized until the window is replaced.

Out of contour.

Excessive heat or pressurization, high humidity, water absorption, chemical or solvent absorption.

If window is out of contour beyond limits specified by detail A-A or detail B-B in Figure 2 immediate replacement of the window is mandatory before further flight.

Distortion (Flight Compartment).

Improper restoration methods; chemical or solvent damage.

Immediate replacement of the window is mandatory before further flight.

Distortion (Cabin).

Improper restoration methods; chemical or solvent damage.

If distortion is caused by poor restoration techniques, the window may be replaced at will with no flight limitations. If distortion is the result of chemical or solvent damage, the airplane MUST be operated unpressurized until the window is replaced.

Cracks through the thickness.

Improper maintenance; stress fatigue, object impact.

Immediate replacement of the window is mandatory before further flight.

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WINDOWS FLIGHT COMPARTMENT DESCRIPTION AND OPERATION

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1. GENERAL A. Storm Window Seals The primary seal for the storm window is installed onto the inner surface of the fuselage pan assembly. The seal performs two functions: it keeps the cabin pressure from escaping during flight and prevents water from entering the cockpit when the airplane is parked. A newer style storm window includes a secondary seal. This secondary seal is installed on the outer surface of the window and mates up against the primary seal. The secondary seal is designed to increase contact pressure and friction on the primary seal, providing improved noise reduction and leak prevention. This secondary seal can be installed to existing storm windows as required. Although the secondary seal is designed to improve the sealing of the storm window, it is possible that in some installations the secondary seal may actually worsen the seal or increase noise. Because of this, the secondary seal should be considered optional and may be removed if desired.

B. Windshield Surface Seal Coating PPG offers the Surface Seal© Coating System, which includes customized kits for coating application, refurbishing and efficiency measurement. The DSS1040 Master Kits contain reusable equipment, replenishable supplies for approximately 10 windshields and chemicals to coat 2 windshields. The DSS1015 Application Kits (Kit A) contains supplies and chemicals to coat 1 windshield. The DSS1027 Master Kit Refill contains replenishable supplies to prepare approximately 10 windshields, less the Surface Seal© preparation and coating solutions. The DSS2000 and DSS2999 Curing Kits contain protective films and customized heating blankets. All part number 101-384025-17/-18 and subsequent windshields have the Surface Seal© Coating System protection applied. For inspection and application refer to the Model 1900 Airliner Series Component Maintenance Manual, Chapter 56.

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WINDOWS FLIGHT COMPARTMENT MAINTENANCE PRACTICES

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1. WINDSHIELD A. Removal (1) Cut the sealing compound free from the retainer strip and window frame. (2) Remove the screws at the front of the windshield. (3) Inside the flight compartment, remove the screw that secures the ground jumper wire at the outboard corner of the windshield and detach the power supply lead from the electrothermal windshield at the corner post. Gently push the windshield panels outward one at a time. NOTE: During removal/cleaning of fiberglass windshield retainer frames, spacer bushings may fall out. Retain all spacer bushings if using fiberglass windshield retainer frames. Metal windshield retainer frames do not require the use of spacer bushings. (4) Pull the rubber seal and rubber filler free from the window frame and channel. (5) Scrape away any remnants of sealing compound, filler or seal that may have adhered to the window frame. (6) Wash the window frame and channel with solvent (14, Table 1, Chapter 91-00-00).

B. Installation Proper performance of this procedure should ensure a stress-free windshield installation. Kit No. 101-5041-1 provides the parts and detailed instructions for installation of both windshields. Kit No. 101-5041-3 provides the parts and detailed instructions for installation of either windshield. The windshields are not part of the kits. In the event of conflict between the procedures in this manual and the instructions provided with the kits, the kits instructions should be followed. Prior to installing the windshield, inspect the windshield frames for old sealer or other obstructions that could cause the windshield to bind or seat improperly. Check all windshield mounting holes with a screw to ensure that all nutplates are in alignment and undamaged. It is recommended that the heating element and temperature sensor resistance of the new windshield be checked (Ref. Chapter 30-40-00, WINDSHIELD HEATING ELEMENTS RESISTANCE CHECK). NOTE: After mixing the sealer used to install the windshield, two people must work constantly to complete the installation of the windshield before the sealer sets up. (1) Mask off the exterior area around the windshields with masking tape. Mask off the nutplates by covering the screw holes on the inside of the windshield frames with masking tape. (2) Cover any exposed windshield glass with masking tape. Press the masking tape around each of the countersunk holes. Cut the masking tape from the countersink of each screw hole with a sharp knife until all masking tape has been removed from the screw holes. (3) Trim the excess protective paper from the inside of the windshield along the raised lip.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (4) Install the AN-15A bolts (10 places) through holes 5, 7 and 12 on the top and bottom and through holes 4 and 10 on each side of the windshield (Ref. Figure 201). Tape the bolt shanks to hold them in place with the bolt heads against the inside surface of the windshield. (5) Fabricate four guide pins by cutting the heads from 3 inch long AN3 bolts. Insert a guide pin and secure it in place in hole 1 at the top and bottom of the frame for each windshield half. (6) Using the guide pins for alignment, carefully center the windshield in the frame as evenly as possible. Check for a minimum clearance of 0.020 inch between the inside glass and the lower frame member. If necessary, place a shim (drill rod or drill shank of sufficient diameter) between the edge of the inside glass and the frame near the lower guide pin. Mark the location of the shim for future reference. If the required clearance cannot be maintained, or interference between the windshield and adjacent structure is encountered, contact Beechcraft Corporation Product Support. (7) With the windshield centered, use a 3 inch long straightedge to match the contour of the windshield with that of the windshield frame. If the windshield rocks on the bolt heads, center the windshield as evenly as possible to equalize the free play. Mark at least four places (one or more on at least three edges) where the exterior surface of the airplane is level with the metal retaining ring of the windshield. If there are not four places that are level, mark the places where the contours most nearly match and identify them as high or low in inches as measured. (8) Remove the windshield from the windshield frame. Remove the ten bolts taped to the windshield. Remove the shims (if installed in Step (6)). Do not remove the guide pins from the windshield frame. (9) Using solvent (14, Table 1, Chapter 91-00-00), clean the contact surfaces of the windshield frame and the new silicone seals. The 50-420066-341 silicone seals are to be applied to the sides and centers of the windshield frame, the 50-420066-343 seals to the bottom and the 50-420066-345 seals to the top. (10) Glue the silicone seals flush with the inside periphery of the windshield frame with silicone adhesive (57, Table 1, Chapter 91-00-00). The ends of the dam material used at the corners should be butted together snugly. (11) The 50-420066-347 sponge rubber strips are to be applied to the sides and center of the windshield, the 50-420066-349 strips to the bottom, and the 50-420066-351 strips to the top. Apply a 0.25 inch wide strip of adhesive (16, Table 1, Chapter 91-00-00) to the inside periphery of the sponge rubber strips and apply a similar strip of cement to the windshield between the screw holes and inside periphery of the windshield where the glass thickens. Glue the sponge rubber strips to the windshield by matching the cement strip on the rubber to the cement strip on the windshield glass. (12) Trim the sponge rubber strips on the ends so that the ends are butted against each other. Trim the outer edge of the sponge rubber strips until a 0.064 inch overlap is maintained around the periphery of the windshield. (13) Fabricate a hollow drill by sharpening one end of a piece of steel tubing with an inside diameter of 0.064 inch. Cut the tubing to the proper length. Fabricate a drill bushing from a 2 to 3 inch length of aluminum tubing having an outside diameter that will fit freely into the screw holes in the windshield and an inside diameter sufficient for insertion and rotation of the drill tube just fabricated. Insert the fabricated drill and bushing into a screw hole of the windshield to ensure that the drill will not bind in the windshield mounting holes.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (14) Using the hollow drill with a block of wood as a backup, drill out all the screw mounting holes through the sponge rubber strips. It may become necessary to punch out the accumulated rubber from inside the drill. (15) Mix approximately 1,200 grams of sealer (38, Table 1, Chapter 91-00-00) in accordance with the manufacturer's recommendations. Apply a 0.25 inch deep bead of sealer around the periphery of the windshield frame in the valley between the silicone rubber seals and the frame. CAUTION: Fiberglass windshield retainer frames require spacer bushings to prevent attaching screws from collapsing the fiberglass frame on installation, which may cause the windshield to crack some time later. Metal windshield retainer frames do not require the use of spacer bushings. (16) Using the guide pins installed in Step (5) for alignment, carefully position the windshield in the frame. Replace the shims removed in Step (8) between the windshield and frame. The shims are to remain in place until the sealer is completely cured. Start, but do not tighten, all screws in the open mounting holes. To facilitate installation of the screws through the rubber seals, lubricate all the screw threads with liquid soap. When installing the screws, carefully note the depth to which the screws penetrate the frame before actually entering the nut plate. In areas where the windshield will remain high after installation, screws with a longer grip length will be required. This can be gaged by visually comparing the amount these screws protrude above the frame after the marked screws are tightened. As a general rule, holes where the windshield is more than 1/16-inch above contour will require the longer screws. (17) Tighten the screws at the locations marked in Step (7) until the windshield is positioned for the contour match as marked. Do not tighten any of the remaining screws. (18) Remove the guide pins and install screws in the guide pin holes. Tighten all screws only sufficiently to seat the screw heads in the countersunk holes of the retainer. Do not tighten the screws as this could create stress by pulling the windshield tighter into the frame. After these screws have been seated into the countersinks, the screws tightened in the previous Step should still be seated and snug against the retainer. (19) Mask off all screw heads with masking tape and smooth and fair the sealer between the windshield and windshield frame. Liquid soap may be used to facilitate smoothing the sealer. Remove the masking tape and clean up the area when the smoothing and fairing process has been completed. NOTE: Following windshield replacement, the aircraft may continue operation Unpressurized, while the sealer achieves a final cure. To allow continued operation, prior to final cure follow Step (20) without deviation, otherwise proceed to Step (21). (20) The sealant must be in a setting stage (usually 1 to 2 hours after application). Ensure adjacent surfaces are clean and free of any oily film and apply aluminum speed tape (187, Table 1, Chapter 91-00-00) as follows: CAUTION: The 2 inch wide tape must be applied evenly, centered on the seal. This refers to the seal between the windshield frame and the airframe. By following this procedure the tape cannot come in contact with the windshield glass. Under no circumstance must the tape be allowed to contact the windshield glass. (a) Adhere the tape evenly paying particular attention to edge sealing.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL NOTE: Flight altitude temperatures may be different than ground or hanger temperatures and must be considered when calculating cure time requirements. The aircraft can be pressurized and tape removed when the appropriate cure time has been achieved. CAUTION: Do not attempt to pressurize the cabin or tighten the screws until the sealer has completely cured. (b) When sealant cure time is achieved, proceed to Step (22) to complete the windshield installation. NOTE: For every 10°F rise in temperature above 77°F, reduce the cure time by half; for every 10°F below 77°F, double the cure time. The cure time of the sealer can be accelerated through the use of heat lamps or circulation of warm air. Do not allow the temperature to exceed 140°F. (21) Allow sealer (38, Table 1, Chapter 91-00-00) to cure for 30 hours or more as required at an ambient temperature of 77°F, at 50% humidity. (22) After the required cure time has elapsed, torque the windshield mounting screws to 20 inch pounds. After torquing, check all screws to ensure the countersunk heads are properly seated and that all screws are tight. If shims were used between the windshield and frame, remove them and fill the gap with sealer. (23) Touch-up paint in the windshield area as required. Paint the windshield retainer in the following manner: CAUTION: Never use aluminum foil to mask electrothermal windshields during painting. Most metal brighteners will combine with aluminum to form a hydrogen gas that corrodes the stannous oxide used as an antistatic coating on the windshields. (a) Mask the windshield and surrounding structure with paper or pasteboard masking material. (b) Remove the old paint with paint stripper (8, Table 1, Chapter 91-00-00). (c) Sand the windshield retainer with 120 to 280 grit paper to remove the original anodized coating. (d) Clean the sanded retainer with solvent (14, Table 1, Chapter 91-00-00) to remove any residue. (e) Prime the retainer with primer (5, Table 1, Chapter 91-00-00). (f) Paint the windshield retainer with the appropriate color urethane paint (6, Table 1, Chapter 91-00-00). (24) Attach the ground jumper wire at the outboard corner of the windshield and connect the power supply lead at the windshield corner post.

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Figure 201 Windshield Installation

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2. WINDSHIELD WEATHER SEAL INSPECTION AND REPLACEMENT A. Seal Inspection Inspect the weather seal between the glass and windshield retainer at the intervals specified (Ref. Chapter 05-20-00, CONTINUOUS INSPECTION SCHEDULE). (1) Turn off all electrical power to the airplane. NOTE: The entire weather seal should be inspected to determine if it is well bonded to glass and retainer surfaces. If there is evidence that portions of the seal can be lifted with finger or spatula pressure, the seal must be repaired or replaced. If the Hump Seal is damaged, a temporary repair (good for 7 days) can be made. This temporary repair will allow the operator to fly the airplane until it can be scheduled into maintenance such that either the hump seal repair or hump seal replacement can be carried out. Refer to the TEMPORARY HUMP SEAL REPAIR procedure. It is essential that repairs are made as needed to ensure maximum windshield service life. (2) Thoroughly inspect the weather seal for debonding, cracks, or deterioration. Particular attention should be given to areas eroded by the wiper blades along the top and bottom of the windshield. To ensure the windshield is serviceable if the seal is damaged or debonded, refer to WINDSHIELD REPLACEMENT CRITERIA for moisture damage. If no moisture damage is noted you may either repair or replace the hump seal. Refer to Step (3) below. NOTE: Weather seals made from silicone material must be repaired with sealer (38, Table 1, Chapter 91-00-00). Weather seals made from polysulfide material must be repaired with sealer (124, Table 1, Chapter 91-00-00). It is easy to distinguish the difference between the seal material used. The silicone seal will be flush with the windshield and the frame surface and will appear translucent white or a milky color. The polysulfide seals will be humped over the seal area and appear either gray or black depending on the date of manufacture. (3) To repair silicone seals, refer to SILICONE SEAL REPAIR. Replace the hump seal using Seal Kit No. 101-5172-1 (LH) or -2 (RH), or repair (Ref. HUMP SEAL REPAIR/REPLACEMENT). NOTE: It is acceptable to repair or replace existing gray or black hump seals, or to replace gray hump seals with black hump seals on all 1900 Airliner models except for some very old models that do not have the latest windshield design. All hump seals are to be installed per kit instructions.

B. Silicone Seal Repair (1) Turn off all electrical power to the airplane. CAUTION: To prevent scratches, nicks, or cracks to the windshield, do not use a metallic scraper when removing old sealant material. (2) Carefully remove any damaged or deteriorated sealant. (3) Clean the damaged area of the seal with solvent (54, Table 1, Chapter 91-00-00). Wipe with a clean, dry cloth before the solvent evaporates.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (4) Apply sealer (38, Table 1, Chapter 91-00-00) to the damaged area until flush with the existing surface and overlapping the windshield and retainer by 0.125 to 0.250 inch. Smooth sealer to match the original contour. Liquid soap may be used to smooth sealer. NOTE: At 75°F with 50% relative humidity, sealer (38, Table 1, Chapter 91-00-00) is tack free in 8 hours. Cure time is 30 hours. For every 10°F rise in temperature (to a maximum of 140°F) the cure time is cut in half. For every 10°F decrease the cure time is doubled. (5) Allow the sealer to cure before flight or perform Step (20) of WINDSHIELD INSTALLATION.

C. Temporary Hump Seal Repair (UB-1 thru UB-74 and UC-1 thru UC-174) NOTE: This is a temporary hump seal repair and may be used for a maximum of 7 days. After which time, a permanent repair must be accomplished. The aluminum speed tape should be only wide enough to ensure full coverage over the channel between the glass and the frame to prevent the ingression of moisture. (1) Apply aluminum speed tape (187, Table 1, Chapter 91-00-00) to the outside periphery of the windshield. There are no tape length limitations.

D. Hump Seal Repair/Replacement NOTE: This procedure applies to installed windshields that have hump seals as the primary weather seal between the glass and retainer. The seals will be either gray or black depending on the date of manufacture. (1) Turn off all electrical power to the airplane. CAUTION: To prevent scratches, nicks or cracks to the windshield, use extreme care if using metallic tools to remove old sealant material. (2) Apply masking material to the outer surface of the windshield leaving 1 1/2 inches of exposed glass along the glass edge interface. CAUTION: Extreme care must be taken if removing sealant along the bottom of the windshield. Antistatic grounding tabs are located as shown (Ref. Figure 202). Be careful not to abrade or scratch the outer glass ply surface with the Scotch-Brite pad or sandpaper. (3) With a single-edge razor blade held at a very shallow angle, carefully remove any damaged or deteriorated sealant. This can be made easier by slightly bending the razor blade at the center to ensure the corners do not scratch or cut into the outboard glass surface. All sealant can be removed from the windshield surface with the razor blade. The small amount of sealant that remains on the retainer can be removed using: 80 grit sandpaper on fiberglass frames, or a Scotch-Brite pad (150, Table 1, Chapter 91-00-00) soaked in solvent (30, Table 1, Chapter 91-00-00) on metal frames. Clean a small area at a time, followed by alcohol rinse (30, Table 1, Chapter 91-00-00), then dry with a clean cloth. Discard soiled cloths regularly to prevent redepositing contaminants. Sealant in the gap between the windshield and retainer does not need to be removed and it is permissible to leave a small amount of sealant on the antistatic tabs. CAUTION: In Step (4), ensure a slurry is maintained. Rubbing the windshield with dry cleanser will scratch the windshield.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (4) Soak cheese cloth with water and using a slurry of cleanser (149, Table 1, Chapter 91-00-00) and water, polish the glass surface by hand only in the repair area where the seal material will be applied until a water-break free surface is observed (a water-break free surface completely wets or sheets over the glass with no sign of drawing up into droplets showing dry areas in between). Achieving a water-break free surface is critical to insure adhesion of sealant to the outer glass surface. (5) After a water-break free surface is obtained, clean the entire area with solvent (30, Table 1, Chapter 91-00-00) and allow to air dry. (6) Mark the retainer in the repair area to match the outer edge of the remaining seal and mark the windshield in the repair area to match the inner edge of the remaining seal. Apply 1 inch masking tape along the marks on the retainer and along the marks on the windshield (you may use several layers of tape to form an edge to guide the hump seal forming tool). The area left open should match the original area where the hump seal was. Ensure the remaining area of glass (daylight opening) is masked to protect the glass surface during the repair operation. (7) Wet a piece of clean cheese cloth with primer (151, Table 1, Chapter 91-00-00) and in one smooth, continuous motion apply a thin coat of primer to the surface of the windshield and retainer. The primer should dry clear. If a haze, or streaks appear, repeat Steps (3) thru (6) and apply primer. (8) Mix sealant (124, Table 1, Chapter 91-00-00) per manufacturers instructions. Using a plastic spatula, apply sealant to repair area. (9) Using the hump seal tool (make from 2 inch plastic spatula or equivalent by filing the contour of the original hump seal into the flat edge), form the hump seal, matching the contour of the existing weather seal as closely as possible. (10) Blend the ends of the repair area to match the existing seal by rubbing very lightly with a cellulose sponge saturated with water. (11) After the seal is formed, immediately remove the 1 inch masking tape while the sealant is wet. (12) Smooth the surface of the sealant by rubbing the surface of the sealant very lightly and briskly with a cellulose sponge saturated with water. (13) Allow sealant to cure. Cure time reference is based on 75°F and 50% relative humidity. For every 10°F to 15°F rise in temperature (to a maximum of 140°F), the cure time is cut in half. For every 10°F to 15°F decrease in temperature, cure time is doubled. Sealant (124, Table 1, Chapter 91-00-00) is tack free in 8 hours; full cure time is 30 hours. (14) After sealant is cured, remove masking and clean as necessary. Inspect for voids and repair as necessary.

3. WINDSHIELD A. Replacement Criteria The electrically heated laminated glass windshield is subject to a gradual process of delamination due to the effect of chemical action and differentials of temperature and pressure incurred during pressurized flights at varying altitudes and under varying weather conditions. This delamination is not detrimental to the structural integrity of the windshield, although it may significantly decrease visibility or the deicing capability of the windshield. Beyond certain limits, either of these effects will require the Page 208 Nov 1/13

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL replacement of the windshield. The following information is provided to evaluate the condition of the windshield. Refer to the 1900 Airliner Series Airworthiness Limitations Manual P/N 129-590000-133, Chapter 04-00-00, CRACKED OR SHATTERED WINDSHIELD. . (1) Visibility Impairment: Replace the windshield when areas of delamination are enlarged to the point of impairing vision, whether or not these areas extend to the edge of the glass. Smaller delamination appearing as bubbles is no cause for concern. The rate of growth of delaminated areas should be used as a guide for scheduling windshield replacement. (2) Heating Impairment Limits: Replace the windshield when the lost heating area exceeds 1/4 to 1/3 of the total heated area. (3) Cracked Glass: (a) Applicable to P/N 114-384020-3/4/5/6 Both glass plies are chemically strengthened and are structural plies. Depending on the stress level on the glass at the time of fracture, the break pattern will generally be of relatively large size fragments adhered to the interlayer. Refer to the 1900 Airliner Series Airworthiness Limitations Manual P/N 129-590000-133, Chapter 04-00-00, CRACKED OR SHATTERED WINDSHIELD. (b) Applicable to P/N 101-384025-(all dash numbers) If the outboard glass ply should fracture, the break pattern will generally be of relatively large in size fragments adhered to the interlayer. If the inboard glass ply should fracture, the break pattern will generally be of small size fragments adhered to the interlayer. Refer to the 1900 Airliner Series Airworthiness Limitations Manual P/N 129-590000-133, Chapter 04-00-00, CRACKED OR SHATTERED WINDSHIELD. (c) Cracking of the interlayer is usually located in the PPG112 interlayer. Cracking is caused by moisture attacking the interlayer and causing the interlayer to degrade. The manufacturer recommends the same criteria be used for this defect as is currently applied to delamination. (4) Scratches: (a) Applicable to P/N 114-384020-3/4/5/6 Scratches or nicks on either surface as deep as 0.005 inch are considered acceptable as long as vision is not seriously impaired. Scratches deeper than 0.005 inch require the windshield to be replaced. Refer to the 1900 Airliner Series Airworthiness Limitations Manual P/N 129-590000-133, Chapter 04-00-00, CRACKED OR SHATTERED WINDSHIELD. (b) Applicable to P/N 101-384025-(all dash numbers) There are no limits on these defects for this part numbered windshields as long as vision is not seriously impaired. Refer to the 1900 Airliner Series Airworthiness Limitations Manual P/N 129-590000-133, Chapter 04-00-00, CRACKED OR SHATTERED WINDSHIELD. (5) Glass Adhesion Chips: Glass adhesion chips can occur on the inboard surface of the outer ply, or the outboard surface of the inner ply. Chips are caused by the vinyl interlayer pulling glass out of the glass surface and usually form at or near the glass edge. The chip will continue to expand/propagate until the stress that caused the chip is relieved. A chip in a non-structural ply is not cause for windshield replacement unless vision is seriously impaired. A chip in a structural ply causes major glass strength reduction, and is cause for windshield replacement. Refer to the 1900 Airliner Series Airworthiness Limitations Manual P/N 129-590000-133, Chapter 04-00-00, CRACKED OR SHATTERED WINDSHIELD.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (6) Moisture Damage: The primary cause of moisture ingress is a deteriorated weather seal. Moisture ingress caused by a deteriorated weather seal will eventually lead to one or all of the following windshield degradation modes: (a) Delamination between the glass and PVB interlayer (Separation between glass and interlayer characterized by a separation line with a different reflection characteristic on each side of the line). (Ref. Visibility Impairment). (b) Corrosion of the bus bar - heating wire junctions (black spots at the bus bar - heating wires interface). (c) Failure of the electric heating system (lost heating area exceeds 1/4 or 1/3 of the total heated area). (Refer to Heating Impairment Limits). (d) Outer glass ply fracture (Ref. Cracked Glass). If the windshield is severely degraded due to any or all of the above conditions, the useful service life is limited and is cause for windshield replacement. Refer to the 1900 Airliner Series Airworthiness Limitations Manual P/N 129-590000-133, Chapter 04-00-00, CRACKED OR SHATTERED WINDSHIELD.

4. ANTISTATIC TAB A. Installation (Aluminum Frames) NOTE: Kit No. 90-5048-1S provides the parts to install two new removable type antistatic tabs on the removable or bonded tab windshields. (1) Remove the existing antistatic tab from windshields that have removable tabs (Ref. Figure 202). Clean the windshield glass with solvent (30, Table 1, Chapter 91-00-00). A wood or plastic scraper may be used to aid in removing the sealer. (2) Center the new antistatic tab on the mark on the windshield retainer that tested 100 megohms or less in ten different locations. A 0.59 inch section of the antistatic tab must cover the windshield glass. Mark the antistatic tab top end at a point 0.5 inch from the outer side of the windshield retainer. Drill a 0.098 inch diameter hole through each side of the tab (Ref. Figure 203). (3) Using the two holes drilled in the antistatic tab as a guide, mark the windshield retainer. (4) Drill two 0.0935-inch diameter holes at the marks on the windshield retainer. Vacuum the area to remove the debris created by the drilling operation. (5) Tap the two holes in the windshield retainer with a No. 4-48 flat bottom tap. NOTE: If the windshield frame is buffed, buff the antistatic tabs before installation. (6) Attach the antistatic tab to the windshield with two MS35207-212 screws furnished with the kit. Mask off the antistatic tab with masking tape so that, when the tab is removed, the periphery of the tab location will be outlined by the tape. Remove the antistatic tab and lightly sand the bottom end. Sand the windshield retainer until bare metal is exposed.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (7) Clean inside the masked off area with solvent (30, Table 1, Chapter 91-00-00) and apply a strip of electrical conductive tape (143, Table 1, Chapter 91-00-00) approximately 1.125 inches long. Center the tape in the masked off area. Install 0.47 inch of the conductive tape on the windshield glass (Ref. Figure 203). Press the electrical conductive tape down with finger pressure only. (8) Clean the bottom surface of the antistatic tab and the top surface of the electrical conductive tape with solvent (30, Table 1, Chapter 91-00-00). Apply an electrical conductive sealant, such as sealer (144, Table 1, Chapter 91-00-00), to the top surface of the electrical conductive tape. (9) Install the antistatic tab on the windshield retainer, sealant should flow from under the edges of the tab. If the sealant does not flow out around the tab, remove the tab and apply additional sealer. NOTE: Do not press on top of the tab in an attempt to squeeze out the sealer. (10) Remove tape from around the antistatic tab and allow sealant to cure as recommended by the product instructions. (11) Mask off the four sides of the tab, leaving 0.0937 inch clearance between the edge of the masking tape and the edge of the tab. Mask off the top surface of the antistatic tab. Do not mask the area between the masking tape and the tab. (12) Apply adhesive (145, Table 1, Chapter 91-00-00) or equivalent to the four sides of the antistatic tab to provide a weather seal. Remove the masking tape and allow the adhesive to cure as recommended by the manufacturer’s instructions. (13) If necessary, paint the antistatic tab to match the finish on the windshield retainer frame. Do not buff the tab after installation on the airplane. (14) Clean the windshield (Ref. 56-00-00, CLEANING WINDSHIELDS).

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Figure 202 Windshield Antistatic Tab

Figure 203 Windshield Antistatic Tab Installation

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5. WINDSHIELD ANTISTATIC COATING AND TAB A. Inspection (Aluminum Frames) Inspect the electrically heated windshields as specified in Chapter 05-20-00. (1) Turn off all electrical power and disconnect the battery. (2) Clean the exterior of the windshield glass with solvent (30, Table 1, Chapter 91-00-00). CAUTION: To prevent damage to the windshield, refer to CLEANING WINDSHIELDS in 56-00-00. NOTE: A megometer capable of reading 100,000,000 ohms (100 megohms) will be required for the following inspection. (3) Set the megometer range dial on 100M and adjust the needle reading to zero. (4) If necessary, the finish may be removed from a spot on the windshield frame to expose bare metal to be used as a ground point. (5) Attach one test lead to the bare metal on the windshield frame. Attach the other test lead to a copper wire mesh pad, such as a Brillo scrub pad. (6) Contact the windshield glass with the wire mesh pad. Check the meter reading. If the meter reading is 0 to 100 megohms in at least 10 different locations, no further action is required. If the finish was removed from a spot on the windshield frame for this test, repaint the spot to match the surrounding finish. NOTE: Resistance between the windshield and windshield frame should not exceed 100 megohms. Consistency in the readings is desirable. (7) If the megometer reading is more than 100 megohms, proceed with the following Steps: NOTE: Do not remove the windshield from the airplane as a result of this inspection. Contact the Customer Support Department of Beechcraft Corporation for consultation. (8) If some of the readings were more than 100 megohms, draw a diagram noting the locations of the excessive readings, and contact the Customer Support Department of Beechcraft Corporation for consultation. (9) If all the megometer readings were more than 100 megohms, perform the following antistatic tab inspection Steps: (a) Remove the megometer test lead from the wire mesh pad. (b) Place the megometer test lead on the windshield glass within 0.50 inch of the antistatic tab. Do not press down on the antistatic tab during this inspection. (c) If the meter reading is 10 megohms or less as the lead is drawn around the antistatic tab, the antistatic tab is providing sufficient electrical continuity between the windshield retainer and the antistatic coating on the windshield glass. If the meter reading is more than 10 megohms, continue with the following Steps.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL NOTE: During the following test, ensure that the hands and body are completely insulated from the airplane, windshield, and meter probes. If more than one antistatic tab is installed, check each tab. (d) Connect the meter test leads to wire mesh pads. Place one pad within 0.50 inch of the antistatic tab. Place the other wire mesh pad in 10 different locations on the windshield and note the resistance readings. If the 10 resistance readings are 100 megohms or less, a new antistatic tab must be installed. (10) If intermittent readings of 100 megohms or less were indicated, repeat Step (9). If all resistance readings were more than 100 megohms, proceed with the following Steps: (a) Move the megometer test lead and pad approximately four inches away from the edge of the antistatic tab and within 0.25 inch of the windshield retainer. Hold the pad stationary on the windshield glass. (b) Place the other pad and test lead in 10 different locations on the windshield and note the resistance readings. (c) Place a mark on the windshield retainer in the vicinity of the stationary pad. At the mark, write the number of resistance readings that were 100 megohms or less. If all ten resistance readings were 100 megohms or less, install a new antistatic tab. Center the new tab on the mark made on the windshield retainer. (d) If ten resistance readings of 100 megohms or less were not found, move the ohmmeter test lead approximately four inches from the previous test point and take an additional ten resistance readings with the wire mesh pad. The other ohmmeter test lead must remain within 0.25 inch if the windshield retainer. (e) Connect the meter test leads to wire mesh pads. Place one pad within 0.50 inch of the antistatic tab. Place the other wire mesh pad in 10 different locations on the windshield and note the resistance readings. If the 10 resistance readings are 100 megohms or less, a new antistatic tab must be installed. NOTE: If meter readings of 100 megohms were not indicated at any test point, do not remove the windshield from the airplane until contacting the Customer Support Department of Beechcraft Corporation for consultation. (11) Restore electrical power as required.

B. Coating Inspection (Lightweight Fiberglass Frame) The electrically heated windshields are to be inspected at the interval specified in Chapter 05-20-00. This inspection may be accomplished as follows: (1) Turn off all electrical power and disconnect the battery. (2) Clean the exterior of the windshield glass with solvent (30, Table 1, Chapter 91-00-00). CAUTION: In order to avoid damaging the windshield, use extreme care when cleaning the windshield glass and performing this inspection. Do not remove the windshield from the airplane.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (3) Open the nose avionics compartment access door and locate a suitable location to serve as a grounding point for the volt-ohmmeter test lead. NOTE: A megometer capable of reading 100,000,000 ohms (100 megohms) will be required for the following inspection. (4) Set the range dial on the meter to 100M and adjust the reading needle to zero. (5) Attach one meter test lead to the ground point inside the nose avionics compartment. Attach the other test lead to a copper wire mesh pad, such as a Brillo scrub pad. (6) Contact the windshield glass with the wire mesh pad. Check the meter reading. If the meter reading is 0 to 100 megohms in at least ten different locations, no further action is required. NOTE: Resistance between the windshield and ground should not exceed 100 megohms. Consistency in the readings is desirable. (7) If the meter reading is more than 100 megohms, check the connection between the anti-static bus bar and ground. If the connection is not secured properly, or if it is corroded, use appropriate measures to repair the connection. If the connection is good, the anti-static coating on the windshield is no longer functioning properly. The windshield must be removed in order to have it coated or replaced. (8) Restore electrical power as required.

6. BONDED ANTISTATIC TAB A. Repair (Aluminum Frames) Inspect the electrically heated windshields as specified in Chapter 05-20-00. (1) Tape off an area around the fiberglass cover tab. The area should be large enough to protect the glass and contain excessive sealant flow. (2) Using a plastic wedge, gently pry the fiberglass tab away from the windshield surface. Do not disconnect the tab from the metal edge attachment. (3) Fold the fiberglass cover tab back over the metal edge attachment and tape it down out of the work area. (4) Using a plastic wedge, pry the copper tab away from the windshield surface. Do not break or sever the braided wire attaching the copper tab. Tape the tab out of the work area. (5) Gently scrape the sealant from the windshield surface with the plastic wedge. (6) Remove any residual sealant by rubbing the area with a soft cloth dampened with solvent (14, Table 1, Chapter 91-00-00). The glass may exhibit some staining that cannot be removed; however, it is not detrimental to the tape bond. (7) Using a 40 watt pencil type soldering iron, melt the solder attaching the copper tab to the braided wire and separate the wire from the tab. (8) Cut a piece of copper foil tape (146, Table 1, Chapter 91-00-00) 1/2 inch long.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (9) Remove the tape backing, being careful not to contaminate the adhesive with body oils or dirt. Place the copper tape over the area originally used for tab bonding. Rub the tape down to the surface of the windshield with a plastic wedge. (10) Position a section of wire solder on the center of the copper tape (installed in the previous Step). Apply the tip of the pencil soldering iron to the solder just long enough to melt a puddle of solder with a maximum diameter of 1/8 inch. Remove the soldering iron and the solder simultaneously. (11) Using tweezers or some other suitable tool, hold the braided wire to the top of the solder spot. Apply the tip of the soldering iron to the braided wire at the solder spot. Remove the soldering iron as soon as the solder has blended with the braided wire. (12) Remove any excess flux with a cotton swab dampened with solvent (14, Table 1, Chapter 91-00-00). Wipe the interior of the taped-off area with solvent (30, Table 1, Chapter 91-00-00). CAUTION: Use only minimal amounts of solvent and isopropyl alcohol, as either could destroy the adhesive bond of the copper tape. (13) Thoroughly clean the bond surface of the fiberglass cover tab with solvent (30, Table 1, Chapter 91-00-00). Using a cotton swab, apply a moderate coat of primer (147, Table 1, Chapter 91-00-00) to the bond surface of the fiberglass cover tab. Allow the primer to dry for approximately two minutes. (14) Apply a 1/16 to 1/8 inch thick layer of electronic grade silicone rubber (148, Table 1, Chapter 91-00-00) over the copper tab and inner boundaries of the taped off area. (15) Fold the fiberglass cover tab back to its original position and tape it in place. Apply sufficient pressure to the cover tab with the tape to assure a weather-proof seal. Allow the sealant to cure for 24 hours before removing the tape. (16) Remove excessive sealant flow from the tape by gently rubbing it with a plastic wedge. After the excessive sealant has been removed, remove the tape. (17) Clean the windshield (Ref. 56-00-00, CLEANING WINDSHIELDS).

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7. FLIGHT COMPARTMENT SIDE WINDOW A. Removal (1) Remove the headliner from its retainer strip at the top of the window. NOTE: If the RH window frame is to be removed, remove the screws from the RH circuit breaker panel and move it slightly away from the upholstery before removing the upholstery panel from its retainer. (2) Remove the upholstery panel from its retainer strip at the bottom of the window frame. (3) Remove the attaching screws from both the upper and lower retainer strips and remove the retainer strips (Ref. Figure 204). (4) Remove the bolts and washers from the storm window stop and remove the stop. (5) Work the plastic window frame away from the airplane structure. NOTE: The inner window will be removed along with the window frame. If the inner window requires removal, it may be pressed out of its retaining groove in the plastic window frame with hand pressure. (6) Remove the screws from the outer window retainer and remove the retainer. (7) Carefully remove the outer window from the airplane. (8) Clean all residual sealer from the inner retainer ring and the outer window frame.

B. Installation WARNING: Because of manufacturing variables, there may be some dimensional variation in the flight compartment side window cutouts. These cutouts may be somewhat larger than standard on some airplanes. A window with standard dimensions must not be installed in an oversize cutout because the milled shoulder on the standard window will not be sufficient to assure structural integrity. To determine the correct replacement window, measure the originally installed window (Ref. Figure 205). (1) Apply sealant (124, Table 1, Chapter 91-00-00) around the periphery of the side window frame. (2) Install a new 90-380030-3 seal around the rim of the window with the sealing ridges on the outboard side of the window. All surfaces of the seal must lay flat against the window when installed (Ref. Figure 204). (3) Position the window in the outer window frame. Ensure that the window is centered in the frame. (4) Install the metal retainer ring and secure it with the attaching screws. CAUTION: Use care not to cut or scratch the window pane when cutting away center portion of window seal. (5) Using a single-edge razor blade, carefully cut away the center portion of the window seal covering the outer surface of the window pane. Cut the seal flush with the cutout in the airplane skin.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (6) Remove the protective paper from the outer window pane. Thoroughly clean the outer window pane and the inner window. (7) If necessary, place the inner window in the groove in the plastic window frame and snap it into place with hand pressure. (8) Install the plastic window frame in place around the side window. Position the upper and lower upholstery retainer strips in place on the window frame and secure each with the attaching screws. (9) If necessary, install the RH circuit breaker panel and secure it with the attaching screws. (10) Install the upholstery panel and the headliner in their respective retainer strips.

8. WINDOW ATTACH FRAMES A. Inspection and Repair (1) Press on the cabin windows and the flight compartment side windows from the outside of the airplane. If no movement is observed, no further action will be required. If there is an indication of looseness, proceed as indicated in the following Steps. (2) Remove each loose window as indicated in the preceding portions of this chapter. (3) With the frame pressed back against the outer skin, drill 0.098 inch diameter holes through the frame and skin at a distance of 0.20 inch from the edge of the frame with a No. 40 drill. Drill the holes 6 inches apart in the areas where the frame has separated from the skin. (4) Countersink the holes 100 degrees on both sides. The holes on the frame assembly side need be countersunk only enough to hold the frame in place while the epoxy adhesive dries and the rivet butts are made flush. (5) Pry the frame and skin apart enough to sand their mating surfaces with 380-grit sandpaper until all the adhesive is removed. CAUTION: Pry the frame and skin apart only enough to permit repair. Any attempt to break the two free from one another may result in damage to either or both. (6) Clean the mating surfaces with solvent (17, Table 1, Chapter 91-00-00) and a clean cloth until all the residue of adhesive is removed. (7) Mix epoxy adhesive (93, Table 1, Chapter 91-00-00) in accordance with the manufacturer's directions and apply it onto the mating surfaces (Ref. Figure 204). (8) Install NAS1097 rivets and flatten the rivet butts enough to fill countersink. (9) Flush the rivet butts to the frame assembly and remove all the excess epoxy adhesive with naphtha and a clean cloth. (10) Install the windows as outlined in the preceding portions of this chapter.

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Figure 204 Flight Compartment Side Window Installation

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Figure 205 Flight Compartment Side Window

9. STORM WINDOW A. Inspection (1) Inspect storm window (1) for cracks or crazing of the surface (Ref. Figure 206). (2) Check latch assembly (11), hinge halves (2 and 16) and hold open detent (17) for proper operation. (3) Inspect the primary (4) and secondary (optional) (3) seals for deterioration and proper fit to the frame and storm window (1). (4) Check the adjustment of the storm window. There should be no wind noise or water leakage. (5) Inspect the storm window frame for evidence of corrosion or trapped water.

B. Removal (1) Open storm window (1) until the hold open detent (17) engages (Ref. Figure 206). (2) Remove screws (5), washers (6), nuts (7) and gasket (19) from storm window (1) and storm window hinge half (2). Remove the storm window from the airplane. (3) If the hinge assembly is to be removed, remove bolts (18) securing hinge (16) to the airplane structure and remove the hinge. (4) If the storm window is to be replaced with a new window, remove capnuts (15), washers (14) and screws (13) from latch assembly (11). Remove the latch assembly and gasket (12) from the storm window.

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C. Installation and Adjustment (1) Inspect storm window (1), primary seal (4) and the optional secondary seal (3) for condition. Replace as necessary (Ref. Figure 206). (2) If storm window (1) is being replaced with a new window, install rubber gasket (12) and latch assembly (11) on the window with attaching screws (13), washers (14) and capnuts (15). Do not overtighten. (3) If the storm window frame hinge half (16) was removed, position the hinge half onto the airplane structure and secure with three attaching bolts (18). CAUTION: When attaching the acrylic storm window to the hinge half (2), torque the attaching screws 2 to 4 inch-pounds above the torque required to turn the nut with the screw. Exceeding the recommended torque values will cause the window to crack at the hinge attaching points. (4) Position gasket (19) and storm window (1) on window hinge half (2) and install the five attaching screws (5), washers (6) and nuts (7). Torque the screws 2 to 4 inch-pounds above that required to turn the nut with the screw. (5) Close and latch the storm window. (6) If primary seal (4) is the only seal installed, inspect storm window (1) for a tight fit against the window frame seal. If both primary and optional secondary (3) seals are installed, inspect the storm window by observing that both seals fit tightly together. (7) Adjust the hinge assembly along the three slotted holes, in hinge half (16), as required to achieve an effective seal or to allow the window to be closed and latched without undue force. If the adjustment is at the end of its travel, the slots in the hinge assembly may be lengthened slightly until the window can be closed and latched without undue force. The window should be positioned in the center of the cutout in the skin. (8) With the storm window closed and latched, pour water on the outside of the window. From inside the cockpit, check the lower edge of the storm window for leaks. (9) If leaks are found, repeat Steps (7) and (8) and then perform the STORM WINDOW STOP ADJUSTMENT procedure.

10. STORM WINDOW PRIMARY SEAL A. Removal (1) With storm window (1) opened or removed, remove the primary seal (4) and adhesive from the window frame. (2) Remove any remnants of the seal or adhesive from the window frame.

B. Installation (1) Ensure the storm window frame is completely free of any traces of old adhesive.

56-10-00

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (2) Apply adhesive (16, Table 1, Chapter 91-00-00) to the storm window primary seal (4) in the area it will contact the window frame and to the mating area of the window frame (Ref. Figure 206). Allow the adhesive to dry until tacky. (3) Install the storm window primary seal (4). Ensure that the seal fits flush against the storm window. Press the seal firmly to bond the seal to the window frame. (4) After the adhesive has completely cured, perform the STORM WINDOW INSTALLATION AND ADJUSTMENT procedure in this section.

11. STORM WINDOW SECONDARY SEAL A. Removal (1) Perform the STORM WINDOW REMOVAL procedure in this section. (2) Completely remove secondary seal (3) and adhesive from storm window (1). Ensure the window is completely free of any traces of seal or adhesive (Ref. Figure 206).

B. Installation (Optional) (1) Clean the perimeter of the storm window (1) (Ref. Figure 206) using solvent (54, Table 1, Chapter 91-00-00). (2) Using a cloth that is dry and clean, wipe the cleaned area dry before the solvent has evaporated. NOTE: Do not exceed a total adhesive thickness of 0.25 inch. (3) Apply adhesive (185, Table 1, Chapter 91-00-00) to the secondary seal (3) and the cleaned area of storm window (1). (4) Install storm window secondary seal (3) on storm window (1), ensure the seal fits flush against the storm window. Press the seal firmly to bond the seal to the window (Ref. Figure 206). (5) Allow the adhesive to cure for 30 minutes minimum at 77°F. (6) With primary seal (4) installed and the adhesive on secondary seal (3) completely cured, perform the STORM WINDOW INSTALLATION AND ADJUSTMENT procedure in this section.

12. STORM WINDOW STOP A. Adjustment (1) Open storm window (1) and loosen bolts (9) securing the storm window stop assembly (8) in its slotted mounts (Ref. Figure 206). NOTE: If the window stop assembly is at the end of its travel adjustment, the window stop may be bent outboard to obtain a tighter seal. (2) Slide the storm window stop assembly (8) outboard to obtain a tighter seal. (3) Tighten the storm window stop assembly mount bolts (9). (4) If leaks are found, perform Steps (7) and (8) contained in the STORM WINDOW INSTALLATION AND ADJUSTMENT procedure in this section. Page 222 Nov 1/13

56-10-00

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

1. STORM WINDOW 2. WINDOW HINGE HALF 3. SECONDARY SEAL (OPTIONAL) 4. PRIMARY SEAL 5. SCREWS (5 PLACES) 6. WASHER (5 PLACES) 7. NUT (5 PLACES) 8. STOP ASSEMBLY 9. BOLT (2 PLACES) 10. WASHER (2 PLACES) 11. LATCH ASSEMBLY 12. GASKET 13. SCREW (2 PLACES) 14. WASHER (2 PLACES) 15. CAPNUT (2 PLACES) 16. FRAME HINGE HALF 17. HOLD OPEN DETENT 18. BOLT (3 PLACES) 19. GASKET

A 3

19 16

1 5

2

6

17

7 4 18

13 12 14

15

9

11

10 8 DETAIL

A

UC56B 070467AA.AI

Figure 206 Storm Window Installation

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

WINDOWS CABIN MAINTENANCE PRACTICES

56-20-00 200200

1. CABIN WINDOW A. Removal (1) Remove the plastic trim around the opening of the window (Ref. Figure 201). (2) Remove the screws holding the trim strip positioned above the window at the top of the window panel and at the bottom of the headliner. Remove the trim strip. (3) Remove the screws in the lower trim strip located below the window. Remove the trim strip. (4) Remove the screws from the panel surrounding the window and remove the panel. (5) Remove the six screws from the window retainer and remove the retainer and the Lexan ring. (6) Remove the Plexiglas window from the window frame by pushing in on the window from the outside or by prying from the inside. Be careful not to damage the window frame, fuselage skin or any of the upholstery.

B. Installation (1) Install a new seal on the new window. Install the seal so the outboard surface of the window is covered by the protective portion of the seal (Ref. Figure 201). (2) Place the window and seal in the installed position in the window frame. (3) Place the Lexan ring in its installed position against the window. Place the retainer against the window frame. The ends of the retainer should be up and the flat surface should be towards the window. (4) Position one end of the retainer in the window frame so that a screw and washer can be started in one of the end holes. Tighten the screw no more than finger tight. (5) Push the retainer into the window frame so that a screw and washer can be started in the second hole from the end of the retainer. Tighten the screw finger tight. NOTE: It may be necessary to strike the retainer with a rubber mallet or wooden hammer handle in order to make it fit inside the window frame. Be careful not to damage the Lexan ring. (6) Working around the window and towards the other end of the retainer, insert and partially tighten the remaining screws and washers. (7) When all the screws and washers have been installed, tighten all the screws. (8) Replace the window panel with the screws and washers which were removed from the panel. (9) Replace the trim strips above and below the window. (10) Replace the plastic trim around the opening of the window.

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Figure 201 Cabin Window Removal and Installation

Page 202 Nov 1/09

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CHAPTER 57 - WINGS TABLE OF CONTENTS SUBJECT

PAGE

WINGS 57-00-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Wing Fuel Provisions (Effectivity: UA-1 and After, UB-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Wing Fuel Provisions (Effectivity: UC-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

1 1 1 1

MAIN FRAME 57-10-00 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Landing Gear Keel . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201

PLATES/SKINS 57-30-00 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Wing and Center Section . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Access . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .201

AILERON 57-50-00 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Aileron . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Balancing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Checking Balance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Balancing Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Equipment Required To Perform Check Balancing By Force Measurement Method . . . . . . . . . . . . . . Force Measurement Method . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Equipment Required To Perform Check Balancing By Counterbalancing Method . . . . . . . . . . . . . . . . Counterbalancing Method . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

57-CONTENTS

201 201 201 201 201 201 201 202 202

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List of Effective Pages CH-SE-SU

PAGE

DATE

57-LOEP

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57-CONTENTS

1

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57-00-00

1

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57-10-00

201 and 202

May 1/10

57-30-00

201 thru 204

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201 thru 205

May 1/11

C5

57-LOEP

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

WINGS DESCRIPTION AND OPERATION

57-00-00 00

1. GENERAL This wing is fabricated and assembled as a one-piece unit spanning from left Wing Station 291.735 to right Wing Station 291.735. The upper and lower main spar caps are continuous from left Wing Station 215.376 to right Wing Station 215.376. The lower main cap is made from three extrusions bonded together, with any two of the three capable of withstanding the limit load. The upper cap is a single machined extrusion. The rear spar caps are also machined extrusions. Outboard rear spar caps are joined to the center section caps in permanent joints outboard of the nacelles. These joints use nested angles as splice plates. Aluminum alloy sheet metal webs and stiffeners complete the spar assemblies. The primary structural materials in the wing are 2024 alclad aluminum alloy sheet, with 2014 and 7075 aluminum alloys used for forgings and extrusions. Outboard of the nacelle, the wing is a two-cell semimonocoque box structure of conventional design. The leading edge assembly and the main section assembly are joined to the wing spar with riveted lap joints. A subspar is installed forward of the main spar. The space forward of the subspar is utilized to route wiring, plumbing and engine controls. The nacelle keel members are machined aluminum alloy plate and incorporate the landing gear hinge point support structure. Formed sheet metal formers and stringers establish the nacelle contours and a cavity for the main landing gear. The ailerons are symmetrical sections, except in the wing tip area. They are hinged at the centerline and have overhanging aerodynamic balance. Each aileron has a single-channel-section spar as a main structural member. The aileron is attached to the wing at three hinge points. The aileron spar and skins are 2024 aluminum alclad sheet. An adjustable trim tab is located at the inboard end of the left aileron. The single-slotted flaps are fabricated in four sections, one on each outboard wing panel and one on each side just outboard of the fuselage. The structural material is 2024 alclad aluminum and 6061 aluminum sheet. Each flap has a single-channel-section spar. The outboard flap skins have chordwise beads in both top and bottom skins. The inboard flap skins are beaded chordwise aft of the spar on the bottom side only. Each flap is mounted on two tracks that are attached to the wing rear spar. Antifriction rollers in the flap roll in slots in the flap tracks.

2. WING FUEL PROVISIONS (EFFECTIVITY: UA-1 AND AFTER, UB-1 AND AFTER) Between the subspar and main spar, bladder fuel cells are installed, extending from just outboard of the nacelle to the wing tip assembly. Aft of the main spar, an integral fuel tank is a part of each outboard main section. Inboard of the integral tank, two bladder fuel cells are installed. Chordwise, the main section tanks extend from the front spar to the rear spar.

3. WING FUEL PROVISIONS (EFFECTIVITY: UC-1 AND AFTER) The wing fuel system is composed of an integral wet wing. Tubing is routed through the box section to be utilized for wire routing. A subspar is installed forward of the rear spar from Wing Station 114.25 through Wing Station 159.51 to provide utilization for control cable routing. Chordwise, the main section tanks extend from the leading edge skin to the rear spar. The main and rear spars in the center section of the wing are parallel in the plan view. A subspar located forward of the rear spar provides a tunnel for control cables and flap drive shafts and also serves as a fuel wall for the aft side of the center section wet fuel tank. The subspar extends from the root rib to the nacelle.

57-00-00

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

WINGS MAIN FRAME MAINTENANCE PRACTICES

57-10-00 200200

1. LANDING GEAR KEEL WARNING: Cracks found in the trunnion bolt area are not repairable. If cracks are found in this area, the entire keel must be replaced.

A. Inspection NOTE: Zone inspection area: 700 (1) Degrease keel area inside of wheel well. (2) Inspect the inboard and outboard keels in the left and right wheel wells. Inspect for cracks, corrosion, damage and loose or missing rivets. Inspect the entire keel area, paying particular attention to cracks emanating from the trunnion bolt area (Ref. Figure 201). (3) Using a 10X power or greater magnifying glass, inspect the flanged areas of the keel for cracks.

57-10-00

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 201 Landing Gear Keel

Page 202 May 1/10

57-10-00

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

WINGS PLATES/SKINS MAINTENANCE PRACTICES

57-30-00 200200

1. WING AND CENTER SECTION A. Access Refer to Figure 201 for views of the access openings in the wing and center section.

57-30-00

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 201 (Sheet 1 of 3) Wing and Center Section Access Openings

Page 202 Nov 1/09

57-30-00

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 201 (Sheet 2 of 3) Wing and Center Section Access Openings

57-30-00

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 201 (Sheet 3 of 3) Wing and Center Section Access Openings

Page 204 Nov 1/09

57-30-00

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

WINGS AILERON MAINTENANCE PRACTICES

57-50-00 200200

1. AILERON A. Balancing After repainting and/or repair, the finished aileron must be checked to ensure that its static moment about the hinge line is within the prescribed limits. The prescribed limits are 8.80 to 10.00 pound-inches nose-heavy. The static moment of the aileron is determined by multiplying the unbalanced weight of the aileron assembly times the perpendicular distance from the hinge centerline to the center of gravity when the chord line is horizontally level. The weight is measured in pounds and the distance in inches.

B. Checking Balance The balance must be checked in a draft-free area with the aileron completely assembled in flying condition. All painting, including stripes and touch-up, must be completed. The tab, tab pushrod, static wicks, and hinge bolts must be attached. The chord line must be horizontally level and the hinge line must be properly supported when the static moment is measured. Although many different methods of check balancing exist, they can be categorized under the following two headings: (1) Counterbalancing - The application of a known force or weight at a measured distance from the hinge line to counter the unbalance moment of the aileron assembly. (2) Actual Force Method - Measurement of the force applied by the aileron surface on a single support at a known distance from the centerline of the hinge. NOTE: Counterbalancing is the simplest method of check balancing.

2. BALANCING PROCEDURES A. Equipment Required To Perform Check Balancing By Force Measurement Method (1) A stand with knife-edge supports as illustrated in Figure 201. The knife edges should be in the same horizontal plane. (2) A certified beam balance calibrated in units of 0.01 pound(s) or less. The balance should have a flat weighing platform and its capacity should equal tare plus 2.0 pound(s) minimum. (3) A support spindle similar to the illustration and levelling blocks, as required. (Blocks + spindle = tare). (4) A straightedge, rule and spirit level.

B. Force Measurement Method (1) Locate the chord line at the inboard end of the surface by placing a straightedge at the inboard end so that one edge bisects the center of the hinge point and the trailing edge. (2) Mark the chord line.

57-50-00

Page 201 May 1/11

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (3) Place a small platform scale under the trailing edge if the control surface is tail heavy or under the leading edge if it is nose heavy. (4) Place the upper end of the spindle under the trailing edge of the tail-heavy surfaces or under the leading edge of the nose-heavy surfaces. The spindle must be vertical throughout the balancing procedure. Hold a spirit level against the marked chord line and level it by extending or contracting the spindle, or by using blocks and shims under the spindle. (5) Measure the perpendicular distance from the hinge centerline to the point supported by the spindle. Ensure that the spirit level and rule are removed from the surface and read the reaction on the beam balance. (6) Calculate the static underbalance moment M from the formula: M = D (R-T) pound-inches where, D = Perpendicular distance from the hinge centerline to the spindle point (inches). R = Reaction (Pounds) read from the beam balance. T = Tare, i.e. spindle plus levelling blocks or shims on the scale platform (Pounds). EXAMPLE: D is 10.0 inches, R is 7.680 lb. and T = 6.780lb. M = 10.0 (7.680 - 6.780); M = 9.00 pound-inches. M is within the prescribed range which is satisfactory. If M is not within the prescribed range, refer to Step (9) under BALANCING PROCEDURE COUNTERBALANCING METHOD in this section.

C. Equipment Required To Perform Check Balancing By Counterbalancing Method (1) A stand with knife-edge supports as illustrated in Figure 201. The knife edges must be in the same horizontal plane. (2) A paper cup or similar lightweight container. (3) Approximately one pound of lead shot. (4) A certified beam balance weighing device calibrated in units of 0.01 pound or less. (5) A straightedge, ruler, and spirit level.

D. Counterbalancing Method (1) Locate the chord line by placing a straightedge at the inboard end of the aileron assembly so that one end is on the trailing edge and the other end is centered on the leading edge. Mark the chord line with a suitable marker such as a grease pencil, then remove the straight edge. (2) Secure the trim tab (LH Only) in its neutral position with a small piece of masking tape. (3) Fit the correct size bolts in the hinge brackets and mount the aileron on the knife-edge supports. Ascertain that the aileron is free to rotate about the hinge line.

Page 202 May 1/11

57-50-00

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (4) To determine if adjustment weights should be added or removed, suspend a paper cup from a point near the center of the aileron trailing edge. Use a short length of small diameter string secured to the surface with a piece of masking tape as illustrated in Figure 201. The cup must be free to hang vertically. (5) Add small quantities of lead shot to the cup until the aileron balances with the chord line level. Check this by holding the spirit level aligned with the marked chord line. (6) Carefully measure the perpendicular distance D within 0.1 inch from the hinge line to the point of suspension of the cup. (7) Remove the cup, contents, and string, then weigh them to within 0.05 pound. NOTE: Since any weighing error is magnified by the distance D, weighing is most important and must be done carefully on scales that are certified for accuracy. (8) Calculate the static balance as follows: (a) The weight of the cup and contents is designated by W. (b) The distance between the centerline of the hinge and suspension point of the cup is designated by D. (c) The over- or underbalance moment is designated by M. (d) M = W x D (e) The following is a typical example of a balancing calculation: Assume the aileron is overbalance (nose heavy) and the paper cup was suspended from the trailing edge. Assume that the aileron balances with the chord line level at W = 0.900 pound and D = 10.0 inches, then... M = 0.900 x 10.0 M = 9.00 pound-inches. (The product of W x D must be accurate to within 0.05 pound-inches.) In this instance, M is within the required static balance range and is therefore acceptable. (9) Adjustment weights must be added or removed and the balance rechecked if the static balance does not fall within the prescribed limits 8.80 to 10.00 pound-inches nose heavy. NOTE: If the balance is not within the limits stated above, lead adjustment weights may be added or removed from the forward end of the inboard main leading edge closure rib to attain the required balanced condition. A steel cover plate must be installed over the adjustment weights. If necessary to meet the conditions of balancing, one or two cover plates may be used without the addition of weights, or two cover plates in addition to the weights may be used. Refer to Table 201 or Table 202.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 201 Aileron Balancing

Page 204 May 1/11

57-50-00

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Table 201 COVER PLATES, ADJUSTMENT WEIGHTS, AND SCREWS USED FOR AILERON BALANCING (EFFECTIVITY: UA-1 AND AFTER; UB-1 AND AFTER) Part Number of Cover Plate(s) (2 Maximum)

Part Number of Weight(s) (6 Maximum)

Number of Weights Used

Part Number of Screws (4 each) Required with One Cover

Part Number of Screws (4 each) Required with Two Covers

101-130001-219

101-130001-201

None

MS27039-1-08

MS27039-1-09

101-130001-219

101-130001-201

1

MS27039-1-09

MS27039-1-10

101-130001-219

101-130001-201

2

MS27039-1-10

MS27039-1-11

101-130001-219

101-130001-201

3

MS27039-1-11

MS27039-1-12

101-130001-219

101-130001-201

4

MS27039-1-12

MS27039-1-13

101-130001-219

101-130001-201

5

MS27039-1-14

MS27039-1-14

101-130001-219

101-130001-201

6

MS27039-1-15

MS27039-1-15

Table 202 COVER PLATES, ADJUSTMENT WEIGHTS, AND SCREWS USED FOR AILERON BALANCING (EFFECTIVITY: UC-1 AND AFTER) Part Number of Cover Plate(s) (2 Maximum)

Part Number of Weight(s) (11 Maximum)

Number of Weights Used

Part Number of Screws (4 each) Required with One Cover

Part Number of Screws (4 each) Required with Two Covers

118-130000-127

118-130000-129

1

MS27039-1-08

MS27039-1-09

118-130000-127

118-130000-129

2

MS27039-1-09

MS27039-1-10

118-130000-127

118-130000-129

3

MS27039-1-10

MS27039-1-11

118-130000-127

118-130000-129

4

MS27039-1-11

MS27039-1-12

118-130000-127

118-130000-129

5

MS27039-1-12

MS27039-1-14

118-130000-127

118-130000-129

6

MS27039-1-14

MS27039-1-15

118-130000-127

118-130000-129

7

MS27039-1-15

MS27039-1-15

118-130000-127

118-130000-129

8

MS27039-1-15

MS27039-1-15

118-130000-127

118-130000-129

9

MS27039-1-16

MS27039-1-16

118-130000-127

118-130000-129

10

MS27039-1-17

MS27039-1-17

118-130000-127

118-130000-129

11

MS27039-1-18

MS27039-1-18

57-50-00

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CHAPTER 61 - PROPELLER TABLE OF CONTENTS SUBJECT

PAGE

GENERAL INFORMATION 61-10-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Settings . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Propeller . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Hub Lubrication . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202 Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 207 Dynamic Balancing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 207 Test Preparation (UA-3, UB-1 thru UB-74, UC-1 thru UC-150 without Kit 114-9032-1 Installed) . . . . . 207 Test Preparation (UA-3, UB-1 thru UB-74, UC-1 thru UC-150 with Kit 114-9032-1 Installed and UC-151 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 208 Test Procedure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 210 Ground Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 210 Flight Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 211

PROPELLER CONTROLLING 61-20-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Primary Governor . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Overspeed Governor . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Low Pitch Stop . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4 Ground Fine Stop System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4 Reversing System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4 Synchrophaser . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Governor . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Operational Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202 Feather Detent Adjustment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 203 Overspeed Governor . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 205 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 205 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 205 Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 206

PROPELLER AUTOFEATHERING 61-21-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . High Pressure Switch . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Low Pressure Switch . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Annunciation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Control and Arming-Light-Out Relays . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

61-CONTENTS

1 1 1 1 1 2 2 2

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CHAPTER 61 - PROPELLER TABLE OF CONTENTS (CONTINUED) SUBJECT PAGE Troubleshooting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 101 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 101 Operational Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 101 Auto-Feather System Troubleshooting, Engines Off . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 104 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Dump Valve . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Pressure Switch . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202 Adjustment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202

PROPELLER SYNCHROPHASER 61-22-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Troubleshooting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 101 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 101 Synchrophaser . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 101 Functional Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 101 Synchrophaser Checks . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 102 Governor Speed Biasing Coil Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 104 Magnetic Pickup Voltage Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 105 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Pickup . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Installation and Adjustment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201

PROPELLER INDICATING 61-40-00 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Throttle Switch . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Adjustment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201

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List of Effective Pages CH-SE-SU

PAGE

DATE

61-LOEP

1

Nov 1/10

61-CONTENTS

1 and 2

Nov 1/09

61-10-00

1 201 thru 215

Nov 1/09 Nov 1/09

61-20-00

1 thru 5 201 thru 206

Nov 1/09 Aug 1/10

61-21-00

1 thru 5 101 thru 105 201 and 202

Nov 1/09 Nov 1/10 Aug 1/10

61-22-00

1 101 thru 109 201

Nov 1/09 Nov 1/10 Nov 1/09

61-40-00

201

Nov 1/09

C4

61-LOEP

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

PROPELLER GENERAL INFORMATION DESCRIPTION AND OPERATION

61-10-00 00

1. GENERAL The 1900 airliner uses two four blade propellers of composite construction with standard hubs. Each constant speed, full feathering, reversible propeller is controlled by engine oil from a single acting engine driven governor. Backing up the engine driven governor is an overspeed governor and a fuel topping governor, also known as the power turbine governor, which is a component within the normal or primary governor. A servo piston mounted on front of the propeller spider hub moves the propeller blades through links connected to the trailing edge. Centrifugal counterweights on each blade, in conjunction with a feathering spring on the servo piston, increase pitch (decrease rpm) to the feathered position as governor oil pressure is relieved. The feathering spring completes the feathering operation when centrifugal twisting moment is lost as the propeller stops rotating. To further expedite feathering, the normal feathering mechanism is backed by an automatic feathering system that provides a means of immediately dumping oil from the propeller governor to enable the feathering springs to start feathering the propeller blades as soon as engine torque meter oil pressure drops below 3.1 ± 0.6 psi at power settings of 85 to 90% N1. The servo piston of the reversing propeller is connected by four spring loaded sliding rods to a low pitch stop collar (feedback ring) located behind the propeller. The movement of the low pitch stop collar is transmitted by one carbon block through a reversing lever and connecting linkage to a Beta valve, which is positioned to maintain the blade angle while propeller RPM is lower than that of the governor pilot valve as selected by the control. The reversing lever is also connected to the fuel topping governor to limit propeller RPM in the reverse position. A push-pull cable extends from the reversing lever aft to a control Beta cam box connected to the power lever control and fuel control unit. Movement of the controls is transmitted through the Beta cam box and interconnecting linkage to the fuel control unit and governors to regulate propeller speed and pitch.

A. Settings Full Feathered....80° ± 0.5° at the Angle 42 inch station Mechanical Reverse....-14° ± 0.50° Pitch Stop The angle at which the prop dome just touches the 4 low pitch stop rod nuts which move the mechanical low pitch stop collar is 8.0 ± 1° when the blades are held toward the decrease position at the 42 inch station. Low RPM....1400 (at detent) Maximum RPM....1700 (at take off) NOTE: When a propeller synchrophaser is installed there is no change in governor setting procedures. The synchrophaser must be turned OFF while making governor settings.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

PROPELLER MAINTENANCE PRACTICES

200200

1. PROPELLER A. Hub Lubrication (1) Follow all the Steps of the 100 hour inspection procedure. CAUTION: Lubrication procedures must be followed correctly to maintain accurate dynamic balance of the propeller blade assemblies and hub assembly. Improper lubrication practices can create grease leaks, out-of-balance conditions, and/or internal corrosion. Use approved lubricants only. (2) Proceed as follows to lubricate the propeller assembly: CAUTION: Failure to remove one of the fittings on each of the blades for venting purposes will result in either leaking past the blade shank-to-hub seal or forcing grease into the hub cavity. Both conditions will result in grease leakage and can affect propeller operation both in pitch change and balance. (a) Remove the grease fittings from the engine side of the hub unit. NOTE: The same amount of grease must be applied to all blades of a propeller assembly at the time of lubrication. Failure to apply the same amount of grease to each blade may result in balance changes which result in increased vibration during operation. (b) Add an equal number of pumps of grease through each of the grease fittings on the cylinder side of the hub unit (Ref. Chapter 12-20-00, FLIGHT COMPARTMENT ENGINE CONTROLS AND PROPELLER LUBRICATION). NOTE: When pumping the grease into the fitting, observe the grease coming out of the opening. Lubrication is complete when the grease emerges in a steady flow with no air pockets or moisture and has the color and texture of the new grease. (c) Work a probe (loop) of wire in and out of the openings on the engine side of the hubs to help release air pockets in the grease. (d) Install all grease fittings, making sure that each grease fitting is properly installed and that the ball of each fitting is seated properly against the opening of the fitting. NOTE: The interval between lubrication is an important factor in minimizing internal corrosion. Observe the proper lubrication intervals of 300 hours. (e) Make an entry in the Log Book verifying that these inspection and lubrication procedures have been completed.

B. Removal (1) Remove the upper and lower forward engine cowling (Ref. Chapter 71-10-00, COWLING REMOVAL).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (2) Remove the nuts, washers and screws securing the propeller deicer brush block assembly to the engine bracket (Ref. Chapter 30-60-00, PROPELLER DEICER BRUSH REPLACEMENT). CAUTION: To prevent damage to the deicer brushes after removal, tape them in place after removing the brush block. (3) Remove the screws and fiber washers from around the rear circumference of the spinner dome. Carefully remove the spinner dome and store it to prevent damage. (4) Disconnect the engine beta linkage and carbon block assembly from the feedback ring. (5)

Support the propeller assembly with a suitable sling and mobile hoist with a capacity of 300 pounds or more.

(6) Install the beta system puller (P/N CST2987) or feedback ring puller (P/N TK1573-918-1) (Ref. Figure 202). Pull the feedback ring forward to expose the eight double hex headed propeller mounting bolts and washers. CAUTION: To avoid damaging the propeller, ensure the tool is not cocked. Avoid bending or otherwise damaging the spring loaded rods and the feedback ring (Ref. Figure 202). (7) Remove the safety wire from the propeller mounting bolts. (8) Remove the eight bolts and washers securing the propeller, and carefully remove the propeller assembly from the engine. NOTE: If the propeller is being overhauled, discard the mounting bolts and washers as new bolts and washers are to be installed after propeller overhaul. (9) Decompress and remove the beta system compression tool. (10) Remove and discard the propeller mounting packing seal.

C. Installation NOTE: For correct propeller synchrophaser pickup installation for left hand or right hand propellers (Ref. 61-22-00). (1) Place the new packing seal over the engine shaft. (2) Pull the low pitch stop collar fully forward with the beta system puller (P/N CST2987) or the feedback ring puller (P/N TK1573-918-1) (Ref. Figure 202). CAUTION: To avoid damaging the propeller, make sure that the tool is not cocked. Take the precautions necessary to avoid bending or otherwise damaging the spring loaded rods and the low pitch stop collar (brass ring). NOTE: The B3001-2 low pitch stop collar is not installed on new propellers delivered from the factory as a precaution against damage during shipment. Only one of the FOUR jam nuts is installed on one of the B3002-2 rods. This one nut is locked in position on the rod with thread locking compound (59, Table 1, Chapter 91-00-00) for proper positioning of the B3001-2 collar. (3) When installing a propeller that was previously removed, proceed to Step (4). If a new propeller is being installed, proceed as follows:

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (a) Install the two jam nuts on the two B3002-2 rods. (b) Loosen the nut on the front end of each rod. (c) Install the B3001-2 collar on the ends of the four B3002-2 rods. The rods can be screwed into the collar by applying a thin wrench to the flats provided on the rods. DO NOT TIGHTEN THE JAM NUTS AT THIS TIME. (d) Install the propeller on the engine as indicated in Steps (4) thru (7) below. (e) Tighten the one jam nut that is locked in position on the rod to 12 foot-pounds. (f) Using a dial indicator of the type used by machinists, check the runout of the low pitch stop collar. The total runout should not exceed 0.010 inch. In the event of excessive runout, adjust the three remaining B3002-2 rods and jam nuts as required to obtain the specified runout. (g) Torque the three remaining jam nuts to 12 foot-pounds. NOTE: In the event that the black colored jam nut (coated with Loctite) becomes loose by accident or mistake, it can be repositioned at the exact point on the rod by measuring the distance between the outer surface of the low pitch stop collar (brass ring) and the outer surface of the propeller flange (Ref. Figure 203). The distance must be 1.88 inch. (h) Torque the nut on the front of each B3002-2 rod to 12 foot-pounds. (4) Pull the low pitch stop collar fully forward with beta system puller (P/N CST2987) or the feedback ring puller (P/N TK1573-918-1) (Ref. Figure 202). (5) Install new prop shim between propeller and engine shaft as per P&WC S.B. 13116R2. (6) Install the propeller on the engine by inserting the two mounting studs on the propeller into the mounting holes on the drive shaft of the engine. (7) Coat the entire shaft of the eight propeller mounting bolts, threaded and unthreaded section, and the washer with antiseize thread compound, graphite petrolatum (76, Table 1, Chapter 91-00-00). (8) Install the eight propeller mounting bolts and washers, with the chamfered edge of the washer toward the bolt head, through the engine mount flange into the propeller mount flange. (9) Using a torque adapter (26 or 27, Table 7, Chapter 91-00-00) torque all bolts to 40 foot-pounds and then to 80 foot-pounds following torque sequence A (Ref. Figure 201).

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(10) Final torque all bolts to 100 to 105 foot-pounds following torque sequence B (Ref. Figure 201). After final torque, lockwire (178, Table 1, Chapter 91-00-00) all mounting bolts (Ref. Chapter 20-07-00, LOCKWIRE). CAUTION: The torque on the bolts shall not exceed 105 foot-pounds. (11) Remove the feedback ring puller and connect the propeller reversing lever to the propeller control linkage. CAUTION: With the carbon brush block held against the low pitch stop collar, check the clearance between the low pitch stop collar and the metal retaining clip of the carbon block. If at any point this clearance is 0.005 inch or less, replace the carbon block to prevent damage to the low pitch stop collar by the retaining clip. Give clearance between carbon brush block and brass ring wall at the nearest point. (12) Install the deicer brush block assembly (Ref. Chapter 30-60-00, PROPELLER DEICER BRUSH REPLACEMENT). (13) Check the propeller reversing linkage on the front end of the engine for proper rigging. (14) Position spinner dome and install the screws and fiber washers around the circumference. (15) Install lower forward engine cowling (Ref. Chapter 71-10-00, COWLING INSTALLATION). (16) Perform the necessary engine run up checks as outlined in the applicable Pilot's Operating Handbook. (17) Install upper forward engine cowling (Ref. Chapter 71-10-00, COWLING INSTALLATION).

Figure 201 Propeller Mounting Bolt Torquing Sequence

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 202 Feedback Ring Puller

61-10-00

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Figure 203 Beta Control Feedback Ring

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D. Inspection Check propellers for condition, and attachment; check condition of mechanical feedback ring and stop rods and springs; check reversing linkage for evidence of binding and security of attachment; check the valve and plumbing for security of attachment and leakage (Ref. Chapter 5-20-00).

2. DYNAMIC BALANCING NOTE: Dynamic balancing frequency should be based on individual operator service experience. Operators are encouraged to initially implement a dynamic propeller balance program at intervals not to exceed 1200 service hours. Thereafter dynamic propeller balancing should be accomplished at appropriate intervals, based on individual operator’s field experience or any time propeller maintenance is accomplished which could effect propeller balance; such as blade repairs. Balancing is also recommended after lubrication. Dynamically balanced propellers shall be rebalanced whenever the engine and/or propeller is changed. Dynamic balancing of the propellers must be accomplished ONLY by personnel who are totally familiar with the applicable service equipment and data. Personnel must be thoroughly instructed as to the use and care of the test equipment being used. Different brands of dynamic balancing equipment are available for purchase and may be used as directed by the manufacturer of the equipment. The equipment listed below is almost totally automatic and, when connected properly, will furnish Step by Step instructions on the readout screen as test procedures are performed. UC-151 and After are equipped with permanent hardware provisions installed for propeller dynamic balancing. Two kits are available for propeller balancing on UA-3, UB-1 thru UB-74 and UC-1 thru UC-150. Kit 114-9032-1 has permanent dynamic balancing provisions similar to installations on UC-151 and After. For the airplanes which do not have the Kit 114-9032-1 installed, Kit 114-9035-1 provides temporary provisions to dynamically balance the propellers. It is recommended that the manufacturer's instructions for the equipment being used be followed closely during the balancing procedure. The following balancing procedure has been developed for use with the Chadwick-Helmuth Model 8500 Vibration Analyzer (8, Table 7, Chapter 91-00-00). When using a vibration analyzer other than a Chadwick-Helmuth Model 8500, make adjustments to the following procedures based on the equipment manufacturer’s instructions on the use of this equipment.

A. Test Preparation (UA-3, UB-1 thru UB-74, UC-1 thru UC-150 without Kit 114-9032-1 Installed) Kit 114-9035-1 provides information on how to modify the cowling and to temporarily route the cables on airplanes without Kit 114-9032-1 installed. (1) Remove the upper forward cowling assembly. (2) Following Chadwick-Helmuth procedures, install the velocimeter and propeller R.P.M. sensor to the forward lower stud on the tach generator pad with the arrow on the sensor pointing down. (3) Route the cables into the flight compartment through the lower forward cowling as shown on the kit drawing. If the cowling is not equipped to route the cables properly, install an access panel per Kit 114-9035-1. (4) Clamp the cables as required in the cowling area and tape the cables to the airplane skin, keeping the cables as flat as possible. Use 2 inch wide tedlar tape (112, Table 1, Chapter 91-00-00) or equivalent for this purpose.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (5) Install the upper forward cowling assembly if the balancing data is to be recorded in flight.

B. Test Preparation (UA-3, UB-1 thru UB-74, UC-1 thru UC-150 with Kit 114-9032-1 Installed and UC-151 and After) Kit 114-9032-1 provides information and parts required to install permanent provisions similar to those used on UC-151 for dynamically balancing the propeller assembly. NOTE: The Kit 114-9023 provides information to relocate the propeller deice brush block to a location that will reduce the dirt and grime on the brush block. This kit MUST be installed prior to, or in conjunction with, Kit 114-9032-1. (1) Remove the upper forward engine cowling. (2) Unplug the connector and remove the dummy vibration sensor from the bracket on the tach generator pad (Ref. Figure 204). (3) Install the vibration sensor furnished with the balancing equipment. Ensure that the arrow on the sensor points down when the sensor is secured to the bracket. (4) Plug the connector into the vibration sensor, and install the upper cowling assembly. NOTE: This installation is identical for both engines. (5) Place the Vibration Analyzer in the flight compartment. (6) Connect the adapter cable (Chadwick-Helmuth P/N 10390) to the receptacle in the sidewall behind the copilot's seat. (7) Connect the left engine leads from the adapter cable to channels 1 and A and the right engine leads from the adapter cable to channels 2 and B on the switch box (Chadwick-Helmuth P/N 9110). (8) Connect the patch cord (furnished with the 9110 switch box) to the 9110 switch box and to the input connection on the Vibration Analyzer.

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Figure 204 Vibration Sensor Connection

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3. TEST PROCEDURE NOTE: Propeller maintenance, such as blade repairs, lubrication, painting, etc., must be accomplished prior to dynamic balancing procedures. Engine and/or propeller changes will require balancing. Propeller dynamic balancing procedures may be accomplished on the ground or in flight. Generally, better results are obtained with the in flight procedure since vibration readings are taken when the propeller is operating at typical cruise flight blade angle and aerodynamic loads. Readings taken during high power ground running may not be as consistently accurate due to blade angle and/or ground buffeting. NOTE: Hawker Beechcraft Corporation recommends that the propeller be balanced to 0.2 inch per second (IPS) or less when in flight and 0.1 IPS or less during ground operation.

A. Ground Test (1) Perform a normal engine start. Taxi to the run up pad and position the airplane crosswind to perform the ground balancing procedure. (2) Ground run the downwind engine at the most frequently used parameters for cruise power and propeller RPM. Press START on the analyzer and the analyzer will automatically acquire a reading and calculate a propeller solution. Press PRINT on the analyzer to print the solution. (3) Turn the airplane around so the opposite engine will be on the downwind side. Again run the downwind engine at the most frequently used parameters for cruise power and propeller RPM. (4) Press SEL CHART on the analyzer (as indicated by the cursor on the analyzer display) and switch to the opposite propeller on the switch box. Press START on the analyzer and the analyzer will automatically acquire a reading and calculate a propeller solution. Press PRINT on the analyzer to print the solution. CAUTION: Do not alter static balance to achieve dynamic balance. (5) Compute the weight amounts and locations required for dynamic balance of the propeller in accordance with Figure 205 and Table 201. NOTE: Dynamic balancing weights (washers) may be attached to the spinner bulkhead; twelve possible weight locations may be utilized (Ref. Figure 206). The balancing weights should be located on the engine side of the bulkhead and should not exceed 0.9 ounce per location. Nut plates, riveted to the propeller side of the bulkhead, may be used in place of nuts for screw attachment. (6) Adjust the weight installation until satisfactory balance readings are confirmed. Ensure that the decal is installed on the aft side of the spinner bulkhead to indicate that the propeller has been dynamically balanced. NOTE: Table 201 indicates the combined weight of the balance weights (washers) and the attaching screws. (7) Remove the forward engine cowling and disconnect and remove the vibration sensor from the tach generator bracket. Connect and install the dummy sensor and install the forward engine cowling. (8) Make appropriate log book entries denoting that the propellers have been dynamically balanced. Record the amount of weight added or removed at each location and the IPS of the final balance.

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B. Flight Test (1) Fly the airplane at the altitude and cruising speed at which the airplane is normally operated. (2) When the propeller speed has stabilized at cruise RPM, press START on the analyzer and the analyzer will automatically acquire a reading and calculate a propeller solution. Press PRINT on the analyzer to print the solution. (3) Press SEL CHART on the analyzer (as indicated by the cursor position on the analyzer display) and switch to the opposite propeller on the switchbox. Press START on the analyzer and the analyzer will automatically require a reading and calculate a propeller solution. Press PRINT on the analyzer to print the solution. CAUTION: Do not alter static balance to achieve dynamic balance. (4) Compute the weight amounts and locations required for dynamic balance of the propeller in accordance with Figure 205 and Table 201. NOTE: Dynamic balancing weights (washers) may be attached to the spinner bulkhead; twelve possible weight locations may be utilized (Ref. Figure 206). The balancing weights should be located on the engine side of the bulkhead and should not exceed 0.9 ounce per location. Nut plates, riveted to the propeller side of the bulkhead, may be used in place of nuts for screw attachment. (5) Adjust the weight installation until satisfactory balance readings are confirmed. Ensure that the decal is installed on the aft side of the spinner bulkhead to indicate that the propeller has been dynamically balanced. (6) Disconnect and remove all of the test equipment from the airplane. (7) Remove the forward engine cowling and disconnect and remove the vibration sensor from the tach generator bracket. Connect and install the dummy sensor and install the forward engine cowling. (8) Make appropriate log book entries denoting that the propellers have been dynamically balanced. Record the amount of weight added or removed at each location and the IPS of the final balance. Table 201 Allowable Propeller Balance Weight NO. OF WEIGHTS

AMT. OF WEIGHT IN GRAMS

0

3.2

1

8.0

2

12.4

3

17.2

4

21.5

nut

0.8 NOTE

Table 201 indicates the combined weight of the balance weights (washers) and the attaching screws.

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Figure 205 Balance Chart

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Figure 206 Spinner Bulkhead Weight Locations

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PROPELLER PROPELLER CONTROLLING DESCRIPTION AND OPERATION

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1. GENERAL A. Primary Governor An increase in the flow of oil regulated by the primary governor passes through the oil transfer housing and the hollow center of the propeller shaft to move the propeller blades toward the low pitch (high rpm) hydraulic stop and reverse positions. The primary governor, mounted on top of the gear reduction housing, regulates N2 by varying propeller blade pitch. The governor consists of a gear-type oil pump with flyweights mounted on a rotating head, and a spring-loaded pilot valve that regulates the flow of oil to and from the propeller servo piston. The position of the pilot valve is controlled by the rotating flyweights, in conjunction with the spring-load imposed by the external speed control lever. When the engine rpm drops to an underspeed condition below the control setting, spring force overcomes flyweight force to lower the pilot valve plunger and open the port in the governor drive gear shaft, through which oil flows to the propeller servo piston and decreases blade angle. The decrease in pitch decreases the load on the engine. The resultant increase in engine rpm also increases the centrifugal force of the rotating flyweights, which then lifts the pilot valve plunger to cover the port in the governor drive gear shaft and shut off the flow of oil to the propeller. The forces exerted on the pilot valve plunger by the flyweights and speeder spring then balance to initiate the on-speed cycle of the governor. An overspeed condition occurs with a decrease in propeller load or with movement of the propeller control to decrease rpm. Flyweight force then overcomes speeder spring force and raises the pilot valve plunger to open the port through which oil drains from the propeller through the governor to the sump. The load on the engine increases and rpm drops as the counterweights and feathering spring increase propeller pitch. The pilot valve then centers in the governor drive gear shaft to block the flow of oil to and from the propeller as governor flyweight and speeder spring force reach a state of equilibrium. The primary governor contains a fuel topping governor which has an airbleed orifice that opens to change the effect of the fuel control valve in the Fuel Control Unit (FCU) to reduce fuel, power, and N1. During normal flying, instantaneous changes in the atmosphere density cause high N2 rpm. When the primary governor senses the overspeed condition, the governing section moves an airbleed lever to open the airbleed orifice. When the propeller control lever is at high rpm, the airbleed will open as N2 increases 6% above the selected value. When the propeller control lever is at low rpm, the airbleed will open when N2 reaches 4% below the selected value. When the engine control lever is moved into reverse, the Beta cambox push/pull cable pulls the reset linkage aft to hold the airbleed orifice open (Ref. Figure 1).

B. Overspeed Governor A propeller overspeed governor, mounted on the left side of the reduction gear housing, acts as a safeguard against propeller overspeed should the primary governor fail. The overspeed governor regulates the flow of oil to the propeller pitch-change mechanism by means of a flyweight and speeder spring arrangement similar to that of the primary governor. The overspeed unit governs at 104% N2 speed (approximately 1768 rpm). Since it has no mechanical controls, the overspeed governor is equipped with a testing solenoid that resets the normal overspeed setting to approximately 1564 rpm for ground testing (Ref. Figure 2).

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Figure 1 Propeller Governor

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Figure 2 Propeller Governor System

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C. Low Pitch Stop Propeller blade angle will be on the low pitch stop anytime the propeller rpm is lower than that selected. With the engine condition levers on low setting, the gas turbine section N1 cannot drive the propeller rpm N2 to the setting selected by the propeller control lever. At that point, the propeller rpm falls below the range of the governor and reaches the low pitch stop. As N1 power is reduced, N2 is reduced. The primary governor senses this underspeed condition, lowers the pilot valve in the governing section to reduce pitch (reduce the load). The decrease in blade angle allows the propeller rpm N2 to increase and match the rpm as selected in the cockpit. The Beta ring on the propeller gives mechanical feedback of the blade position to the reverse lever of the Beta valve. When the blade pitch reaches the desired position as determined by the rigging, the reverse lever moves the Beta valve outward to stop the oil flow to the propeller and hold the blade at that pitch. At that point the governor has no effect because the Beta valve prevents oil from going to or leaving the propeller.

D. Ground Fine Stop System The ground fine stop system uses an electrical solenoid mounted on the front of the reversing push/pull cable to limit the propeller blade angle to 0 to 7°. The solenoid is connected to the propeller reversing lever by means of a slotted clevis which allows the reversing lever to be pulled aft, resetting the Beta valve. The electrical solenoid can be energized by two ground paths. One is through the RH landing gear squat switch. The other occurs by pulling the power levers to the ground idle fine switch. The solenoid in each case energizes and pulls the reverse lever of the Beta valve aft to reset the blade angle.

E. Reversing System When the power lever controls are lifted for placement in the reversing range, the power levers actuate the pedestal switches that break the circuit of the secondary low pitch stop system. This enables the propeller servo piston to reverse the pitch of the propeller blades with engine oil pressure from the governor sump. The power lever controls are connected to the engine control Beta cam box which is also linked to the fuel control unit. A push-cable extends from the Beta cam box forward to the reversing lever to which is attached the brush that rides in the propeller brass feedback ring. Movement of the power control lever is transmitted through the Beta cam box and interconnecting linkage to the fuel control unit. Propeller control levers to the governors regulate propeller speed and pitch.

F. Synchrophaser The propeller synchrophaser automatically matches the RPM of both propellers and positions the propellers at a pre-set phase relationship between the blades of the left and the right propellers. This phase relationship is designed to decrease cabin noise and is not adjustable by the pilot. The control box senses pulses which are generated by pickups mounted on both engines at identical locations in relation to the engine centerline. Steel targets, mounted on the propeller spinner bulkheads, provide the pulse reference for the pickups. The magnetic pickups operate on the magetoelectric principle: when the steel target passes in close proximity to the magnetic field of the pickup, an alternating current is induced in the pickup. The magnitude of the peak-to-peak voltage should ideally be a minimum of 4 volts. The control box does not respond to amplitude of the input signals, but to phasing of the waveforms. A specific amount of voltage change, approximately 0.7 volts, is necessary to trigger the control box. Maintaining a time-phase relationship is the control box’s only priority. The control box senses the

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL pulses from each pickup and attempts to superimpose the waveforms by trimming the speed of each propeller. Current flow through each trim coil (independent of polarity) either pulls the governor flyweights inward, or pulls on a disc attached to the pilot valve to simulate an under speed condition. Speed trim of the propellers is accomplished by the control box with correction commands to each propeller governor. The character of the correction commands is in the form of pulse width-modulated 28 volt direct current: the duration of the pulse of current is regulated to produce the proper amount of speed trim. The amount of speed trim per pulse width modulation by the control box is a function of the governor and will always be within a very narrow range (holding range, approximately 25 ± 2 RPM). The governor servo can increase, but never decrease, the speed set by the propeller control lever. The RPM of one propeller will follow the changes in RPM of the other propeller over the predetermined holding range of the governor. This limited holding range prevents either propeller from losing more than limited RPM if the RPM of the other propeller is manually reduced, such as in power changes or propeller feathering, while the synchrophaser is on. The synchrophaser system is controlled through a toggle switch placarded PROP SYNCH-ON-OFF. To operate the system, synchronize the propellers manually and set the RPM of each propeller to within 10 RPM of each other and turn the syncrophaser on. To change RPM, adjust both propellers at the same time. This will keep the setting within the holding range of the system. If the synchrophaser is on, but will not synchronize propellers, the propeller speeds are not within the limits required for the system to assume control (outside capture range). Turn the synchrophaser off, synchronize the propellers manually, then turn the synchrophaser on. Engagement will automatically occur when the relative phase of the input signals is within 30 rotational degrees of the control box internal phase setting. Engagement usually occurs within seconds of system turn-on; however it may take as long as 30 seconds for the input signals to drift within phasing range. When input conditions are satisfied, the system engages. Both propeller speeds are increased by one-half the holding range of the system. Each control box output signal is the inversion of the other. As propeller RPM is increased on one side, it is decreased on the other. Syncrophaser engagement or operation can never reduce propeller RPM below manual speed settings except in rare instances where torque or temperature limiting is approached.

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PROPELLER PROPELLER CONTROLLING MAINTENANCE PRACTICES

200200

1. GOVERNOR A. Removal (1) Disconnect the lever (1) from the Beta control valve (2) by removing the cotter pin, washer and clevis pin (Ref. Figure 201). (2) Disconnect the lost motion link (3) from the reset arm by removing the cotter pin, nut, washer, spacer and bolt. (3) Disconnect the PY tube (5) from the governor (19) by removing the lockwire and disconnecting the coupling nut. (4) Disconnect the speed control rod (6) from the propeller governor lever by removing the lockwire, bolt, and washer. (5) Remove bolts (7), plate (8), oil tube (9) and preformed packing (10). (6) Remove the self-locking nuts (17) using the retaining nut wrench (Pratt & Whitney P/N 30114-16). (7) Remove the washers (18), governor (19) and gasket (20). Discard the gasket. (8) If the original governor is to be replaced with a new part, remove the following: (a) Note angular position of the elbow (12), then loosen the jam nut and remove the elbow, jam nut, preformed packing (13) and back-up ring (14). (b) Remove the straight adapter (16) and preformed packing (15).

B. Installation (1) If a new or replacement is to be installed, assemble the adapter and elbow as follows: (a) Install new preformed packing (15) on straight adapter (16). Screw the adapter into the body of the governor, tighten and torque the adapter to 65 to 75 inch-pounds and lockwire the adapter (Ref. Figure 201). (b) Install the jam nut (11), back-up ring (14) and preformed packing (13) on elbow (12). Refer to Chapter 70-06-00, Standard Practices, of the Pratt & Whitney Maintenance Manual P/N 3032842 for elbow assembly procedures, and install the elbow at the same angle as noted at disassembly. Tighten the jam nut and torque to 70 to 80 inch-pounds. Safety lockwire the nut and elbow. (2) Install new gasket (20) on reduction gearbox studs, with raised surface uppermost. (3) Install propeller governor (19), while rotating the propeller shaft so that the splines engage correctly. Ensure proper engagement by checking that there is no gap between the flange and the governor.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (4) Secure the governor with the washers (18) and self-locking nuts (17). Tighten the nuts and torque to 170 to 190 inch-pounds using the retaining nut wrench (Pratt & Whitney P/N 30114-16). (5) Install a new Preformed packing (10) on oil tube (9). Connect upper end of tube to the elbow (12) and insert the lower end of the tube into the reduction gearbox boss. Secure lower end with plate (8) and two bolts (7). Tighten the bolts and torque to 36 to 40 inch-pounds and lockwire. Tighten the tube coupling nut and torque to 90 to 100 inch-pounds and lockwire. (6) Connect the PY governor air tube to the adapter (16), tighten the coupling nut, torque the coupling nut 90 to 100 inch-pounds and lockwire. (7) Connect the speed control rod (6) to the propeller governor lever using the bolt and washer. Torque to 20 to 30 inch-pounds and install safety wire. (8) Connect the lost motion link (3) to the reset arm (4) using the bolt, spacer, washer and nut. Torque to 12 to 18 inch-pounds and install the cotter pin. (9) Connect the lever (1) to the governor Beta cam clevis by installing the pin, bushing, washer and cotter pin. (10) Determine the proper electrical connector from the applicable wiring diagrams and connect it to the governor. Safety wire the electrical plug. WARNING: Ensure that the proper electrical connectors are attached to the propeller governor, overspeed governor, and the autofeather low pressure switch by checking the applicable wiring diagrams. The connectors to these plugs are identical and they could be swapped. CAUTION: Excessive tightening will result in damage to threads of mating parts. (11) Check engine control rigging as necessary.

C. Operational Check Use the following checks to assess the operation of the governors: (1) Start the engines in accordance with the operator's manual. (2) Place the propeller levers in HIGH RPM position. (3) Advance the power levers until propeller tachometers stabilize at maximum rpm. (4) Pull the propeller levers aft to the detent. (5) Observe that the primary governors reduce propeller tachometer rpm readings. (6) Return propeller levers to HIGH RPM. (7) Shut down engines in accordance with the operator's manual.

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D. Feather Detent Adjustment Position the propeller lever in the full high rpm position and advance the power lever to obtain 1600 rpm. Pull the propeller lever back against the detent, then adjust the detent as required to obtain 1400 to 1425 rpm (at detent) (Ref. Figure 202). Pull the propeller lever past the detent-feathered position and observe that the propeller is in the fully feathered position.

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Figure 201 Removal/Installation of Propeller Governor

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Figure 202 Autofeathering Detents

2. OVERSPEED GOVERNOR A. Removal (1) Remove the upper forward engine cowling (Ref. Chapter 71-10-00). (2) Remove the safety wire and disconnect the electrical plugs from the governor solenoids. (3) Remove the four self-locking nuts and plain washers securing the governor and remove the governor from the left side of the reduction gear housing.

B. Installation (1) Install a new gasket on the mounting pad. (2) Apply lubricating grease (64, Table 1, Chapter 91-00-00) sparingly to the governor splined drive. NOTE: The propeller shaft may require rotating to ensure that the splines align properly. (3) Position the governor on the mounting pad and install the four plain washers and self-locking nuts. Apply a torque of 170 to 190 inch-pounds to the mounting nuts.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (4) Determine the proper electrical connectors from the applicable wiring diagrams and install. Precaution should be taken not to over tighten the electrical plugs when connecting the overspeed governor. This could cause damage to the overspeed governor solenoids. The solenoid with the receptacle on top is the speed reset solenoid used to test governor operation and the remaining solenoid is used for auto feathering. Safety wire the electrical plugs. WARNING: Ensure that the proper electrical connectors are attached to the propeller governor, overspeed governor, and the autofeather low pressure switch by checking the applicable wiring diagrams. The connectors to these plugs are identical and they could be swapped. (5) Install the upper forward engine cowling (Ref. Chapter 71-10-00).

C. Check NOTE: The overspeed unit, mounted on the left side of the reduction gear housing, is preset at the factory to govern at 104% N2 speed (approximately 1768 rpm). This setting can be reduced to 1520 to 1610 rpm for testing purposes by means of a testing solenoid valve. A second solenoid valve attached to the governor provides a means of immediately dumping oil from the unit for autofeathering. Perform the overspeed governor check as follows: CAUTION: Do not force the power levers into the full reverse position while performing this test. (1) Start the engines in accordance with the operator's manual. (2) Press the PROP GOV TEST and BUS TIE circuit breakers on the right circuit breaker panel. (3) Run the engine with the power lever set below 1700 rpm and the propeller control in the full forward position. CAUTION: Do not exceed the engine torque and ITT limits during this test. (4) Hold the PROP TEST switch on the pilot's subpanel in the OVERSPEED position. (5) Advance the power lever until propeller rpm stabilizes at 1520 to 1610 rpm. Observe increase in torque. (6) Retard the power lever to the IDLE position, then release the propeller test switch. Repeat the preceding check on the other engine. (7) Shut down engines in accordance with the operator's manual.

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PROPELLER PROPELLER AUTOFEATHERING DESCRIPTION AND OPERATION

61-21-00 00

1. GENERAL The automatic feathering system provides a means of immediately dumping oil from the propeller governor to enable the feathering springs to start feathering the propeller blades as soon as engine torquemeter oil pressure drops below a preset point, equivalent to approximately 260 foot-pounds. of torque, at power settings of 85 to 90% N1 or greater. The system is primarily intended for use during takeoff and landing. It should be kept on until the airplane has gained enough altitude that the loss of one engine and its resultant drag will not present an immediate problem to the pilot.

2. DESCRIPTION The autofeathering system is controlled through a four-pole three-position control switch with modes of ARM (the up position), OFF (the center position), and TEST (the momentary down position) (Ref. Figures 1, 2, and 3). The system is armed by placing the arming switch, located in the subpanel, in the ARM position. This closes the circuit from a 5-ampere circuit breaker in the right circuit breaker panel to the power lever switches mounted on a bracket in the control pedestal; however, the autofeathering system will remain inactive as long as the power levers are retarded below the 85 to 90% N1 position. When the power levers are advanced to the 85 to 90% N1 position, a mechanical actuator connected to each power lever will close its respective switch and complete the circuit to the high and low pressure switches mounted adjacent to the torque pressure transmitter just forward and above the left exhaust outlet on each engine. These pressure switches monitor torquemeter oil pressure.

A. High Pressure Switch The high pressure switch is a double-pole, double-throw switch that actuates to the normally closed set of contacts when the torquemeter oil pressure drops below 4.7 ± 0.4 psi and closes to the normally open set of contacts when the pressure rises again to 6.1 ± 0.5 psi. When closed to its normally open set of contacts (torquemeter oil pressure above 6.1 ± 0.5 psi), the high pressure switch provides current from the number 1 power lever switch to the opposite autofeather system annunciator while making current available to energize the opposite system's autofeather dump solenoid valve and arm-light-out relay. Should a fault cause the torquemeter pressure to drop below 4.7 ± 0.4 psi on either autofeather system, the high pressure switch on the faulted side will return to its normally closed position, thereby removing current from the opposite side to disable that system and providing current to its own feather dump solenoid control relay; consequently, the possibility of automatically feathering both propellers is precluded.

B. Low Pressure Switch The autofeather low pressure switch provides for the switching of ground to the control relay and the autofeather dump solenoid valve. The low pressure switch is a single-pole, double-throw pressure actuated switch permanently set to actuate on rising pressure to its normally open contact at 4.0 ± 0.8 psi. On falling pressure, it will return to its normally closed contact at 3.1 ± 0.6 psi. In its normally closed position (torquemeter pressure below 3.1 ± 0.6 psi), the low pressure switch completes the ground circuit to the autofeather dump solenoid valve and to its respective control relay. The normally open contact of this switch is unused.

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C. Annunciation When the autofeather control switch is in the ARM position and the engine power levers are advanced above 85 to 90% N1, the annunciator light for one side receives operating current through the high pressure switch in the opposite side system. Grounding of the annunciator lights is accomplished through the number two power lever switches. When the autofeather system on one side activates, its control relay closes, energizing its respective arm-light-out relay which removes current to extinguish the annunciator.

D. Control and Arming-Light-Out Relays The relays for the RH autofeathering system are installed on relay panel A121 (1900), A120 (1900C), located in the PCB rack beneath the center aisle floorboard. The relays for the LH autofeathering system are mounted on relay panel A119 in the same pcb rack.

3. OPERATION When the engines are operating at some power in excess of 85 to 90% N1 and the autofeather system is armed, the autofeather pressure switches will all be closed to the normally open position. Energizing current is available at one pole of the control relays for operation of the feather dump solenoid valves and the arm-light-out relays by way of the opposite system high pressure switch. The annunciators are also energized through the opposite high pressure switch in this mode. Should the torquemeter pressure begin to drop on one engine due to an engine fault, the high pressure switch on the faulted side will close at approximately 4.7 psi. At this point, the unfaulted side of the system is disabled and its annunciator extinguished because the energizing current provided through the high pressure switch of the faulted side of the system is no longer available; due to the switching of the high pressure switch on the faulted side, this current is now applied to the coil of the control relay of the faulted side of the system. The faulted side of the system annunciator is still illuminated since it is receiving its power from the unfaulted side. At this point, should the torquemeter pressure on the faulted side continue to fall, the low pressure switch will actuate at approximately 3.1 psi. Actuation of the low pressure switch completes the ground circuit of the already energized control relay, closing it to energize the feather dump solenoid that immediately dumps the oil from the propeller governor. This same current is applied to the coil of the arm-light-out relay, closing it to remove current from, and extinguish, the annunciator. For testing purposes, the autofeather control switch is equipped with a TEST mode by which the power lever switches in the pedestal can be bypassed to check out the system with the power levers retarded to below the 90% N1 position.

Page 2 Nov 1/09

61-21-00

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 1 Propeller Autofeathering System Schematic - OFF MODE

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Page 3 Nov 1/09

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 2 Propeller Autofeathering System Schematic - ARMED MODE

Page 4 Nov 1/09

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 3 Propeller Autofeathering System Schematic - ENGINE-OUT MODE

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Page 5 Nov 1/09

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

PROPELLER PROPELLER AUTOFEATHERING TROUBLESHOOTING

100100

1. PROCEDURES The flight crew will probably be first to note a fault in operation of the autofeather system; the system must be tested before each flight. Flight crew observations may include comments such as “no annunciation during test” or “propeller not feathering during test”. Regardless of the nature of the particular complaint, the same troubleshooting Chart (Chart 1, in Figure 101) will enable the technician to properly differentiate between faulted components of the autofeather system. As a result of the interplay between RH and LH sides of the system, a certain logical continuity must be maintained during the troubleshooting procedure in order to eliminate confusion and needless replacement of components; therefore, Hawker Beechcraft Corporation strongly advises the technician to follow the troubleshooting Chart until the Chart suggests some terminal activity which should lead to fault correction.

A. Operational Check The following check procedure should be performed prior to beginning the troubleshooting process; this should more clearly define the exact nature of the complaint. This procedure is the only procedure recommended for evaluating the switching level of the torque pressure switches and evaluating adjustment of the autofeather power lever switches. NOTE: The first half of this procedure evaluates the torque pressure switches. (1) Start both engines. (2) Place the power levers in IDLE. (3) Hold the autofeather control switch in the TEST position. (4) The propellers should remain unfeathered, and the autofeather annunciator lights should remain extinguished. (5) Increase engine torque slightly above 650 foot-pounds and hold the autofeather control switch in TEST. Both autofeather annunciators should illuminate, confirming the switching of the high pressure torque switches to their normally open position. (6) Slowly retard the LH engine power lever while holding the control switch in the TEST position, and check that the RH AUTOFEATHER light extinguishes between 450 and 650 foot-pounds, thus signaling the switching of the high pressure switch to its NC position. Continue retarding the LH engine power lever and check that both the LH and RH autofeather annunciator lights extinguish and the propeller begins to feather between 220 and 420 foot-pounds: this indicates that the low pressure switch has closed. NOTE: Any indication that the pressure switches have not reacted in the predicted manner will necessitate replacing the affected pressure switch. As the propeller blades rotate toward feather, the engine torque will increase above the low pressure switch setting and result in system cycling. (7) Repeat the preceding check with the RH engine. NOTE: The balance of this check evaluates power lever switch adjustment.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (8) Place the autofeather control switch in the ARM position. (9) Advance the LH power lever 89 to 91% N1; the propeller should remain unfeathered and the annunciator lights should remain extinguished. (10) Retard the LH power lever and perform the previous check with the RH power lever; the same results should be observed. (11) Advance both power levers 89 to 91% N1; both autofeather annunciators should illuminate. NOTE: Any indications inconsistent with the expected results of this check will necessitate replacing or readjusting the autofeather power lever switches.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

NO

28 vdc during test at pin: 13 of pcb A119(left)(1900/C) or 31 of pcb A121(right)(1900); 31 of pcb A120(right)(1900C) NO

Fault in high pressure switch on opposite side

Did annunciator illuminate during test

YES

NO YES

Did annunciator extinguish during test

YES

Fault in pcb

NO

Propeller will not feather or feathers randomly with prop sync on

Ground signal during test at pin: 25 of pcb A119(left)(1900/C) or 43 of pcb A121(right)(1900); 43 of pcb A120(right)(1900C)

YES

NO

Fault in low pressure torque switch

YES

28 vdc during test at pin: 12 of pcb A119(left)(1900/C) or 30 of pcb A121(right)(1900); 30 of pcb A120(right)(1900C) NO

Fault in feather dump solenoid

YES

Engine harness connectors P13 and P18 switched

Fault in pcb

Fault in high pressure torque switch

* Perform the PROPELLER AUTOFEATHERING OPERATIONAL CHECK in this chapter before beginning this troubleshooting procedure.

UB61B 017156AA.AI

Figure 101 Troubleshooting - Propeller Autofeathering

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

B. Auto-Feather System Troubleshooting, Engines Off NOTE: The following Steps can be used to test the Auto-Feather System electrically without running the engines. (1) Place the power levers in the full forward position and the auto-feather switch in ARM or secure the auto-feather test switch in the TEST position. (2) Disconnect all four auto-feather pressure switches. (3) Install a jumper wire between pins E and F on both high pressure switches. (4) Perform the APPLYING GROUND POWER procedure (Ref. Chapter 24-40-00). (5) Both auto-feather annunciators should be illuminated. (6) On the RH high pressure switch, move the jumper from pin F to pin D. (So that the jumper is from pin E to pin D). (7) The LH auto-feather annunciator should extinguish. (8) On the RH low pressure switch, place a jumper wire between pins A and B. (9) The RH auto-feather annunciator should extinguish and the RH dump valve should activate. (10) Remove the jumper from the RH low pressure switch and return the jumper to pin F on the RH high pressure switch. (So that the jumper is from pin E to pin F again). (11) Both auto-feather annunciators should be illuminated. (12) On the LH high pressure switch, move the jumper wire from pin F to pin D. (So that the jumper is from pins E to D). (13) The RH auto-feather annunciator should extinguish. (14) On the LH low pressure switch, place a jumper wire between pins A to B. (15) The LH auto-feather annunciator should extinguish and the LH dump valve should activate. (16) Remove the jumper from the LH low pressure switch and return the jumper to pin F on the LH high pressure switch. (So that the jumper is from pin E to pin F again). (17) Both auto-feather annunciators should be illuminated. (18) Remove the jumper wires from both the RH and LH low pressure switches. (19) Remove the jumper wires from both the RH and LH high pressure switches. (20) Perform the REMOVING GROUND POWER procedure (Ref. Chapter 24-40-00). (21) Connect all four auto-feather pressure switches. (22) Place the power levers in the idle detent position and place the auto-feather switch to the OFF position.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (23) If the auto-feather test switch has been secured then unsecure the auto-feather test switch and place the switch to the OFF position. (24) Return the airplane to service.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

PROPELLER PROPELLER AUTOFEATHERING MAINTENANCE PRACTICES

200200

1. DUMP VALVE A. Removal (1) With the battery and generator switches OFF and the external power source disconnected, remove the lower forward cowling (Ref. Chapter 71-10-00). NOTE: The dump valve is the lower of the two solenoid units mounted on the overspeed governor. (2) Remove safety wire and disconnect the electrical connector of the dump valve. (3) Remove safety wire and remove the screws securing the dump valve in place.

B. Installation (1) Secure the dump valve solenoid to the overspeed governor with the attaching screws, then safety the screws. (2) Determine the proper electrical connectors from the applicable wiring diagrams. Connect the electrical connector to the solenoid and safety it. WARNING: Ensure that the proper electrical connectors are attached to the propeller governor, overspeed governor, and the autofeather low pressure switch by checking the applicable wiring diagrams. The connectors to these plugs are identical and they could be swapped.

2. PRESSURE SWITCH A. Removal NOTE: Removal and Installation for the high pressure and low pressure switches are identical. The high pressure switch is mounted adjacent to the torque pressure transmitter just forward and above the left exhaust outlet on each engine. The low pressure switch is mounted just below and to the left of the high pressure switch. (1) Perform the REMOVING GROUND POWER procedure (Ref. Chapter 24-40-00). (2) Perform BATTERY DISCONNECTION procedure (Ref. Chapter 24-31-00). (3) Set generator switch to OFF. (4) Remove the upper forward cowling (Ref. Chapter 71-10-00). (5) Disconnect the electrical connector from the switch. (6) Remove safety wire and loosen mounting nut at the base of the switch. (7) Remove the switch by unscrewing switch from the torque manifold.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (8) Install a protective plug in the switch mounting hole to prevent contamination of the torquemeter chamber oil supply.

B. Installation (1) Remove the protective plug from the switch mounting hole in the torque manifold. (2) Screw the switch into the torque manifold and tighten the mounting nut. (Torque the mounting nut to 30 inch-pounds maximum.) Lock the nut in place with safety wire. (3) Determine the proper electrical connectors from the applicable wiring diagrams. Connect the electrical connector to the switch. WARNING: Ensure that the proper electrical connectors are attached to the propeller governor, overspeed governor, and the autofeather low pressure switch by checking the applicable wiring diagrams. The connectors to these plugs are identical and they could be swapped. (4) Install the upper forward cowling (Ref. Chapter 71-10-00). (5) Perform BATTERY CONNECTION procedure (Ref. Chapter 24-31-00). (6) Perform AUTOFEATHER OPERATIONAL CHECK (Ref. 61-21-00).

C. Adjustment Access to the autofeather power switches is through the cover on the left side of the pedestal. The switches may be adjusted as follows: NOTE: When the outside temperature is near - 30°C / - 22°F and colder, the engines will reach maximum torque before attaining 90% N1. For cold weather operations, to ensure positive system activation, the reference line should be set at approximately 86% N1. (1) Run the engines 85 to 90% N1 and mark a reference line on the pedestal with masking tape. It is recommended that the reference line be as close to 90% N1 as possible. (2) Shut down the engines. (3) Disconnect the P16 connectors at the right and left high pressure switches and jumper pins E to F. (4) Apply external and/or ship electrical power to the aircraft. (5) With the autofeather control switch in the ARM position, adjust each power switch in its mounting slots until its respective autofeather light in the pedestal annunciator panel illuminates when power levers are moved to the masking tape reference mark from cutoff.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

PROPELLER PROPELLER SYNCHROPHASER DESCRIPTION AND OPERATION

61-22-00 00

1. GENERAL The propeller synchrophaser automatically matches the RPM of both propellers as a result of maintaining a specific phase relationship between the blades of the left and right propellers. The control box senses pulses which are generated by pickups mounted at identical locations on both engines, in relation to the engine centerline. Ferrous metal targets, mounted on the propeller spinner bulkheads, provide the pulse reference for the pickups. The magnetic pickups operate on the magnetoelectric induction principle: when the ferrous metal target passes in close proximity to the magnetic field of the pickup, an alternating current is induced in the pickup. The waveform of the current is essentially a sine wave with a very narrow bandwidth at the peaks of the wave. The amplitude of the positive signal may not necessarily be equal to the amplitude of the negative signal; therefore, it will be necessary to measure peak-to-peak voltages (the magnitude of difference between the voltage of the positive peak and the voltage of the negative peak) in order to properly assess the operating outputs of the pickups. The duration of the waveform, from the time it leaves the baseline until it returns to the baseline after the negative peak, is approximately 1/215 of the time the waveform remains at the baseline. The frequency of the alternating current is dependent upon propeller RPM and will equate to 25 Hz. at 1500 RPM. Attesting to the sensitivity of the control box is the fact that the entire waveform cycle occurs during a time span of less than 1/5000 second. As a result of the peculiar characteristics of the AC waveform produced by the pickups, an oscilloscope must be used to measure peak-to-peak voltages. The magnitude of the peak-to-peak voltage should be between 4 and 110 volts. The control box does not respond to amplitude of the input signals, but to phasing of the waveforms. A specific amount of voltage change, approximately 0.7 volts, is necessary to trigger the control box. Maintaining a time-phase relationship is the control box's only priority. The control box senses the pulses from each pickup and attempts to superimpose the waveforms by trimming the speed of each propeller. Speed trim of the propellers is accomplished by the control box with correction commands to each propeller governor. The character of these correction commands is in the form of pulse width modulated 28 volt direct current: the duration of the pulse of current is regulated to produce the proper amount of speed trim. The amount of speed trim per pulse width modulation by the control box is a function of the governor and will always be within a very narrow range (holding range of approximately 25 ± 2 RPM). The governor servo can increase, but never decrease, the speed set by the propeller control lever. The RPM of one propeller will follow the changes in RPM of the other propeller over the predetermined holding range of the governor. This limited holding range prevents either propeller from losing more than a limited RPM if the RPM of the other propeller is manually reduced, such as in power changes or propeller feathering, while the synchrophaser is on. The synchrophaser system is controlled through a toggle switch placarded PROP SYNCH-ON-OFF. To operate the system, synchronize the propellers in the normal manner and turn the synchrophaser on. To change RPM, adjust both propellers at the same time. This will keep the setting within the holding range of the system. If the synchrophaser is on, but will not synchronize propellers, the propeller speeds are not within the limits required for the system to assume control (outside capture range). Turn the synchrophaser off, synchronize the propellers manually, then turn the synchrophaser on.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

PROPELLER PROPELLER SYNCHROPHASER TROUBLESHOOTING

100100

1. PROCEDURES The first indication that a fault may exist in the synchrophaser system will most likely be squawks from the flight crew; however, performing a functional check of the operation of the synchrophaser system before troubleshooting may yield information essential to differentiating faults within the system. The functional check which follows should be performed as the first phase of troubleshooting. The second phase of troubleshooting will include checking the wiring and peripheral components. A breakout box can be fabricated in the shop and will provide convenient access to the various circuits of the synchrophaser system (Ref. Figure 102). Should the checks fail to reveal the nature of the fault, it will be necessary to begin the third phase of trouble-shooting, which will be to evaluate the reaction of the propeller governors. This is done by using the breakout box to check the effect of the synchrophaser control box on the speed-biasing coil of the propeller governor. If no faults became apparent during previous troubleshooting attempts, the fourth phase of troubleshooting is to evaluate the character of the outputs from the magnetic pickups. In the event none of the previous checks or procedures reveals the exact nature of the fault, the fault may be assumed to be in the control box, and a control box, known to be operational, should be installed. The troubleshooting diagram, refer to the Chart in Figure 104, proposes a logical sequence of checks and activities to be performed during the troubleshooting process and may aid the technician in eliminating confusion and maintaining logical order. All manners of faults have been investigated, but only the most probable causes of faults in the system are detailed in this diagram.

2. SYNCHROPHASER A. Functional Check The proper operation of the synchrophaser system may be ground checked as follows: CAUTION: Do not overtorque the engines. (1) Start the engines and increase power until N1 torque reads approximately 1500 foot-pounds. With the propeller levers in HIGH RPM, the propellers will be at about 1700 RPM. (2) Using the propeller levers, reduce propeller RPM on both engines to as nearly 1500 RPM as possible. (3) Turn the synchrophaser switch to the ON position. (4) Using the propeller lever, increase the RPM of the left propeller slightly and check that the synchrophaser increased the RPM on the right propeller to match. NOTE: Should the previous check yield negative results, turn the synchrophaser off and repeat the above procedure. (5) If the previous check indicated that the synchrophaser is operating, continue increasing the propeller RPM with the propeller lever until synchronization is lost and record this RPM.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (6) Using the propeller lever, slowly decrease the propeller RPM on the left engine below the RPM of the right engine until synchronization is again lost and record this RPM. (7) The difference between the two propeller speeds at which synchronization was lost is the holding range and should be at least 23 to 27 RPM. This check evaluates the ability of the right propeller governor to respond to correction commands from the control box, as well as the control box's ability to issue commands properly. NOTE: Failure of this check will necessitate performing the synchrophaser continuity check. (8) Repeat the functional check procedure by varying the propeller RPM of the right engine while letting the left propeller governor respond. (9) During this functional check, determine that the communications and navigation equipment is not adversely affected by the synchrophaser operation and that synchrophaser operation is not adversely affected by the operation of other electrical equipment.

B. Synchrophaser Checks Use this check to locate the source of the fault should the synchrophaser system fail the functional check. Use a multimeter with an ohm function or an ohmmeter to accomplish these checks. (1) Check that the battery master switch is off, and open the synchrophaser circuit breaker. (2) Remove the connector from the control box. (3) Check the synchrophaser system wiring for conformity to the schematic in Figure 101, Sheet 1 of 2 or Sheet 2 of 2. (4) Perform the continuity checks as indicated in Table 101 for UC-1 thru UC-40 or Table 102 for UC-41 and After. NOTE: Resistance readings noted in Table 101 or Table 102 may be as much as 20% higher during heat soak following engine shutdown. (5) After performing the continuity checks, close the synchrophaser circuit breaker, turn the battery master switch on and check for battery voltage between pins A and B. (6) Correct any wiring faults that may have been revealed during these continuity checks. Replace any component which failed to meet the proper resistance specified for that component. Restore the system to operational status and perform the functional check again. CAUTION: DO NOT connect the plug to the control box before turning the battery master switch off and opening the synchrophaser circuit breaker. (7) If no faults were found by the continuity checks, perform the propeller GOVERNOR SPEED BIASING COIL CHECK.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Table 101 Synchrophaser Continuity Checks (UC-1 thru UC-40) Test Between Receptacle Pins

Ohmmeter Indication*

Component

A and airplane ground

0 to 2 ohms

A and B

Infinity with switch off; will see some value with switch on

C and G

52 to 68 ohms

Right Pickup

D and E

52 to 68 ohms

Left Pickup

J and R

115 to 205 ohms

Right Governor

K and L

115 to 205 ohms

Left Governor

M and S

0 to 2 ohms

C, G, D, E, K, R, J, or L to ground

Infinity

* Resistance may be as much as 20% higher during heat soak after engine shutdown

P151 GROUND

A

R PRIMARY PROP GOV SIGNAL RET

R

BLUE

A

R PRIMARY PROP GOV VOLTAGE +

J

WHITE

B

R PICKUP -

C

BLUE

BLK

R PICKUP +

G

WHITE

WHT

L PRIMARY PROP GOV VOLTAGE +

L

WHITE

B

L PRIMARY PROP GOV SIGNAL RET

K

BLUE

A

L PICKUP +

E

WHITE

WHT

L PICKUP -

D

BLUE

BLK

28 VDC IN

B

R GOV

R PICKUP

L GOV

L PICKUP

5A SYNCHROPHASER CONTROL BOX

CONTROL SWITCH

UC61B 091902AA.AI

Figure 101 (Sheet 1 of 2) Control Box Wiring Schematic (UC-1 thru UC-40)

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Table 102 Synchrophaser Continuity Checks (UC-41 and After) Test Between Receptacle Pins

Ohmmeter Indication*

Component

A and airplane ground

0 to 2 ohms

A and B

Infinity with switch off; will see some value with switch on

C and G

52 to 68 ohms

Right Pickup

E and D

52 to 68 ohms

Left Pickup

K and R

115 to 205 ohms

Right Governor

J and L

115 to 205 ohms

Left Governor

M and S

0 to 2 ohms

C, G, D, E, K, R, J, or L to ground

Infinity

* Resistance may be as much as 20% higher during heat soak after engine shutdown

P151 GROUND

A

R PRIMARY PROP GOV VOLTAGE +

K

BLUE

A

R PRIMARY PROP GOV SIGNAL RET

R

WHITE

B

R PICKUP -

C

BLUE

BLK

R PICKUP +

G

WHITE

WHT

L PRIMARY PROP GOV VOLTAGE +

J

WHITE

B

L PRIMARY PROP GOV SIGNAL RET

L

BLUE

A

L PICKUP +

E

WHITE

WHT

L PICKUP -

D

BLUE

BLK

28 VDC IN

B

R GOV

R PICKUP

L GOV

L PICKUP

5A SYNCHROPHASER CONTROL BOX

CONTROL SWITCH

UC61B 092145AA.AI

Figure 101 (Sheet 2 of 2) Control Box Wiring Schematic (UC-41 and After)

C. Governor Speed Biasing Coil Check The following checks assess the operation of the governors. Perform this check before evaluating the outputs of the magnetic pickups: (1) Connect the breakout box between the airplane harness and the control box.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (2) Start the engines and increase the propeller RPM to 1600 RPM with the propeller levers in high RPM. (3) Using the propeller levers, reduce the RPM to 1500 RPM to insure that the propellers are being controlled by the governors. (4) Install a jumper between jacks R and A. The speed of the right propeller should noticeably increase. (5) Install the jumper between jacks L and A of the breakout box. The speed of the left propeller should noticeably increase. (6) A failure of either propeller governor to increase propeller speed indicates that there is a fault in the speed biasing coil of that governor. (7) If no faults were revealed by this operational check, perform the MAGNETIC PICKUP VOLTAGE CHECK before condemning the control box.

D. Magnetic Pickup Voltage Check It is important that the output voltages of the magnetic pickups be within the 4 to 110 volt peak-to-peak (Vpp) operating range. In addition, it is also important that the output voltage waveforms generated begin with a 0 to +V transition and the Vpp is the same for each magnetic pickup to within ± 0.10 Vpp (Ref. Figure 103). This procedure requires a ground runup of the engines with an oscilloscope in the cockpit to measure the voltage waveforms. (1) Ensure that all electrical power is off and the PROP SYNCH circuit breaker is disengaged. (2) Disconnect the wire harness plug from the synchrophaser control box and install the breakout box between the plug and the control box. (3) Set the oscilloscope to read 1 volt/division and scan once every 125 milliseconds. (4) Connect the oscilloscope leads to the breakout box to observe the left magnetic pickup. NOTE: Match the polarity of the oscilloscope leads to the polarity of the magnetic pickup. (5) Restore electrical power to the airplane. (6) Start the engines in accordance with the operator's manual. (7) Manually set both propellers as close to 1500 rpm as possible. NOTE: It may be necessary to adjust the sweep of the oscilloscope to obtain a steady picture. (8) Observe the waveforms on the oscilloscope. Record the peak-to-peak voltage (Vpp). The ideal Vpp should be between 4 and 110 volts. The magnetic pickup waveforms should depart smoothly from the baseline, rise sharply to a positive peak (+V), descend sharply to a negative peak (-V), and return smoothly to the baseline (Ref. Figure 103). The waveforms should be symmetrical; if they are not: (a) A ragged waveform indicates a faulty pickup.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (b) An asymmetrical waveform indicates the target is not parallel to the tip of the magnetic pickup. (c) A secondary peak may indicate a damaged target. (9) Repeat the magnetic pickup voltage check for the right magnetic pickup. (10) Shut down the engines in accordance with the operator's manual.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 102 Synchrophaser Breakout Box Schematic

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 103 Magnetic Pickup Voltage Waveform

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

NO

Is circuit breaker closed

Engines synchronize after a few seconds (Note 1)

YES

System operating normally

YES

NO NO

Close circuit breaker

Are engines within capture range

YES

Check circuit integrity (Note 2)

Turn system off, manually resync engines and turn system off

NO

Is resistance specified in Chart 1 between pins J & R or pins K & L low

NO

YES

YES

Correct wiring fault or replace faulty pickup or governor

NO

NO

Are engine harness connectors P13 and P18 switched YES

Reconnect P13 to the prop governor and P18 to the autofeather dump valve

Are synchrophaser checks correct

Replace the faulty governor

NO

Do both governors react to control box (Note 3) YES

Check pickup outputs (Note 4)

Are peak-to-peak voltages within tolerance YES

NO

Can voltages be adjusted by varying pickup gap (Note 5)

NO

YES

Are AC waveforms identical in appearance YES

Is face of target parallel to face of pickup

If system still inoperative replace control box

YES

Adjust or replace targets

Replace control box

Replace pickup

NO

Correct target alignment

UB61B 017157AA.AI

Figure 104 Troubleshooting Flow Chart - Propeller Synchrophaser Note 1

Perform SYNCHROPHASER FUNCTIONAL CHECK.

Note 2

Perform SYNCHROPHASER CHECK.

Note 3

Perform GOVERNOR SPEED BIASING COIL CHECK.

Note 4

Perform MAGNETIC PICKUP VOLTAGE CHECK.

Note 5

Refer to SYNCHROPHASER PICKUP INSTALLATION AND ADJUSTMENT procedure.

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PROPELLER PROPELLER SYNCHROPHASER MAINTENANCE PRACTICES

200200

1. PICKUP A. Installation and Adjustment The synchrophaser pickups are located in a bracket adjacent to each propeller deicer brush block. The pickup should be installed and adjusted as follows: (1) Ensure that battery and generator switches are off and the external power source is disconnected. (2) Remove the lower forward engine cowling in accordance with COWLING REMOVALS, Chapter 71-10-00. (3) Install the pickup into the mounting bracket and secure with the mounting locknuts. NOTE: For proper function of the propeller synchrophaser, the pickup must be installed as follows: 5° behind the blade for the left hand propeller; 20° ahead of the blade for the right hand propeller. (4) Adjust the magnetic pickup to obtain a reference gap of 0.050-inch between the magnetic pickup tip and the target. (5) Tighten the locknuts at 25 inch-pounds. (6) Connect the two electrical leads. (7) Perform the MAGNETIC PICKUP VOLTAGE CHECKS procedure in the TROUBLESHOOTING section. Record the peak-to-peak voltage (Vpp) values for each pickup. Optimum voltage output is between 4 and 110 volts. (8) Readjust the pickups as required to obtain matching voltage outputs from each pickup. Move the pickup away from the target to decrease the voltage output and closer to the target to increase the voltage output. Altering gap by 0.007-inch will increase/decrease voltage by approximately 0.5 volt. (9) Repeat Steps (7) and (8) as required to match the voltage outputs from each pickup. Install the cowling in accordance with COWLING INSTALLATION in Chapter 71-10-00. (10) Torque the locknut on the pickup to a maximum of 25 inch-pounds and safety wire the locknut. (11) If the connector plug was previously disconnected, connect the plug to the control box. (12) Perform the SYNCHROPHASER FUNCTIONAL CHECK procedure in the TROUBLESHOOTING section.

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1. THROTTLE SWITCH A. Adjustment Power to the low pitch stop system is provided through the throttle switch actuated by the power levers. (1) Remove the panel from the left side of the pedestal. (2) Adjust the throttle switch in its mounting slots until it is in the full overtravel position (approximately 0.06 inch beyond the point of actuation) with the power levers in the forward position. Lift either power lever and check that the switch clicks at a point approximately 0.20 inch before the end travel by the lever attached to the actuating arm.

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POWER PLANT 71-00-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Engine . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 203 Buildup . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 213 Fuel Control Unit . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 213 Purge and Soak Procedure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 213 Handling Fuel Nozzles . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 214 Engine Washing Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 214 Engine Bonding Strap Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 215

COWLING 71-10-00 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Cowling . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Forward Cowling Latch . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Adjustment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 201 201 205 205

MOUNTS 71-20-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Engine Truss Assembly . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Engine Mounts . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Truss Assembly . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Engine Truss Bolt Torque Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202 Inspection Criteria . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 204 Isolators Only . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 204

FIRESEALS 71-30-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1

ELECTRICAL HARNESS 71-50-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1

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ENGINE DRAINS 71-70-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Engine Fuel Purge System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Fuel Purge Tank . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Cleaning . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Fuel Purge System Air Filter . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Cleaning . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Fuel Purge System Check Valve . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202 Inspection, Cleaning and Leakage Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202 Fuel Flow Divider/Purge Valve . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 205 Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 205

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1. GENERAL The PT6A-65B engines installed on the airplane are of the “free turbine” type. Each engine utilizes two independent turbine sections: one driving the compressor in the gas generator section and the second (two-stage power turbine) driving the propeller shaft through a reduction gearbox (Ref. Figure 201, in the Maintenance Practices section). Each engine is self-sufficient since its gas generator driven oil system provides lubrication for all areas of the engine, pressure for the torquemeter and power for propeller pitch control. Inlet air enters the engine through an annular plenum chamber, formed by the compressor inlet case, where it is directed forward to the compressor. The PT6A-65B engine compressor consists of four axial stages combined with a single centrifugal stage, assembled as an integral unit. The engine is equipped with a compressor wash ring at the compressor air inlet case. A row of stator vanes, located between each stage of compression, diffuses the air, raises its static pressure and directs it to the next stage of compression. The compressed air passes through diffuser tubes which turn the air through 90° in direction and convert velocity to static pressure. The diffused air then passes through straightening vanes to the annulus surrounding the combustion chamber liner assembly. The combustion chamber liner consists of two annular wrappers bolted together at the front dome-shaped end. The outer wrapper incorporates an integral large exit duct. The liner assembly has perforations of various sizes that allow entry of compressor delivery air. The flow of air changes direction 180° as it enters and mixes with fuel. The fuel/air mixture is ignited and the resultant expanding gases are directed to the turbines. Fuel is supplied by a dual manifold consisting of primary and secondary transfer tubes and adapters. The fuel is then injected into the combustion chamber liner through 14 individual nozzles arranged in two sets of seven. The fuel/air mixture is ignited by two spark ignitors which protrude into the liner. The resultant gases expand from the liner, reverse direction in the exit duct zone and pass through the compressor turbine inlet guide vanes to the single-stage compressor turbine. The guide vanes ensure that the expanding gases contact the turbine blades at the correct angle, with minimum loss of energy. The still expanding gases are then directed forward to drive the power turbine section. The two-stage power turbine, consisting of the first-stage guide vane and turbine and the second-stage inlet guide vane and turbine, drives the propeller shaft through a reduction gearbox. The compressor and power turbines are located in the approximate center of the engine with their respective shafts extending in opposite directions. The exhaust gas from the power turbine is collected, routed into the exhaust duct assembly and directed to the atmosphere by twin opposed exhaust stubs. Interturbine temperature (T5) is monitored by an integral bus-bar, probe and harness assembly installed between the compressor and power turbines with the probes projecting into the gas path. A terminal block mounted on the gas generator case provides a connection point to the flight compartment instrumentation. All engine-driven accessories, with the exception of the propeller governor, overspeed governor and tachometer-generator, are mounted on the accessory gearbox at the rear of the engine. These components are driven by the compressor by means of a coupling shaft which extends the drive through a tube at the center of the oil tank. The engine oil supply is contained in an integral oil tank which forms the rear section of the compressor inlet case. The tank has a total capacity of 2.35 U.S. gallons and is provided with a dipstick. Refer to the engine Maintenance Manual for complete description and maintenance information on the engine and accessories.

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1. ENGINE A. Removal NOTE: Airplanes idle for more than seven days due to maintenance or other reasons should either have the engines operated or preserved. Refer to the appropriate procedures in Chapter 10-10-00 or the Pratt and Whitney PT6A-65B Engine Maintenance Manual. Engine removal is accomplished as described in the following procedure: (1) Preliminary Steps. (a) Position a suitable stand under the tail of the airplane (Ref. Chapter 91-00-00 Figure 1 Sheet 2 of 10). CAUTION: With both engines removed, the airplane becomes tail heavy. Failure to install a tail stand may result in the airplane tail settling to the ramp. (b) Drain the engine oil. Refer to CHANGING THE ENGINE OIL procedure (Ref. Chapter 79-00-00). (c) Make sure all electrical power to the airplane is disconnected. (2) Perform the PROPELLER REMOVAL procedure (Ref. Chapter 61-10-00). (3) Perform the COWLING REMOVAL procedure (Ref. 71-10-00). (4) Place a suitable cover over the engine inlet air screen to prevent foreign matter from entering the air inlet. (5) Disconnect the following plumbing hoses from the engine: NOTE: Tag or identify all hoses to facilitate and ensure correct installation on the engine. Cap all open hoses and engine ports to prevent contamination. (a) Oil inlet hose - flare fitting. (b) Oil overboard breather hose - clamp. (c) Oil outlet hose - flare fitting. (d) Fuel inlet hose - flare fitting. (e) Fuel return hose - flare fitting. (f) Forward combustion chamber drain hose - flare fitting. (g) Manifold dump valve drain hose - flare fitting. (h) Fuel pump drain hoses - clamps.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (i) Air bleed hose - flare fittings. (j) Fuel Purge System lines. (6) Except when the existing engine is to be replaced with another, all electrical leads may be disconnected by removing the electrical plugs from the firewall receptacles. If the engine harness is to be used on a replacement engine, disconnect the electrical leads and plugs from the components listed below and remove the electrical harness clamps necessary to allow engine removal. NOTE: Tag or identify all electrical connectors and note harness clamp locations to facilitate correct installation. Cap all plugs and receptacles to prevent contamination. (a) Starter-generator. (b) Ignitor control box. (c) Interstage turbine temperature thermocouple (Ref. Figure 204). (d) Gas generator tachometer generator. (e) Oil pressure transmitter. (f) Oil temperature transmitter. (g) Propeller tachometer generator (Ref. Figure 206). (h) Propeller overspeed governor (Ref. Figure 205). (i) Propeller deicer brushes. (j) Torque pressure transmitter. (k) High pressure autofeather switch. (l) Low pressure autofeather switch. (m) Autofeather dump solenoid. (7) Disconnect the engine controls as follows: (a) Disconnect the propeller control cable from the governor control arm. Remove the fire sleeve from the control cable (Ref. Figure 203). (b) Disconnect the power control cable from the power control lever in the mounting bracket for the cam and attendant control linkage. (c) Disconnect the condition lever control cable from the linkage at the fuel control unit, or start control unit. NOTE: Tag and retain the throttle attaching bolt, washer, and nut for installation on the engine.

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(8) Remove the engine from the airplane as follows: NOTE: Identify and retain all engine mounting bolts, washers, and nuts to facilitate and ensure a correct new engine installation. (a) Attach an engine sling to the engine hoisting lugs (Ref. Chapter 91-00-00). Position a suitable hoist directly over the eye of the hoisting sling. (b) Remove the cotter pin, nut, washer and bolt (four places) which secure the truss mount to the engine mounts. Refer to the Chapter 71-20-00 ENGINE MOUNT illustration in the Maintenance Practices section. (c) The engine may now be hoisted to remove from the airplane.

B. Installation (1) Attach an engine sling to the engine hoisting lugs (Ref. Chapter 91-00-00). (2) Hoist the engine over the airplane engine mount and slowly lower it to align the truss mount with the engine mount. (3) With the truss mount and engine mount properly aligned, install the four bolts, washers, and nuts which secure the engine mounts to the truss and torque the bolts. Refer to ENGINE INSTALLATION, BUILD-UP AND TRUSS MOUNT TORQUES Table in Chapter 71-20-00. Install the cotter pins. (4) Install the bolts and screws at the two cowl bulkheads. (5) Connect the following hoses to the engine fittings. NOTE: Apply a light coat of lubricant or compound to the male threads of all fittings (Ref. THREAD LUBRICANTS, Chapter 91-00-00). (a) Oil inlet hose - flare fitting. (b) Oil overboard breather hose - clamp. (c) Oil outlet hose - flare fitting. (d) Fuel inlet hose - flare fitting. (e) Fuel return hose - flare fitting. (f) Forward combustion chamber drain hose - flare fitting. (g) Manifold dump valve drain hose - flare fitting. (h) Fuel Purge System lines. (i) Fuel control drain hoses - clamps. (j) Air bleed hose - flare fitting.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (6) Install the engine harness plugs into the firewall receptacles. Route the electrical harness and secure it with the attaching clamps as it was installed on the replaced engine. Make all electrical connections finger tight to the engine components listed below and safety wire the electrical plug. WARNING: Ensure that the proper electrical connectors are attached, P17 to the propeller overspeed governor and P18 to the autofeather dump solenoid. The connectors are of an identical type and could be swapped. (a) Starter-generator. (b) Ignitor control box. CAUTION: Torque the thermocouple harness connections according to the Pratt and Whitney PT6A-65B Engine Maintenance Manual. (c) Interstage turbine temperature thermocouple (Ref. Figure 204). (d) Gas generator tachometer generator. (e) Oil pressure transmitter. (f) Oil temperature transmitter. (g) Propeller tachometer generator (Ref. Figure 206). (h) Propeller overspeed governor (Ref. Figure 205). (i) Propeller deicer brushes. (j) Torque pressure transmitter. (k) High pressure autofeather switch. (l) Low pressure autofeather switch. (m) Autofeather dump solenoid. (7) Control cables installation: (a) Connect the power control cable to the power lever at the mounting bracket for the cam and attendant linkage. (b) Connect the condition lever control cable to the linkage at the fuel control unit, or start control unit. (c) Connect the propeller governor control cable. Rig the engine and propeller reversing mechanism (Ref. Chapter 61-20-00 and Chapter 76-00-00). (8) Perform the PROPELLER INSTALLATION procedure (Ref. Chapter 61-10-00). (9) Remove the cover over the engine inlet air screen. (10) Perform the COWLING INSTALLATION procedure (Ref. 71-10-00).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (11) If the airplane is not to be released for flight operations, refer to the Pratt and Whitney PT6A-65B Engine Maintenance Manual or STORAGE, Chapter 10-10-00 in this Maintenance Manual for the proper procedures.

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Figure 201 PT6A-65B Engine

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Figure 202 Engine Mount

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Figure 203 Engine and Propeller Controls

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Figure 204 Trim Thermocouple and T5 Thermocouple Terminal Block

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Figure 205 Propeller Governor Adjustment Points

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C. Buildup CAUTION: Consult the Pratt and Whitney PT6A-65B Engine Maintenance Manual before attempting to remove the new engine from the shipping container. Engine buildup consists of the installation of accessories and equipment on the new engine. Refer to the Pratt and Whitney PT6A-65B Engine Maintenance Manual for removal and replacement procedures. After all accessories have been installed, complete the engine buildup as follows: NOTE: Tag or identify all hoses, bolts, nuts, washers, electrical connectors, and note harness clamp locations for installation on the new engine. Cap all open hoses and engine ports to prevent contamination. (1) Remove the engine mount (vibration isolators) from the old engine and install on the new engine. NOTE: All engine mounts (vibration isolators) on an engine must be of the same manufacturer and type. The engine mount (vibration isolator) may be rebuilt by the replacement of the molded rubber assembly. (2) Torque the engine mount (vibration isolator) attaching bolts to engine case 225 to 300 inch pounds (Ref. Table 202, ENGINE INSTALLATION, AND BUILDUP AND TRUSS MOUNT TORQUES, 71-20-00). (3) Remove the bleed air adapter from the old engine and install on the new engine using the same bolt and washer combination, with a new gasket. (4) Remove engine control brackets and supports from the old engine and install on the new engine. (5) Remove the exhaust stacks and install on corresponding (left or right) exhaust ports.

2. FUEL CONTROL UNIT A. Purge and Soak Procedure CAUTION: Throughout the following procedure, observe the starter operating limits of 30 seconds ON, 10 minutes OFF, 30 seconds ON, 30 minutes OFF, 30 seconds ON, 30 minutes OFF, 30 seconds ON, 30 minutes OFF, 30 seconds ON, 30 minutes OFF. (1) To purge the fuel control unit of preservation oil, install the unit on the engine and connect it to the airplane fuel supply. Install a suitable line between the fuel control outlet and a clean container. (2) With the ignition OFF, and the fuel control lever in the IDLE position, motor the engine until clean fuel flows from the fuel outlet. During the motoring run, move the fuel lever from the IDLE to CUTOFF position to ascertain fuel shutoff. (3) Remove the line used to purge the fuel control unit and connect the engine plumbing to the fuel control outlet. Check the installation for leakage. NOTE: An eight-hour soaking period is recommended following the purging procedure. This will enable the bypass valve diaphragm to regain the pliability it possessed during unit calibration and will permit the synthetic packing in the unit to swell for a satisfactory seal. CAUTION: Before attempting a start, motor the engine with the fuel shutoff to ensure that all fuel has been purged from the engine.

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3. HANDLING FUEL NOZZLES Extreme care should be exercised when handling fuel nozzles; even fingerprints on the orifice may produce a poor spray pattern. If fuel nozzles are not to be installed in the engine immediately, they should be placed in a covered container to prevent exposure to dirt. If nozzles are to be shipped or stored, they must be separately and securely wrapped. Packing nozzles together in a single container which allows relative movement will damage orifice faces.

4. ENGINE WASHING PROCEDURES CAUTION: The compressor bleed air must be capped off when washing the compressor. For both external and compressor washing procedures for the PT6A-65B engine, refer to the Pratt and Whitney PT6A-65B Engine Maintenance Manual. CAUTION: For continuous cycling, use of the starter is limited to 30 seconds ON, 10 minutes OFF, 30 seconds ON, 30 minutes OFF. For engine clearing and restart, use of the starter is limited to 30 seconds ON, 3 minutes OFF, 30 seconds ON, 30 minutes OFF. Before any external engine washing, cover or protect all electrical components and plugs located on the engine or in the engine compartment. Do not allow water to enter the engine air inlet or exhaust. High pressure water or solvent should not be directed on electrical components or mechanical parts having air vent holes. After washing the engine blow dry the components with dry shop air. NOTE: The following engine wash cycle limitations are applicable to Lucas Aerospace starter-generator installations.

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Table 201 WASH CYCLE LIMITATIONS

Cycle

Starter Mode

Time Duration1 P/N 230852,4

Time Duration1 P/N 230783

Wash

ON

30 Seconds

30 Seconds

Soak

OFF

15 Minutes

15 Minutes

Rinse

ON

20 Seconds

20 Seconds

Dry

OFF

30 Minutes

5 Minutes

Rinse

ON

20 Seconds

20 Seconds

Dry

OFF

30 Minutes

5 Minutes

Clear

ON

30 Seconds

30 Seconds

Starter Cool

OFF

30 Minutes

15 Minutes

Start

ON

See Note 5

See Note 5

Remarks Inject Fluid at 5% N1

Inject Fluid at 5% N1

Inject Fluid at 5% N1

1. Maximum starter ON and minimum starter OFF times shown. 2. Starter-generator P/N 23085 (UC-1 thru UC-142). 3. Starter-generator P/N 23078 (UC-143 and After and Airplanes With Kit 114-9034-1S Installed). 4. Starter current must be limited to 800 amps. 5. Per normal start limitations (Ref. AFM/POH).

5. ENGINE BONDING STRAP INSPECTION Apply an anti-erosion tape or other approved anti-chaffing material to the problem area(s) of the strap only if the broken and/or chaffed strands of the existing strap installation do not exceed 10% (i.e., 134 strands) of the total strand count (i.e., 134 of 1344 = 10%) (Ref. Figure 206).

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Figure 206 Engine Bonding Strap

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

POWER PLANT COWLING MAINTENANCE PRACTICES

71-10-00 200200

1. COWLING A. Removal The engine cowling consists of five separate sections which must be removed in the proper sequence. The various cowling sections are attached with cam-type fasteners, dzeus fasteners, and screws (Ref. Figure 201 and 202). The forward portion of the cowling houses the engine power turbine and propeller gearing reduction. The middle portion houses the induction system and compressor turbine and the aft portion encloses the accessory drive section and provides access for oil servicing. CAUTION: To prevent damage to the propeller spinner, the cowling edges, and the exterior painted surfaces, it is imperative that the cowling be kept level throughout the removal procedure. This may be accomplished by employing two men, one on each side of the cowling. NOTE: Identify and tag all cowling attaching parts. (1) Release the four cam-type fasteners of the upper forward cowling. (2) Raise the cowling and remove from the airplane. (3) Release the dzeus fasteners, remove the attaching screws, raise the upper aft cowling and remove from the airplane. (4) Release the dzeus fasteners, remove the attaching screws, raise the cowling, disconnect the wing ice light wiring and remove the upper middle cowling from the airplane. (5) Release the dzeus fasteners, remove the attaching screws, and carefully lower the lower forward cowling from the airplane. (6) Remove the lower aft cowling as follows: (a) Disconnect the ice vane linkage at the ice vane control arm. CAUTION: Bushings are installed in the holes of the control arms which attach to the ice vane clevises and may have a tendency to fall out when the clevises are disconnected. (b) Disconnect the starter-generator cooling duct from the lower LH side of the cowling. (c) Release the dzeus fasteners, remove the attaching screws and lower the cowling from the airplane.

B. Installation CAUTION: To prevent damage to the propeller spinner, the cowling edges, and the exterior painted surfaces, it is imperative that the cowling be kept level throughout the installation procedure. This may be accomplished by employing two men, one on each side of the cowling. All skin gaps at the mating surfaces must be within 0.00 to 0.06-inch.

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(1) Install the lower aft cowling as follows: (a) Align the screw holes of the cowling with those of the engine mount truss and secure the dzeus fasteners. Install the attaching screws (Ref. Figure 201 and 202). (b) Connect the starter-generator cooling duct to the lower LH side of the cowling. (c) Install the bushings in the holes where the ice vane clevises attach to the control arms and attach the clevises to the control arms. Connect the wing ice light wiring. (2) Carefully raise the lower forward cowling into position, align the screw holes of the cowling with those of the engine mount truss and secure the dzeus fasteners. Install the attaching screws. NOTE: Check inertial separator linkage for proper travel before installing middle cowling. (3) Position the upper middle cowling to the airplane and secure the dzeus fasteners. Install the attaching screws. (4) Align the holes of the upper aft cowling with those of the engine mount truss and secure the dzeus fasteners. Install the attaching screws. (5) Position the upper forward cowling and secure with the four cam-type fasteners. (6) Remove all cowling identification tags.

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Figure 201 Cowling Installation (UA-1 and After, UB-1 and After)

71-10-00

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Figure 202 Cowling Installation (UC-2 and After) Page 204 Nov 1/09

71-10-00

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2. FORWARD COWLING LATCH A. Adjustment The forward lower cowling incorporates a hook that engages a cam pin (Ref. Figure 203, Detail C). The hook is installed in the forward lower cowling and the cam pin is installed in the upper cowling. Check that the hook is installed so it latches over the cam pin when it is turned in the direction of the arrow that is decaled on the side of the cowling. The head of the cam pin is 6-sided (hexagon) and is installed in a 12-sided hole that locks it in place. The cam pin is held in place with a snap ring (Ref. Detail D). To adjust the latch, remove the snap ring and pull the head of the cam pin out of the 12-sided hole. The pin can be rotated in 30° increments to increase or decrease the latching force. Check that the forward upper cowling guide pins are inserted in the anchor holes and the cowling is snug against the forward lower cowling.

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FORWARD UPPER COWLING

B B

A C

C DETAIL

A

FORWARD LOWER COWLING

A

SNAP RING

CAM PIN 12 SIDED HOLE DETAIL

D

CAM PIN HEAD

D INSIDE VIEW DETAIL

B

HOOK

UPPER COWLING CAM PIN

LOWER COWLING

LEFT SIDE DETAIL

C

RIGHT SIDE

Figure 203 Forward Cowling Latch Adjustment

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UC71B 023350AA.AI

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

POWER PLANT MOUNTS DESCRIPTION AND OPERATION

71-20-00 00

1. GENERAL A. Engine Truss Assembly The engine is supported by a welded tubular truss assembly of 4130 chrome molybdenum steel, which is bolted to the nacelle firewall at four attaching points.

B. Engine Mounts Four engine mounts (vibration isolators) are utilized to secure the engine to the truss assembly and dampen engine vibration to the airframe structure. Each engine mount is attached to the engine by four bolts, and each isolator is anchored to the truss assembly with one bolt (Ref. Figure 201, in the Maintenance Practices section). Individual engine mounts may be replaced or overhauled without removing the engine provided the weight of the engine is adequately supported. For detailed instructions on the removal and installation of engine mounts, refer to ENGINE BUILDUP and ENGINE INSTALLATION, 71-00-00. All engine mounts (vibration isolators) on an engine must be of the same manufacturer and type.

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POWER PLANT MOUNTS MAINTENANCE PRACTICES

200200

1. TRUSS ASSEMBLY A. Installation Should the engine truss assemblies be removed for any reason, or a new truss assembly be installed, the following Steps should be followed to ensure proper alignment with the firewall and the engine cowling: (1) Position the truss assembly up to the firewall and install only those bolts at which the truss assembly attach points are flush with the firewall (Ref. Figure 201). Torque these bolts 365 to 390 inch-pounds. If at least one thread does not protrude through the barrel nut, remove the AN960-616 washer from the bolt and torque as above. (2) Check the remaining attach points. If all attach points are flush with the firewall, the attach bolts may be installed at these points and torqued 365 to 390 inch pounds (Ref. Table 202). (3) If the remaining attach points are not flush with the firewall, measure the gap. If the gap does not exceed 0.100 inch, install a shim (of 321 stainless steel), or combination of shims equating the proper thickness, between the truss and the firewall (Ref. Table 201). The bolt, then, is to be installed and torqued as called out in Step (1). If the gap exceeds 0.100 inch between the truss assembly and the firewall, contact Hawker Beechcraft Customer Service for evaluation and recommended action. (4) If the engine truss assembly is removed from the firewall to install the engine after aligning the truss assembly with the firewall, install the engine into the engine truss assembly and install the truss assembly and engine up to the firewall assembly using the shims that were established for proper alignment on the engine truss assembly. (5) If the engine and engine truss assembly was built-up off the airplane without checking proper alignment with the firewall, the engine and truss assembly must be positioned up to the firewall and those bolts at which the truss attach points are flush with the firewall are to be installed. Remove the attach bolts to the engine isolator and engine truss, lift the engine clear of the engine truss assembly, then perform the procedure to check for proper alignment with the firewall as noted in Steps (1), (2), and (3). (6) Check the fit of the engine cowling to ensure that no binding or twisting is prevalent. If any binding or twisting is noted, repeat the above procedure. Table 201 ENGINE TRUSS TO FIREWALL SHIMS Material

Specification

Product

Vendor

0.020 inch

MIL-S-18729

129-910032-121

Hawker Beechcraft Corporation

0.025 inch

MIL-S-18729

129-910032-123

Hawker Beechcraft Corporation

0.032 inch

MIL-S-18729

129-910032-125

Hawker Beechcraft Corporation

0.040 inch

MIL-S-18729

129-910032-127

Hawker Beechcraft Corporation

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Table 202 ENGINE INSTALLATION, BUILD-UP AND TRUSS MOUNT TORQUES ITEM

TORQUES

Engine Truss Mount and Engine Mount. Installs engine truss to engine mount (vibration isolator).

Barry Mounts (95596) - Torque 500 to 600 inch-pounds. Barry Mounts (93880) - Torque 450 to 500 inch-pounds. Lord Mounts (LM-421-SA25) Torque 450 to 500 inch-pounds. Refer to the ENGINE INSTALLATION procedure in Chapter 71-00-00.

Engine Mount (vibration isolator). Installs the engine mount (vibrator isolator) to engine case.

Torque 225 to 300 inch-pounds. Refer to the ENGINE INSTALLATION procedure in Chapter 71-00-00.

Engine Truss Mount Assembly. Installs the engine truss to the firewall.

Torque 365 to 390 inch-pounds. Refer to the TRUSS ASSEMBLY INSTALLATION.

B. Engine Truss Bolt Torque Check NOTE: This procedure provides data on checking the torque of the truss bolts on a previously installed truss. The weight of the engine and propeller does not change this procedure therefore it is not necessary to support the engine during the torque check. Perform this procedure on each of the four bolts attaching the truss to the nacelle structure, aft of the firewall. (1) Check that the torque applied to bolt (7) is 365 to 390 inch-pounds. (Ref. Figure 201 and Table 202). (2) Verify that at least one thread of bolt (7) protrudes through the barrel nut. (3) If bolt (7) does not protrude through the barrel nut, verify that an MS20002C6 washer (6) is installed under the head of the bolt. If an additional AN960-616 washer is installed on the bolt head, remove the flat washer. If one thread of the bolt still does not protrude through the barrel nut, replace the bolt. (4) Verify that bolt (7) does not touch the barrel nut retainer. (5) If bolt (7) touches the barrel nut retainer, verify that a MS20002C6 washer (6) is installed under the head of the bolt. Verify that an AN960-616 or AN960-616L washer (3) is installed on the bolt. (6) If washer (3) and an AN960-616L washer are installed on bolt (7) and the bolt still touches the barrel nut retainer, replace the bolt.

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Figure 201 Engine Mount

71-20-00

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2. INSPECTION CRITERIA A. Isolators Only The following inspection criteria are intended only to be a general rule of thumb when evaluating vibration isolators. If a particular isolator pair does not quite meet any single criterion, but has a large percentage of several, it would be prudent to change out the isolator. The important point to bear in mind when evaluating isolators for replacement is that degradation due to repeated torque loadings tends to be a gradual phenomenon. The presence of cracks in any of the zones listed does indicate some degradation, but the rate of decrease in stiffness does not accelerate appreciably until a very high percentage of the elastomer/plate interface is visibly separated (Ref. Figure 202). (1) Figure 202, Zone 1: The elastomer separation from the outer (cast) plate must not exceed 75% of the region along the radius of the isolator (shaded area). (2) Figure 202, Zone 2: The elastomer separation from the inner (stamped) plate or cracks in the elastomer along the length of the elastomer/plate interface must not exceed 75% of this distance. (3) Figure 202, Zone 3: Any transverse tears (plate to plate) of the elastomer must not exceed 50% of the elastomer thickness: zone 3 is the most likely place for this to occur. (4) Figure 202, Zone 4: If tears to any degree exist in Zones 1, 2 or 3 and if elastomer separation from the inner plate approaches 100% of the area along the radius of the isolator, the isolator must be replaced.

Figure 202 Vibration Isolator Inspection Criteria

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

POWER PLANT FIRESEALS DESCRIPTION AND OPERATION

71-30-00 00

1. GENERAL The fireseals are fitted to the engine, one forward and the other aft of the engine compressor intake. Each fireseal is constructed of semicircular sections, which are bolted to the engine fireseal flange and to each other to form a complete fireseal. The fireseals also provide a mounting location and support for all lines, controls, and ducts that pass from one engine fire zone to another.

71-30-00

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POWER PLANT ELECTRICAL HARNESS DESCRIPTION AND OPERATION

71-50-00 00

1. GENERAL The power plant electrical harness distributes power to or from the starter-generator, fire detectors, and all engine electrical accessories. The harness is connected at the engine firewall by 2 connectors. The harness is routed from the engine firewall, through the rear and forward fireseals to the reduction gear case and propeller electrical systems.

71-50-00

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

POWER PLANT ENGINE DRAINS DESCRIPTION AND OPERATION

71-70-00 00

1. GENERAL A. Engine Fuel Purge System The fuel purge system is designed to ensure that any residual by fuel in the fuel manifolds is consumed during engine shutdown. During engine operation, compressor discharge air (P3 air) is routed through a filter and check valve pressurizing a small air tank mounted on the engine truss mount. On engine shutdown the pressure differential between the air tank and fuel manifolds causes air to be discharged from the air tank, through a check valve and into the fuel manifold system. The air forces all residual fuel, remaining in the fuel manifolds, out through the nozzles and into the combustion chamber (Ref. Figure 1).

71-70-00

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Figure 1 Engine Fuel Purge System

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POWER PLANT ENGINE DRAINS MAINTENANCE PRACTICES

200200

1. FUEL PURGE TANK A. Removal (1) Remove the bleed air line, filter and check valve from the inlet port of the tank (Ref. Figure 201). (2) Remove the check valve and tube from the outlet port of the tank. (3) Remove the four bolts, nuts and two clamps which secure the tank to the mounting plate. (4) Remove the tank from the airplane.

B. Installation NOTE: Replace all packings removed from this assembly with new packings. (1) Install the fitting and packing (if removed) in the outlet port of the tank. Install the check valve at the outlet port (Ref. Figure 201). (2) Install the check valve and packing in the inlet port of the air tank. Install the packing and filter in the inlet port of the check valve. (3) Position the air tank on the mounting place in the original position. Secure the air tank to the mounting plate with the two clamps, four bolts, washers and nuts. (4) Connect the bleed air line to the filter on the inlet side, two clamps, four bolts, washers and nuts. (5) Connect the tube leading to the flow divider to the check valve on the outlet side of the air tank.

C. Cleaning (1) Clean the fuel purge air tank as follows: (a) Perform the FUEL PURGE TANK REMOVAL procedure in this section. WARNING: Wear rubber gloves and eye protection when using solvents. (b) Flush the air tank with solvent (2, Table 1, Chapter 91-00-00) and blow dry with shop air. (c) Perform the FUEL PURGE TANK INSTALLATION procedure in this section.

2. FUEL PURGE SYSTEM AIR FILTER A. Cleaning (1) Disconnect the bleed air line from the filter (Ref. Figure 201). (2) Remove the filter from the check valve. WARNING: Wear rubber gloves and eye protection when using solvents.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (3) Inspect the filter for rust and corrosion; clean with shop air or solvent (2, Table 1, Chapter 91-00-00). A filter which shows signs of rust or corrosion must be replaced. (4) Connect the filter to the check valve. Install a new packing between the filter and check valve. (5) Connect the bleed air line to the filter.

3. FUEL PURGE SYSTEM CHECK VALVE A. Inspection, Cleaning and Leakage Test NOTE: This procedure applies to both of the check valves located adjacent to the engine fuel purge tanks (Ref. Figure 201). CAUTION: Do not disassemble check valve. WARNING: Perform the following procedures in a protected area to avoid hazard to personnel. For inspection and cleaning intervals (Ref. Chapter 5-20-00). (1) Remove check valve. (2) Inspect for any foreign material and/or corrosion. WARNING: Wear rubber gloves and eye protection when using solvents. (3) Using solvent (2, Table 1, Chapter 91-00-00), pressure flush the check valve to remove carbon particles or sludge residue. This can be accomplished by using a set up similar to that in Figure 202. (4) It is permissible to blow dry the check valve using shop air (80 to 120 psi) after assuring that all of the solvent is drained out. (5) After cleaning as described in Step (3), use a test setup to test the check valve for leakage (Ref. Figure 203). (6) Apply a filtered regulated air source (40 to 100 psi) to the upstream side of the check valve. (7) Fill the downstream side of the check valve with water and look for signs of leakage (bubbles). No leakage is allowed. If leakage is detected, repeat Steps (3) through (7). Replace the check valve if leakage is detected the second time. (8) It is permissible to blow dry the check valve with shop air (80 to 120 psi).

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Figure 201 Engine Fuel Purge System

71-70-00

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Figure 202 Check Valve Pressure Flush

Figure 203 Check Valve Leakage Test

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4. FUEL FLOW DIVIDER/PURGE VALVE A. Inspection This procedure inspects the engine fuel flow divider/purge valve for possible fuel leakage from the purge valve into the fuel purge system (Ref. Figure 204). A leaking purge valve may result in fuel leaking into the fuel purge canister if the check valves fail, or are installed incorrectly. No fuel leakage is allowed. (1) Remove the upper and lower forward cowlings to gain access to the engine flow divider/purge valve (Ref. Chapter 71-10-00, COWLING REMOVAL). (2) Disconnect the fuel purge line from the flow divider. (3) Perform a Wet Motoring Run in accordance with Pratt & Whitney Engine Maintenance Manual. Observe all Cautions, Warnings and Notes. (4) Observe purge valve port for leakage while the engine is motoring. Any leakage while motoring is cause for rejection of unit. NOTE: Small drops of fuel may exit the flow divider purge valve shortly after motoring stops. (5) Install the fuel purge line. (6) Allow the starter a cooling period then follow with a Dry Motoring Run in accordance with Pratt & Whitney Engine Maintenance Manual. Observe all Cautions, Warnings and Notes. (7) Install the upper and lower forward cowlings (Ref. Chapter 71-10-00, COWLING INSTALLATION).

71-70-00

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Figure 204 Flow Divider/Purge Valve

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71-70-00

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

CHAPTER 72 - ENGINE TABLE OF CONTENTS SUBJECT

PAGE

GENERAL INFORMATION 72-00-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1

72-CONTENTS

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List of Effective Pages CH-SE-SU

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72-CONTENTS

1

Nov 1/09

72-00-00

1

Nov 1/09

72-LOEP

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ENGINE GENERAL INFORMATION DESCRIPTION AND OPERATION

72-00-00 00

1. GENERAL NOTE: Refer to the PT6A-65B Engine Maintenance Manual P/N 3032842 for detailed information on these subjects.

72-00-00

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CHAPTER 73 - ENGINE FUEL SYSTEMS TABLE OF CONTENTS SUBJECT

PAGE

DISTRIBUTION 73-10-00 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Engine Driven Fuel Pump . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

73-CONTENTS

201 201 201 202

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List of Effective Pages CH-SE-SU

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73-10-00

201 thru 204

Aug 1/10

C3

73-LOEP

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ENGINE FUEL SYSTEMS DISTRIBUTION MAINTENANCE PRACTICES

73-10-00 200200

1. ENGINE-DRIVEN FUEL PUMP A. Removal The firewall shutoff valve must be closed prior to disconnecting any fuel tubing and/or hoses. CAUTION: Observe all fire precautions and safety practices when working on the fuel system (Ref. Chapter 28-00-00) and (Ref. Chapter 12-10-00, FUEL-HANDLING SAFETY INFORMATION). (1) Gain access to the engine-driven fuel boost pump (1) (Ref. Figure 201) by removing the engine cowling as outlined in Chapter 71-10-00. CAUTION: The fire extinguisher system is armed when the FIRE PULL T handle is pulled. Do not actuate the extinguisher push light/switch with the system armed. If the light/switch is pushed, the fire extinguisher agent will be released. (2) Select the battery switch to ON, close the firewall shutoff valve by pulling the appropriate FIRE PULL T handle. (3) Confirm that the shutoff valve is closed by selecting appropriate side electric standby fuel boost pump to ON and confirm that the red fuel pressure annunciator remains illuminated, indicating that the firewall fuel shutoff valve is closed. (4) Select the boost pump to the OFF position. (5) Select the battery switch to OFF. (6) Perform REMOVING GROUND POWER procedures (Ref. Chapter 24-40-00). Display warning notices prohibiting connection of airplane electrical power. (7) Perform BATTERY DISCONNECTION procedure (Ref. Chapter 24-31-00). (8) Make provisions for catching spilled fuel (rags to absorb fuel). NOTE: Removal of the left-hand and right-hand engine-driven fuel boost pump is typical. CAUTION: Failure to use a back-up wrench when loosening or tightening hoses to fittings may damage the fittings. (9) At the inboard forward firewall, place a container under the fuel line connections. Disconnect the pump outlet line from the inboard forward firewall fitting, and the engine fuel supply line from the main fuel filter and drain the residual fuel from the lines. After draining the fuel, reconnect the lines. (10) Disconnect the engine fuel supply line (14) from the pump inlet fitting (13 or 13A) (Ref. Figure 201). (11) Disconnect the pump outlet line (17) from the pump outlet fitting (2 or 2A). (12) Cap or plug the fuel fitting openings to prevent the entry of foreign matter.

73-10-00

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (13) Remove the hose clamps (11) and disconnect the pump seal drain hose (12) from the hose adapters (10 and 16). (14) Remove four bolts (7A) or nuts (7B) and washers (6) securing the engine-driven fuel boost pump (1) to the number four drive pad (5) of the engine. (15) Remove the engine-driven fuel boost pump (1) from the engine.

B. Installation (1) If a new engine-driven fuel boost pump (1) is to be installed, accomplish the following (Ref. Figure 201): (a) Remove the pump inlet and outlet fittings (13 or 13A and 2 or 2A) and the pump seal hose adapter (10) and the elbow (15) from the old pump (1). CAUTION: Failure to use a back-up wrench when loosening or tightening hoses to fittings may damage the fittings. (b) Install the inlet and outlet fittings (13 or 13A and 2 or 2A) and nuts (3) with new packings (4) on the new pump (1). (c) Install the seal drain hose adapter (10) with a new packing (8) and jam nut (9) on the new pump (1). (d) Install the elbow (15) along with adapter (16) on the regulator housing of the pump (1). (2) Position the pump (1) on the engine number four drive pad (5). (3) Secure the pump (1) with the four bolts (7A) or nuts (7B) and washers (6). Torque the bolts (7A) 40 to 50 inch-pounds and safety wire. (4) Connect the pump seal drain hoses (12) to the hose adapter (10 and 16) and install the hose clamps (11). CAUTION: Failure to use a back-up wrench when loosening or tightening hoses to fittings may damage the fittings. (5) Connect the pump outlet line (17) to the pump outlet fitting (2 or 2A) on the pump (1). Using a back-up wrench on fitting (2 or 2A), torque the pump outlet line (17) b-nut 300 to 450 in.-lbs and nut (3) 285 to 315 in.-lbs. (6) Connect the engine fuel supply line (14) to the pump inlet fitting (13 or 13A) on the pump (1). Using a back-up wrench on fitting (13 or 13A), torque the engine fuel supply line (14) b-nut 300 to 450 in.-lbs and nut (3) 285 to 315 in.-lbs. (7) Remove warning notices prohibiting connection of airplane electrical power and perform BATTERY CONNECTION procedure (Ref. Chapter 24-31-00). (8) Perform APPLYING GROUND POWER procedures (Ref. Chapter 24-40-00). (9) Actuate the appropriate FIRE PULL T handle to open the firewall shutoff valve. NOTE: Confirm the firewall shutoff valves are open by observing the red fuel pressure annunciator extinguishes while performing the next Step.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (10) Operate the appropriate standby fuel boost pump to pressurize the fuel lines. (11) Check all fuel lines and fittings for fuel leaks. (12) Turn the standby fuel boost pump off and remove electrical power from the airplane. (13) Install the engine cowling as outlined in Chapter 71-10-00.

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Figure 201 Engine-Driven Fuel Boost Pump Page 204 Aug 1/10

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CHAPTER 74 - IGNITION TABLE OF CONTENTS SUBJECT

PAGE

IGNITION SYSTEM 74-00-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Autoignition System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Spark Igniters . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Igniter Plug . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Cleaning and Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Autoignition Pressure Switch . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 204 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 204 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 204 AutoIgnition Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 205 Ignition Exciter Control Box . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 205 Removal (UA-1 and After; UB-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 205 Installation (UA-1 and After; UB-1 and After) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 206 Removal (UC-1 and After without Kit No. 129-9100 Installed) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 208 Installation (UC-1 and After without Kit No. 129-9100 Installed) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 208 Removal (UC-1 and After with Kit No. 129-9100 Installed) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 211 Installation (UC-1 and After with Kit No. 129-9100 Installed) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 211

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IGNITION IGNITION SYSTEM DESCRIPTION AND OPERATION

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1. GENERAL The spark ignition system of the PT6A-65B engine provides quick light-ups over a wide temperature range. The system includes a solid state ignition exciter designed to be mounted on the airframe, two individual high tension cable assemblies, and two spark igniters. Although designed to be energized by the 28 VDC electrical system of the airplane, the ignition system will function effectively over a range of 9 to 30 volts. The sealed unit of the ignition exciter encases in epoxy resin the solid state circuitry, transformer, and diodes required to transform the DC input to a high voltage output. When the ignition exciter is energized, a capacitor is progressively charged until the energy stored is sufficient to ionize a spark gap in the unit and discharge the capacitor across the two spark igniters through a dividing and step-up transformer network. The network is designed so that one igniter will continue to function even though the other is open or shorted. The network also enables the capacitor to discharge automatically should either or both igniters become inoperative, or should input voltage be switched off. The electrical energy output from the ignition exciter to the spark igniters on the engine is delivered by two ignition cable assemblies, each consisting of an electrical lead contained in flexible metal braiding. The cables are equipped with coupling nuts at each end for connection to the ignition exciter and spark igniter. The spark igniters are mounted adjacent to the fuel manifold at the 4 and 9 o'clock positions on the gas generator case. Each spark igniter is a double-ended, threaded plug with a central positive electrode enclosed in annular semiconducting material. The electrical potential produced by the ignition exciter is applied across the gap between the central conductor and the igniter shell (ground). A small current passes across the semiconducting material of the spark igniter as the potential developed by the ignition exciter increases. This current increases until the air between the central conductor and the shell ionizes into a high energy discharge between the electrodes. The resultant spark always occurs at some point in the annular space between the central conductor and its shell.

2. AUTOIGNITION SYSTEM The autoignition system is designed to provide automatic ignition when engine torque falls too low. This system ensures ignition during take-off, landing, turbulence, and during adverse conditions of icing weather. The autoignition control switch on the left subpanel is wired to a pressure switch mounted adjacent to the torque pressure transmitter just forward and above the exhaust outlet on each engine (Ref. Figure 1). This switch monitors torquemeter oil pressure. When the control switch is in the ARM position and engine torque drops below 400 foot-pounds, the pressure switch closes to provide power from a 5-ampere circuit breaker on the right circuit breaker panel to energize the engine igniter and illuminate the IGNITION ON light (green) in the annunciator panel located forward of the pedestal. The igniter will continue to function until engine torque increases enough to actuate the pressure switch and break the circuit to the igniter and annunciator lights. For extended ground operation, the autoignition system should be turned off to prolong the service life of the igniter units.

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Figure 1 Autoignition System

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IGNITION IGNITION SYSTEM MAINTENANCE PRACTICES

200200

1. SPARK IGNITERS WARNING: Residual voltage in the ignition exciter may be dangerously high. If a spark igniter is to be removed, remove power from the ignition system and ensure the system has been inoperative for at least six minutes prior to removing any ignition components. Always disconnect couplings at the exciter first and use insulated tools to remove the nuts. Do not touch output connectors or coupling nuts with bare hands. To check the spark igniters for proper operation, disconnect one lead at the ignition exciter and listen for the discharge of the spark igniter that is still connected, then repeat the procedure for the other spark igniter. The igniters will spark at a rate of 0.8 to 1.0 spark per second with 9 volts applied and up to a rate of 1.4 to 4.0 sparks per second with 30 volts applied.

2. IGNITER PLUG A. Cleaning and Inspection The spark igniters, located at the 4 o'clock and 9 o'clock positions on the gas generator case, are air-cooled, threaded plugs with a central positive electrode enclosed in an annular semiconducting shell. The plug should be removed, cleaned and inspected as follows: WARNING: Residual voltage in the ignition exciter may be dangerously high. Remove power from the ignition system and ensure the system has been inoperative for at least six minutes prior to removing any ignition components. Always disconnect couplings at the exciter first and use insulated tools to remove the nuts. Do not touch output connectors or coupling nuts with bare hands. (1) Remove the coupling nut on the ignition cable from the spark igniters. (2) Remove the spark igniters from the gas generator case. (3) Remove the copper gasket from the spark igniter. NOTE: The firing end of the igniters should never be cleaned. Do not remove carbon from the electrodes or annular gap areas because the carbon is an aid to igniter operation. (4) Using a felt swab soaked in solvent (30, Table 1, Chapter 91-00-00), clean the inside surface of the terminal. (5) Dry the igniter with clean, dry compressed air. (6) Ensure that the air cooling holes are not blocked. (7) Inspect the igniter shell for wear. If the wear equals or exceeds the amount shown, replace the spark igniter (Ref. Figure 202). CAUTION: It is recommended that if an igniter plug is dropped, it should be replaced with a new plug. Do not use thread lubricants on the spark igniters.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (8) Install a new copper gasket on the spark igniter. (9) Install spark igniter in the boss on the gas generator case. Torque the igniters to 300 inch-pounds, loosen to zero inch-pounds, and retighten 200 to 240 inch-pounds. (10) Install the coupling nut on the ignition cable.

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Figure 201 Autoignition System

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Figure 202 Spark Ignitor - Wear Limits

3. AUTOIGNITION PRESSURE SWITCH A. Removal This switch is mounted adjacent to the torque pressure transmitter just forward and above the left exhaust outlet on each engine (Ref. Figure 201). NOTE: A single double-pole, double-throw switch functions both as the autoignition pressure switch and the high pressure switch for the autofeathering system. (1) Place the battery and generator switches located on the LH subpanel in the OFF position. (2) Be certain the external power source is disconnected. (3) Remove the upper forward cowling (Ref. Chapter 71-10-00, COWLING REMOVAL). (4) Unsafety and remove the electrical connector from the switch. (5) Unsafety the mounting nut at the base of the switch. Remove the switch from the airplane. (6) Install a protective plug in the switch mounting hole to prevent contamination of the governor oil supply.

B. Installation (1) Remove the protective plug from the switch mounting pole. (2) Position the new packing and install the switch in the hole. Tighten the switch mounting nut to seat the packing seal.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (3) Safety wire the nut on the switch. (4) Connect the electrical plug. The plug is to be tightened finger tight, then safety wired. (5) Install the upper forward cowling (Ref. Chapter 71-10-00, COWLING INSTALLATION).

C. AutoIgnition Check (1) With both engines running, set the power levers at idle and depress the autoignition circuit breakers and the bus feeder circuit breakers. The autoignition circuit breakers are located on the copilot's right sidewall circuit breaker panel. (2) Place the LH and RH autoignition switches in the ARM position. Both IGNITION ON (green) lights in the annunciator panel should illuminate. The annunciator panel is located forward of the pedestal (Ref. Figure 201). (3) Advance the RH power lever until the R IGNITION ON light goes out and check that RH engine torque pressure is between 450 to 650 foot-pounds. (4) Retard the RH power lever until the R IGNITION ON light illuminates and check that RH engine torque pressure is below 650 foot-pounds. (5) Pull the RH circuit breaker and ascertain that the R IGNITION ON light goes out. (6) Depress the RH ignition circuit breaker and ascertain that the R IGNITION ON light illuminates. (7) Repeat the procedures in Steps (2) through (6) with the LH engine autoignition system.

4. IGNITION EXCITER CONTROL BOX A. Removal (UA-1 and After; UB-1 and After) NOTE: The following procedure is applicable to both engines. The ignition exciter control box is located on the forward side of the engine firewall, near the nacelle centerline (Ref. Figure 203). (1) Place the battery and generator switches, located on the pilot’s LH subpanel, in the OFF position. (2) Be certain the external power source is disconnected. (3) Remove the upper aft cowling from the nacelle (Ref. Chapter 71-10-00, COWLING REMOVAL). (4) Tag the igniter cables (6) to facilitate installation (Ref. Figure 203). (5) Remove the safety wire from the igniter cable (6) connectors and disconnect the cables. (6) Disconnect the power supply cable (3). (7) Cap the electrical connectors to protect from damage and contamination. (8) Remove the four bolts (5) and washers (4) and remove the ignition exciter control box (2).

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B. Installation (UA-1 and After; UB-1 and After) NOTE: The following procedure is applicable to both engines. (1) Place the battery and generator switches, located on the pilot’s LH subpanel, in the OFF position. (2) Be certain the external power source is disconnected. (3) Ensure the surface of the firewall (1) (Ref. Figure 203) is properly prepared to provide adequate electrical bonding (Ref. Chapter 20-00-01 ELECTRICAL BONDING PROCEDURES). (4) Position the ignition exciter control box (2) and install the two top bolts (5) and washers (4) (Ref. Figure 203). NOTE: Include the two wire harness clamps (7) in the assembly when installing the two bottom bolts (5) and washers (4). (5) Install the two bottom bolts (5) and washers (4). (6) Remove the protective caps from the electrical connectors. CAUTION: Do not let lubricant contact the center conductor of the exciter connectors. Contact may result in a high resistance which could generate heat and oxidation. (7) Lightly coat the threads of the ignition exciter control box (2) connectors with lubricant (169, Table 1, Chapter 91-00-00). CAUTION: When loosening or tightening the coupling nuts do not let the ignition cable braiding or ferrules turn at the same time. (8) Observing the identification tags, connect the igniter cables (6) and the power supply cable (3) to the ignition exciter control box (2) so they are in the same positions as before removal. Tighten nuts finger tight plus 45 degrees and safety wire. (9) Install the upper aft cowling of the nacelle (Ref. Chapter 71-10-00, COWLING INSTALLATION).

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Figure 203 Ignition Exciter Control Box Installation (UA-1 and After, UB-1 and After)

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C. Removal (UC-1 and After without Kit No. 129-9100 Installed) NOTE: The following procedure is applicable to both engines. The ignition exciter control box is located on the aft side of the engine firewall/baffle, near the nacelle centerline, below the engine accessories (Ref. Figure 204). (1) Place the battery and generator switches, located on the pilot’s LH subpanel, in the OFF position. (2) Be certain the external power source is disconnected. (3) Remove the upper aft cowling from the nacelle (Ref. Chapter 71-10-00, COWLING REMOVAL). (4) Tag the igniter cables (7) to facilitate installation (Ref. Figure 204). (5) Remove the safety wire from the igniter cable (7) connectors and disconnect the cables. (6) Disconnect the power supply cable (2). (7) Cap the electrical connectors to protect from damage and contamination. (8) Remove the four bolts (3) and washers (4) and remove the ignition exciter control box (5).

D. Installation (UC-1 and After without Kit No. 129-9100 Installed) NOTE: The following procedure is applicable to both engines. (1) Place the battery and generator switches, located on the pilot’s LH subpanel, in the OFF position. (2) Be certain the external power source is disconnected. (3) Ensure the surface of the firewall (1) (Ref. Figure 204) is properly prepared to provide adequate electrical bonding (Ref. Chapter 20-00-01 ELECTRICAL BONDING PROCEDURES). (4) Position the ignition exciter control box (5) and install the two top bolts (3) and washers (4) (Ref. Figure 204). NOTE: Include the two wire harness clamps (6) in the assembly when installing the two bottom bolts (3) and washers (4). (5) Install the two bottom bolts (3) and washers (4). (6) Remove the protective caps from the electrical connectors. CAUTION: Do not let lubricant contact the center conductor of the exciter connectors. Contact may result in a high resistance which could generate heat and oxidation.

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(7) Lightly coat the threads of the ignition exciter control box (5) connectors with lubricant (169, Table 1, Chapter 91-00-00). CAUTION: When loosening or tightening the coupling nuts do not let the ignition cable braiding or ferrules turn at the same time. (8) Observing the identification tags, connect the igniter cables (7) and the power supply cable (2) to the ignition exciter control box (5) so they are in the same positions as before removal. Tighten nuts finger tight plus 45 degrees and safety wire. (9) Install the upper aft cowling of the nacelle (Ref. Chapter 71-10-00, COWLING INSTALLATION).

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Figure 204 Ignition Exciter Control Box Installation (UC-1 and After without Kit No. 129-9100 Installed)

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E. Removal (UC-1 and After with Kit No. 129-9100 Installed) NOTE: The following procedure is applicable to both engines. The ignition exciter control box is located on the forward side of the engine firewall/baffle, near the nacelle centerline (Ref. Figure 205). (1) Place the battery and generator switches, located on the pilot’s LH subpanel, in the OFF position. (2) Be certain the external power source is disconnected. (3) Remove the mid side cowling from the nacelle (Ref. Chapter 71-10-00, COWLING REMOVAL). (4) Tag the igniter cables (2) to facilitate installation (Ref. Figure 205). (5) Remove the safety wire from the igniter cable (2) connectors and disconnect the cables. (6) Disconnect the power supply cable (6). (7) Cap the electrical connectors to protect from damage and contamination. (8) Remove the four screws (4) and washers (5) and remove the ignition exciter control box (1).

F. Installation (UC-1 and After with Kit No. 129-9100 Installed) NOTE: The following procedure is applicable to both engines. (1) Place the battery and generator switches, located on the pilot’s LH subpanel, in the OFF position. (2) Be certain the external power source is disconnected. (3) Ensure the surface of the firewall (3) (Ref. Figure 205) is properly prepared to provide adequate electrical bonding (Ref. Chapter 20-00-01 ELECTRICAL BONDING PROCEDURES). (4) Position the ignition exciter control box (1) and install the two top screws (4) and washers (5) (Ref. Figure 205). NOTE: Include the two wire harness clamps (7) in the assembly when installing the two bottom screws (4) and washers (5). (5) Install the two bottom screws (4) and washers (5). (6) Remove the protective caps from the electrical connectors. CAUTION: Do not let lubricant contact the center conductor of the exciter connectors. Contact may result in a high resistance which could generate heat and oxidation.

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(7) Lightly coat the threads of the ignition exciter control box (1) connectors with lubricant (169, Table 1, Chapter 91-00-00). CAUTION: When loosening or tightening the coupling nuts do not let the ignition cable braiding or ferrules turn at the same time. (8) Observing the identification tags, connect the igniter cables (2) and the power supply cable (6) to the ignition exciter control box (1) so they are in the same positions as before removal. Tighten nuts finger tight plus 45 degrees and safety wire. (9) Install the mid side cowling of the nacelle (Ref. Chapter 71-10-00, COWLING INSTALLATION).

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6. POWER SUPPLY CABLE 7. WIRE HARNESS CLAMP (2 PLACES)

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Figure 205 Ignition Exciter Control Box Installation (UC-1 and After with Kit No. 129-9100 Installed)

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CHAPTER 76 - ENGINE CONTROLS TABLE OF CONTENTS SUBJECT

PAGE

ENGINE CONTROLS 76-00-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Power Levers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Propeller Controls . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Condition Levers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Controls . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Removal (Power, Propeller, and Condition Levers) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Power Lever Assembly Teardown . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .202 Power Lever Assembly Build Up . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .202 Installation (Power, Propeller, and Condition Levers) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202 Power Lever Detent Pin . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 203 Inspection and Replacement (UA-1 and After, UB-1 and After, UC-1 and After without Kit 129-5009 Installed) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 203 Inspection and Replacement (UA-1 and After, UB-1 and After, UC-1 and After without Kit 129-5009 Installed) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .204 Engine Control (Power, Propeller and Condition) Cable . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 206 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 206 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 207 Engine Control . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 208 Rigging . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 208 Power Levers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 208 Rigging . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 208 Condition Levers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 210 Rigging . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 210 Propeller Levers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 210 Rigging . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 210 Propeller Ground-fine Solenoid . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 211 Rigging . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 211 Ground-fine . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 211 Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 211 Engine Final Adjustments . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 217 Ground Performance Check Procedure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 217 Engine Operating Parameters Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 219 Estimated Field Barometric Pressure Calculation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 220 Engine Power . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 226 Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 226 Fuel Cutoff . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 227 Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 227 Low Idle . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 228 Adjustment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 228 Engine Torque . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 228 Adjustment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 228 High Idle . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 228 Adjustment (Condition Levers Not Split) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 228 Adjustment (Condition Levers Split) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 229 Maximum Forward and Reverse N1 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 229 Adjustment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 229 Propeller Governor . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 229

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CHAPTER 76 - ENGINE CONTROLS TABLE OF CONTENTS (CONTINUED) SUBJECT PAGE Adjustment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 229 Power Levers . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 229 Alignment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 229 Control Cable Repairs and Checks . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 230 Equipment Required . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 230 Leak Check Procedure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 230 Repairing Control Cables . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 231 Cable Purging . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 232

POWER CONTROL 76-10-00 Maintenance Practices. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Condition Control Catch Gate . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

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ENGINE CONTROLS ENGINE CONTROLS DESCRIPTION AND OPERATION

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1. GENERAL Three engine controls for each engine are mounted on the pedestal between the pilot’s and copilot’s control columns. The engine controls control all adjustable aspects of the operating engine. The controls from left to right on the pedestal are the power levers, propeller levers and condition levers. The different levers are also distinguished from one another by the shape of the knobs on each lever (Ref. Figure 1).

2. POWER LEVERS The power levers control the amount of power the engines are delivering to the propellers. This function (N1) is measured by means of a percentage of maximum gas generator speed. Propeller reversing and reverse power are controlled by the power levers as well. A preset idle speed can be selected through detents on the pedestal control mount. A ground fine detent provides for a further flattening of the propeller blade angle during ground roll, achieving more efficient aerodynamic braking. There should be no change in N1 when the power lever is moved through the idle detent and into the ground fine detent. The idle and ground fine range should produce an N1 of 59% ± 1%. The power levers control the reversing mechanisms by means of linkages attached to the power lever cambox assemblies, mounted forward of the fuel control unit on the rear of the engine. A linkage from the cambox assembly is connected to the beta valve actuating arm, located on the propeller governor. When the power levers are moved through the ground fine detent into the reversing range, the beta valve position is changed, allowing the propeller blades to move from the ground fine blade angle into the reverse pitch blade angle. A control linkage from the cambox assembly is connected to the fuel control unit providing control movements to the fuel control unit. When the power lever is in the maximum reverse position, a tab on the fuel control unit contacts the maximum reverse power stop and the lost motion link (airbleed link), through the beta cable interconnect, resets maximum reverse propeller speed.

3. PROPELLER CONTROLS The propeller control levers are connected to the speed adjusting lever of the propeller governor through control linkages. The propeller governor is located in the 12-o’clock position on the forward portion of the reduction gearbox. High rpm, low rpm and feathered markings on the pedestal indicate the forward and aft positions of the propeller levers. Propeller and governor operation is discussed in Chapter 61 of this maintenance manual.

4. CONDITION LEVERS Control linkages connect the fuel condition levers to lever and stop mechanisms on top of the fuel control units. High idle, low idle and cutoff markings on the pedestal allow for the preset selection of the various fuel conditions. When the condition lever is in the CUTOFF position, the fuel condition cam moves against the cutoff stop and actuates a lever depressing the fuel cutoff valve on the fuel control unit. Gas generator speed is controlled by the amount of fuel metered into the engine. Metering of the fuel is a function of valve position in the fuel control unit and of P3 bleed air pressure.

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Figure 1 Engine Control Levers

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ENGINE CONTROLS CONTROLS MAINTENANCE PRACTICES

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1. CONTROLS A. Removal (Power, Propeller, and Condition Levers) (1) Remove the copilot’s seat assembly (Ref. Chapter 25-10-00). (2) Remove the screw and washer that secures the elevator trim tab control wheel and the left end of the engine control lever pivot shaft to the left side of the pedestal. Note the position of the elevator trim tab control wheel for installation (Ref. Figure 201). (3) Remove the power, propeller, and condition knobs and the friction brake knobs. (4) Remove the edgelight panel and the printed circuit board from the pedestal. Retain all attaching parts. (5) Remove the upholstery panel from the RH side of the pedestal and remove the small instrument panel forward of the pedestal control levers. Retain all attaching parts. (6) Remove the four screws that secure the RH end of the control lever pivot shaft to the right side of the pedestal. Retain the screws. (7) Note the quantity of washers and spacers between the control lever assemblies for reassembly purposes. NOTE: A drop cloth should be used to keep items dropped from the pedestal from falling below the floor. (8) Fabricate a dummy control lever pivot shaft 0.495-inch in diameter and at least 9 inches long. (9) Insert the dummy shaft into the control shaft opening in the left side of the pedestal, pushing the control lever pivot shaft to the right until the dummy shaft is ready to enter the RH or LH control lever which is to be removed. NOTE: If more than one control lever is to be replaced, it is best to remove and replace them individually, RH first, then LH. (10) Disconnect the pushrod from the lower connection to the control lever to be removed. Retain the attaching parts. (11) Catching the spacer washers that fall, pull the dummy shaft back to the left until the control lever assembly to be removed is released. Remove the assembly from the pedestal. (12) Disassemble the control lever from the arm assembly by disconnecting the tension spring, removing the nuts and washers and one cotter pin and washer from the three pins in the slotted holes.

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B. Power Lever Assembly Tear Down (1) Remove the tension spring (2) from pins (9) and (15) (Ref. Figure 214). (2) Remove cotter pin (8), nut (6) or washer (7) and washer (5) from pin (1) and discard cotter pin (8). (3) Remove nut (10) and washer (11) from pin (13). (4) Remove cotter pin (18), washers (19) and (20) from pin (17) and discard cotter pin (18). (5) Separate the power lever (16) from the arm assembly (4) and remove washers (3), (12), (14) and (21).

C. Power Lever Assembly Build Up NOTE: Washer (14) has a coating on one side and should be installed with the coating toward the power lever (16) (Ref. Figure 214). (1) Install washers (3), (12), (14) and (21) on the power lever (16) and arm assembly (4). (2) Mate the power lever (16) and arm assembly (4). (3) Install washers (20), (19) and new cotter pin (18) on pin (17). (4) Install washer (11) and nut (10) on pin (13). (5) Install washer (5), nut (6) or washer (7) and new cotter pin (8) on pin (1). (6) Install the tension spring (2) on pins (9) and (15).

D. Installation (Power, Propeller, and Condition Levers) (1) Assemble the new engine control lever to the arm assembly with the nuts and the washers removed in Step (11) under ENGINE CONTROLS REMOVAL. Use a new cotter pin. (2) Position the complete control lever assembly and the arm and brake assembly in the pedestal and push the control lever pivot shaft through the control lever assembly. Install the spacer washers as noted in Step (7), under ENGINE CONTROLS REMOVAL, until contact is made with the dummy shaft (Ref. Figure 201). (3) Connect the pushrod to the lower end of the control lever with the existing attaching parts. (4) Install the control lever pivot shaft. Install the spacer washers noted in Step (7) under ENGINE CONTROLS REMOVAL, until the dummy shaft is pushed out of the pedestal. (5) Install the four screws which secure the RH end of the control lever pivot shaft to the pedestal. (6) Install the LH and RH pedestal upholstery panels. (7) Install the manual elevator trim control wheel and secure it and the LH end of the control lever pivot shaft with the attaching screw and washer. Position the control wheel as noted in Step (2) under ENGINE CONTROL REMOVAL. (8) Install the small instrument panel just forward of the control levers.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (9) Install the printed circuit board and edgelight panel on the pedestal. Make certain that all lights are positioned properly to avoid breakage. (10) Install all control lever knobs and the friction brake knobs. (11) Check for proper operation of all switches and controls in the pedestal area. (12) Install the copilots seat assembly (Ref. Chapter 25-10-00).

Figure 201 Engine Control Levers

2. POWER LEVER DETENT PIN A. Inspection and Replacement (UA-1 and After, UB-1 and After, UC-1 and After without Kit 129-5009 Installed) (1) Perform the REMOVING GROUND POWER procedure (Ref. Chapter 24-40-00). NOTE: Kit 129-5009 may be installed to upgrade to a replaceable detent pin. (2) Remove the power lever assembly from center pedestal (Ref. ENGINE CONTROLS REMOVAL (POWER, PROPELLER AND CONDITION LEVERS)). (3) Inspect the detent pin. If the pin has a groove of more than 0.03 inch, Install Kit 129-5009 (Ref. Figure 202) or perform the POWER LEVER ASSEMBLY TEAR DOWN procedure in this section and replace the power lever (Ref. Figure 214).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (4) If the power lever assembly was torn down, perform the POWER LEVER ASSEMBLY BUILD UP procedure in this section with the new power lever. (5) Install the power lever assembly in the center pedestal (Ref. ENGINE CONTROLS INSTALLATION (POWER, PROPELLER AND CONDITION LEVER)). (6) Adjust the power levers (Ref. POWER LEVERS RIGGING).

B. Inspection and Replacement (UA-1 and After, UB-1 and After, UC-1 and After with Kit 129-5009 Installed) (1) Perform the REMOVING GROUND POWER procedure (Ref. Chapter 24-40-00). (2) Remove the power lever control assembly from the center pedestal. Refer to ENGINE CONTROLS REMOVAL (POWER, PROPELLER AND CONDITION LEVERS). (3) Remove and inspect the detent pin. If the pin has a groove of more than 0.03 inch, discard the pin. (Ref. Figure 202). (4) Inspect the bearing for wear. Replace if necessary. (5) Assemble the control lever assembly using new hardware as necessary. (6) Install the control lever assembly in the center pedestal. Refer to ENGINE CONTROLS INSTALLATION (POWER, PROPELLER AND CONDITION LEVER). (7) Adjust the power levers (Ref. POWER LEVERS RIGGING).

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Figure 202 Power Lever Detent Pin Replacement

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3. ENGINE CONTROL (POWER, PROPELLER AND CONDITION) CABLE A. Removal (1) Remove the left forward partition. (2) Remove the left flight compartment seat (Ref. Chapter 25-10-00). (3) Remove the flight compartment carpet and center floor panel aft of the pedestal and pedestal side panels. (4) Remove the passenger compartment left or right seats, forward of the main spar (Ref. Chapter 25-20-00). (5) Remove the passenger compartment center aisle and left or right carpet (Ref. Chapter 25-20-01) and floorboards, forward of the main spar (Ref. Chapter 06-50-00). (6) Remove the left or right center wing leading edge panel 54 or 55 (UA-1 and After, UB-1 and After), or 23 or 24 (UC-1 and After) (Ref. Chapter 06-50-00). (7) From the applicable engine, remove the upper forward, upper mid and upper aft engine cowlings (UA-1 and After, UB-1 UB-74 Without Kit No. 114-9016-1 and 114-9016-3) or upper forward, top center and inboard mid-side cowlings (UA-1 and After, UB-1 thru UB-74 With Kit No. 114-9016-1 and 114-9016-3, UC-2 and After) (Ref. Chapter 71-10-00). (8) Disconnect the control cable from its respective control unit input lever by removing the safety wire or cotter pins, bolts, washers and nuts. CAUTION: If cables are to be reinstalled, take appropriate measures to protect control cable threaded ends and seals from damage and contamination. NOTE: When removing a rod end, with a marker, mark a reference line across the rod end and control cables and count the number of turns required to separate the two. Record this information to aid in installation and rigging. Do not allow the control cable to turn. It may be necessary to remove the rod end to clear some frame members under the pedestal. (9) To remove a rod end, loosen the nut and slowly turn the rod end until disconnected, retain the nut and washer. (10) Disconnect the propeller control cable (16) from the clip (1), remove the fire sleeve (11) and grommet assemblies (12, 13, 14 and 15) (Ref. Figure 203). (11) To remove the power control cable, remove U bolt, (7) and the special nut securing the power control cable (9), then remove the deicer boot (6), clamps (8) and spring(s) (5). (12) Remove the grommet assemblies at the firewall (10). (13) Remove the clamps securing the control cable in the wing leading edge and remove sealant from cables. (14) From inside the cabin, remove the control cable clamps between FS 134.00 and 243.25. To remove the propeller control cable (1), remove the lock plate (4) at FS 125.00 (Ref. Figure 204).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL NOTE: When removing a rod end, with a marker, mark a reference line across the rod end and control cables and count the number of turns required to separate the two. Record this information to aid in installation and rigging. Do not allow the control cable to turn. (15) Disconnect the control cable from the pedestal quadrant. If the condition control cables (5) are being changed, disconnect them from bracket (6) and condition levers (Ref. Figure 204). (16) Feed the control cable through the fuselage pressure plate (2) and remove it from the aircraft.

B. Installation CAUTION: Take appropriate measures to protect control cable threaded ends and seals from damage and contamination. (1) Carefully route the engine control cable forward into the crew compartment, through the fuselage pressure plate and into the engine compartment. Place the firewall grommet retainer over the cable once it is in the engine compartment. (2) Install the deicer boot (6) with clamps (8) and the spring(s) (5) (maximum of two) on the power control cable (9) (Ref. Figure 203). (3) If applicable, install the fire sleeve (11) over the propeller control cable (16). (4) If removed, install nuts, washers and rod ends to their proper position (as recorded during removal) then secure the nut. (5) Attach clips on the control cables, as required, by squeezing with pliers. (6) Connect the engine control cables to their respective control levers and input arms and secure (Ref. Figure 201). (7) From inside the cabin, install the control cable clamps between FS 134.00 and FS 243.25 and if the propeller control cable (1) is being installed, install the lock plate (4) at FS 125.00 (Ref. Figure 204). (8) Install clamps securing control cables in the wing leading edge. (9) Apply sealant (19, Table 1, Chapter 91-00-00) around control cables where they pass through the fuselage (2) (Ref. Figure 204). (10) Check rigging, perform POWER LEVERS RIGGING, CONDITION LEVERS RIGGING, PROPELLER LEVERS RIGGING and ENGINE CONTROL RIGGING procedures. (11) Safety hardware as needed (Ref. Chapter 20-07-00). NOTE: Inspect work areas for FOD. before closing panels. (12) Install all removed engine cowlings (Ref. Chapter 71-10-00). (13) Install the left and right center wing leading edge panel 54 to 55 (UA-1 and After, UB-1 and After), or 23 or 24 (UC-1 and After) (Ref. Chapter 6-50-00). (14) Install the passenger compartment center aisle and left or right carpet (Ref. Chapter 25-20-01) and floorboards, forward of the main spar (Ref. Chapter 6-50-00).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (15) Install the passenger compartment left or right seats, forward of the main spar (Ref. Chapter 25-20-00). (16) Install the flight compartment carpet and center floor panel aft of the pedestal and the pedestal side panels. (17) Install the flight compartment seat (Ref. Chapter 25-10-00). (18) Install the left forward partition. (19) Perform engine run-up and adjust as necessary. Refer to the appropriate Pilot’s Operating Handbook/Airplane Flight Manual.

4. ENGINE CONTROL A. Rigging The following rigging instructions produce nominal settings of the operating parameters of the engine. Any time an engine is removed, the fuel control unit is changed, the propeller governor is replaced or any other time the adjustments of one of these units is disturbed, the engine controls rigging procedure should be performed, all or in part. These rigging procedures yield only nominal results; a ground check and final adjustment must be performed any time a rigging procedure is completed. Results of the operating parameters of each engine should be recorded and final adjustments must be made to restore the engine to its original performance parameters. The following final adjustments to the engine operating parameters are permitted to field technicians: •

1500 rpm torque per Chart 2 ± 50 foot-pounds and ± 20 foot-pounds between engines

•

Lo-idle speed of 58 to 61%

•

Hi-idle speed of 70% ± 1%

•

Max N1 of 104%, 95% with the max N1 shim tool in place

•

Max N2 of 1700 rpm.

•

Max reverse N1 of 87% ± 1%

5. POWER LEVERS A. Rigging WARNING: Misadjustment of the beta valve can cause unplanned feathering of the propeller. Resulting in a possible hazard to airplane operation and overtorque damage to the engine (Ref. Figure 208). (1) Install a pin through the rigging hole in the lower LH side of the pedestal and through the two power lever bellcranks. (2) Adjust the pedestal interconnect rods until the power levers line up in the same position. Both power levers should be against the idle stop. (3) Install a pin through the rigging hole in the cambox; a No. 41 drill bit can be used as the rig pin (Ref. Figure 205).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (4) Connect the power lever cable end to the cambox input lever. The lever may be adjusted on the input lever shaft as necessary to facilitate connecting the power lever cable to the input lever. (5) Adjust the power lever cable end as necessary to ensure that the angle of the input lever is equal to the angle of the input lever on the opposite engine. The input levers should be situated in a nearly horizontal position. To control backlash at idle settings, a maximum of two springs may be used on the power control cable (Ref. Figure 205) NOTE: The input lever must be in a 9-o’clock to 10-o’clock position, otherwise the rig pin hole will be covered up by the input lever bolt. (6) Disconnect the beta cable from the cambox to avoid damaging the carbon block in the propeller feedback ring. (7) Adjust the cambox-to-fuel-control-unit of the interconnect rod to a length of 8.25-inches as measured from the centers of the attach holes. (8) The interconnect rod should be installed between the top hole on the fuel control unit arm and the next-to-top hole on the cambox. (9) Remove the rig pins from the cambox and the pedestal power levers. (10) Place a protractor on the cambox input lever. (11) Pull the power lever back into the reverse range as far as necessary in order to get the dead-band-stop screw to move off the dead-band stop. Slowly move the power lever forward until the dead-band-stop screw hits the stop. The screw should be against the stop such that a piece of paper between the screw and the stop is held tightly and any further motion in the reverse direction will release the paper (Ref. Figure 206). (12) Measure the angle of the cambox input lever. (13) Push the power lever forward to the point where the stop screw is about to lift off the stop, but will still grip the piece of paper tightly. (14) Measure the cambox input lever angle again. The difference between the two angular measurements represents the dead-band travel and should be 11° to 12°. Adjust the dead-band-stop screw as necessary to obtain the proper dead-band travel. One full turn out of the dead-band-stop screw will widen the dead band 1.32°. (15) Check that the dead-band-stop screw first contacts the stop slightly before the pedestal power lever goes into idle and that dead-band-stop screw begins to lift off the stop as the pedestal power lever goes into reverse. (16) Place the pedestal power levers in idle. Check that the rig pin will fit into the rig pin hole. If it does not, adjust the input lever on the input lever shaft. (17) Once the amount of dead-band travel has been properly set, the dead-band travel can be repositioned between ground fine and idle by adjusting the serrated washer on the upper fuel control arm (Ref. Figure 206). A movement of one serration will result in a 0.6° change in the position of the input arm. Fine adjustments of the position of the dead band may be made by adjusting the length of the interconnect rod.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL NOTE: If it is impossible to position the dead band between idle and ground fine, it is permissible to widen the dead band by an amount not to exceed 1° forward of the idle detent or 1° aft of the ground fine detent. CAUTION: Once the beta cable is connected to the cambox lever, no further attempts to place the power levers in reverse should be made unless the engine is running. (18) Connect the beta cable to the cambox cam in the second hole from the top. NOTE: When lengthening the beta cable by adjusting the rod ends, leave at least one thread visible inside the inspection hole. (19) The beta cable should be adjusted at either the forward or aft rod ends so that the flat surface of the clevis on the beta valve plunger is flush with the forward surface of the conical cap nut on the propeller governor beta valve. This adjustment should be made with the propeller solenoids de-energized (Ref. Figure 207). (20) Adjust the airbleed link (lost motion link) until a 0.1-inch gap between the upper and lower halves is apparent. NOTE: Three complete turns of either upper or lower end adjustment of the airbleed link is equivalent to 0.1-inch.

6. CONDITION LEVERS A. Rigging (1) Place the condition lever in the LOW IDLE detent. (2) Install a rig pin through the rig pin hole in the lever stop assembly on the left side of the fuel control lever (Ref. Figure 206). (3) Adjust the condition lever cable end or the input lever position as necessary and connect the cable to the fuel lever stop assembly. (4) Remove the rig pin from the fuel lever stop assembly.

7. PROPELLER LEVERS A. Rigging (1) Place the propeller lever in FEATHER. (2) Adjust the cable end so that the feather dump valve plunger is fully compressed and bottomed-out when the cable is connected to the speed adjusting lever on the propeller governor (Ref. Figure 207). (3) Check that the bolt connecting the cable to the governor lever does not strike the governor body.

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8. PROPELLER GROUND-FINE SOLENOID A. Rigging WARNING: Misadjustment of the beta valve can cause unplanned feathering of the propeller. Resulting in a possible hazard to airplane operation and overtorque damage to the engine (Ref. Figure 208). (1) Install the solenoid in the supporting bracket with the aft surface of the solenoid flush with the aft portion of the support bracket. (2) Connect the solenoid arm to the propeller reversing lever (Ref. Figure 207). (3) Position the solenoid bracket on the beta cable housing so that the distance between the forward surface of the solenoid and the center of the clevis pin through the end of the solenoid plunger is 0.5-inch. (4) Assure that all hardware is tight and that the safety wires are installed at the beta cable end as appropriate.

9. GROUND-FINE A. Check (1) Ground fine check shall be done with full fuel (for UA and UB series airplanes) or full mains (for UC series) with the airplane setting on a flat surface. If any wind exists, the check shall be done in a crosswind condition. CAUTION: To avoid overheating the cabin windows, do not run the engines in the feathered position while in a quartering to crosswind condition. (2) Check that N1 remains at idle N1 setting and N2 increases slightly, indicating that the propeller pitch has shifted to a lower setting (N1 should not begin to increase until power levers are lifted and moved aft of the ground fine stop). Advance power levers forward, then pull back to ground idle position. (3) Without brakes, advance the condition levers from low idle to high idle. Check that N2 increases and that the airplane does not move forward or backwards. (4) If the airplane has forward movement, reposition the beta cable on the cambox to the next higher hole (Ref. Figure 205). Adjust the clevis on the aft end of the beta cable so that previous settings are not changed. If the beta cable is in the top hole, no further adjustments are possible. If an adjustment is made, return to Step (2). (5) If the airplane has movement aft and the beta cable has just been moved up one hole, no more adjustments are possible and the clevis shall be positioned to the previous hole down. If the airplane has aft movement and the clevis is in the bottom hole, return to the ground fine rigging and power lever rigging position and recheck. If an adjustment has been made, return to Step (2).

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Figure 203 Engine Controls (Engine End)

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Figure 204 Engine Controls (Pedestal End)

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Figure 205 Cambox Assembly

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Figure 206 Fuel Control Unit

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Figure 207 Propeller Governor Assembly

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10. ENGINE FINAL ADJUSTMENTS Prior to beginning any final adjustment procedures to the engine controls, a ground performance check must be performed. The ground performance work sheet should be used to record engine performance data. Refer to Table 201. The worksheet may be retained following final engine controls adjustments for the purpose of documentation of proper engine performance and controls rigging. WARNING: Never make adjustments or perform any of these maintenance procedures while the engine is running.

A. Ground Performance Check Procedure When performing this procedure, it should be noted that at least two of the performance parameters will be outside the expected limits when a fault is in the engine. If only one performance parameter is suspect, it can be assumed that the problem is in the indicating system for that parameter and a check of that system must be performed. (1) Install the Bauer Howden Inc. max N1 shim tool (WT107901) on the condition lever shaft stop tab so the LH side of the tool is flush with the LH side of the stop tab (Ref. Figure 206). NOTE: If the max N1 shim tool is not available, N1 will have to be checked during the test flight. (2) Verify that the feather valve plungers on the propeller GOVERNORS are fully compressed when the propeller controls are in the FEATHER position. (3) Verify that the fuel cutoff is functioning properly (Ref. FUEL CUTOFF CHECK). (4) Start the engines and allow the oil temperature to increase well into the operating range. (5) Check the dead-band position. N1 should remain constant between the idle and ground fine detents; however, N1 should rise as the power levers are moved forward from IDLE and when moved aft from GND FINE. (6) Idle speed should be 58 to 61%; if not, discontinue this procedure and go to LOW IDLE ADJUSTMENT. (7) Place power levers in idle, condition levers at high idle and note prop N2 rpm. Hold prop test switch in low pitch position, note L&R prop rpm decrease of approximately 200 rpm. Ensure rpm has stabilized. Release prop test switch, note prop rpm returns to original noted value. (8) With the power lever in the idle detent, place the propeller lever in high rpm and the condition lever in high idle. Place the propeller test switch in the low pitch position. Advance the power lever to 1500 rpm N2 record the engine 1500 rpm torque. (9) A setting of 1500 rpm torque should correspond to the value obtained from the 1500 rpm torque graph (Ref. Chart 2 Figure 209) ± 50 foot-pounds, and torque readings between engines should agree within 20 foot-pounds. Refer to ENGINE TORQUE ADJUSTMENT. (10) Place the propeller test switch in the OFF position and record the 1500 rpm torque.

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(11) This 1500 rpm torque should correspond to the values (+ 0 to 150 foot-pounds) on the graph. Refer to Chart 3, in Figure 210. NOTE: In the event the torque is out of tolerance, check for binding at the beta cable and solenoid mechanism, then adjust the solenoid aft in the bracket to lower torque or forward in the bracket to increase torque. If adjusting the solenoid does not result in an acceptable torque reading, refer to the propeller manual for adjusting the propeller low pitch stops, then return to this procedure. (12) With the power levers in the idle detent, advance the condition levers to the HI-IDLE position; this should yield an N1 of 69 to 71%; if not, discontinue this procedure, and adjust the high idle speed. Refer to HIGH IDLE ADJUSTMENT. (13) Check for power lever mismatch by moving the power levers forward and aft while observing the N1 responses; should a mismatch exist, refer to POWER LEVER ALIGNMENT. (14) Advance the power levers to the maximum forward position. Record the N1 readings. (15) Retard the power levers and move the levers into the full-reverse range. Record the N1 readings. (16) Maximum forward and reverse N1 can now be adjusted if no power lever mismatch exits. Maximum forward N1 is 95% with the max N1 shim tool in place; max reverse N1 is 88%. Example: Pressure Altitude A = 1350 ft. Outside Air Temperature B = 27° C (a) Draw a line horizontally from a point on the left side of the graph corresponding to the current pressure altitude, point A. (b) Draw a line upward from a point on the bottom of the graph corresponding to the current IOAT, point B. (c) Adjust the engine 1500 rpm torque to the torque value where the two lines from points A and B intersect, point C. (17) Advance the power levers until maximum propeller rpm is reached. Propeller rpm should advance to 1700 rpm if the propeller governor is properly adjusted; otherwise refer to PROPELLER GOVERNOR ADJUSTMENT. (18) Shut down the engines and remove the max N1 shim tool. (19) Check that all cotter pins and safety wiring are correctly installed and that all clevis end nuts are properly tightened and safetied. (20) Perform the ENGINE OPERATING PARAMETERS CHECK.

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B. Engine Operating Parameters Check NOTE: Engine parameter check data is used to evaluate the effects of component replacement, inaccuracies in the engine instruments or progressive hot section deterioration; however, such data should never be used as the sole criteria for determining the airworthiness of an engine. Refer to the MINIMUM TAKE-OFF POWER Chart and procedures in the appropriate Pilot’s Operating Handbook to determine if the engines are producing sufficient power for airworthy operation. The graphs for engine operating parameter checks are presented as torque versus ambient temperature and pressure so that engine parameters can be checked over a wide range of conditions without overtorquing the reduction gearbox. Periodic engine checks must be carried out and changes in engine parameters noted. All forms of engine deterioration will be accompanied by an increase in interstage turbine temperature and fuel flow at a given power. Compressor deterioration will, in most cases, be due to dirt deposits and will effect an increase in gas generator speed to obtain a given power setting. This form of deterioration can be remedied by compressor washing as described in the engine maintenance manual. Hot-section deterioration will cause a decrease in the gas generator speed, an increase in fuel flow and an increase in ITT for a given power setting. NOTE: Prior to beginning the following checks, the engine cowling must be in place in order to ensure consistency of engine check parameters. (1) Determine the field barometric pressure. NOTE: This pressure may be obtained by setting the altimeter to zero and reading the pressure from the Kollsman window or by contacting the local Flight Service for an uncorrected pressure reading. Do not use the sea level pressure normally reported by the tower. (2) Record the outside air temperature in degrees Celsius. (3) Obtain the target torque reading from Graph 1 of Chart 4, in Figure 211 by plotting temperature against field barometric pressure. (4) Start the engines and allow the oil temperature to increase to operating temperature. (5) Position the airplane crosswind, 90° to the wind direction, to eliminate variation in parameters due to changing wind velocity. (6) Turn off the generator, bleed air and air conditioning on the engine to be checked. (7) Check that the propeller levers are in high rpm. (8) Bring the power levers forward to establish a torque indication equal to the target torque value from Graph 1 of Chart 4, in Figure 211. (9) Check that engine propeller rpm is governing at 1700 rpm. (10) Record the N1 readings, ITT indications and the fuel flows from the engine being checked. (11) Shut down the engines. (12) Determine the reference fuel flow from Graph 2 of Chart 4, in Figure 211 by plotting the field barometric pressure against the outside air temperature.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (13) Determine the reference N1 speed from Graph 3 of Chart 4, in Figure 211 based on outside air temperature. (14) Find the reference ITT value by plotting the outside air temperature to the ITT curve. The graphs in Chart 4, in Figure 211 are based on a typical new installation. Minor differences between engines and installations may produce results that are slightly different from the reference values represented by the graphs; therefore, should the engine meet the minimum takeoff power criteria as stated in the Pilot’s Operating Handbook, large variations of fuel flow and ITT from the graph reference values may be indicative of faults within the engine’s indicating systems. Should an engine parameter begin to drift progressively up or down during subsequent checks, a trend toward engine deterioration may be assumed. A minimum takeoff power check should be performed before the engine is declared nonairworthy. The real value of the graphs on Chart 4, in Figure 211 is in giving the technician an approximation of where the engine parameters should be for the typical new installation. Detailed record keeping of engine performance parameters is all-important in detecting early trends of change in performance. For that purpose, a good system of engine condition trend monitoring should be employed by maintenance personnel so that preventive maintenance measures may be accomplished well in advance of the development of more serious conditions.

C. Estimated Field Barometric Pressure Calculation NOTE: This is an alternate method of obtaining field barometric pressure and should be used only in instances where the Kollsman window is off scale. (1) Set the altimeter to field elevation. Record the elevation and the barometric pressure from the Kollsman window. (a) Example: 1820 ft MSL at 29.84 inches Hg (2) Use 1.0 inch Hg per 1,000 ft MSL of altitude to determine the estimated field barometric pressure. (a) Example: 1820 ft MSL/1,000=1.82 inches Hg (3) Subtract the 1.82 inches Hg from the barometric pressure recorded in Step (1) (a). (a) Example: 29.84 inches Hg - 1.82 inches Hg=28.02 inches Estimated Field Barometric Pressure (4) Apply 28.02 inches Hg to the proper Charts. Refer to ENGINE OPERATING PARAMETERS CHECK.

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MISADJUSTMENT OF THE BETA VALVE CAN CAUSE UNPLANNED FEATHERING OF THE PROPELLER, RESULTING IN A POSSIBLE HAZARD TO AIRPLANE OPERATION AND OVERTORQUE DAMAGE TO THE ENGINE. BETA VALVE CONICAL CAP NUT

BETA VALVE CONICAL CAP NUT

VALVE POSITION

CLEVIS

CLEVIS

FLAT SURFACE OF CLEVIS AFT SURFACE CONICAL CAP NUT VALVE SHOWN IN FULL AFT (UNADJUSTED) POSITION FOR CLARITY

FLAT SURFACE OF CLEVIS FLUSH WITH AFT SURFACE OF CONICAL CAP NUT VALVE SHOWN IN ADJUSTED POSITION UC76B 045953AA.AI

BETA ARM NOT SHOWN FOR CLARITY

Figure 208 Beta Valve Adjustment

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Table 201 Ground Performance Worksheet Airplane S/N: ______________

Engine S/N: LH _____________ RH _____________

Technician: ______________

Date: ______________

OAT: _____________

Pressure Altitude: _______________

Field Elevation: _______________

Fuel Cutoff Check: LH ____________ RH ______________ Engine Start Comments: LH __________________________________________________________________ RH __________________________________________________________________ Engine Oil Temp: LH ______________ RH ______________ GROUND PERFORMANCE CHECKS Idle N1: (Target 59%)

Before Adjustment LH ________ RH ________

After Adjustment LH ________ RH ________

Place Propeller Test Switch in LOW PITCH Position 1,500 RPM Torque from Chart 2: (± 50 AND ± 20 Between Engines)

Before Adjustment LH ________ RH ________

After Adjustment LH ________ RH ________

Place Propeller Test Switch in OFF Position 1,500 RPM Torque from Chart 3: (+0, -150)

Before Adjustment LH ________ RH ________

After Adjustment LH ________ RH ________

Hi-Idle N1: (Target 69 to 71%)

Before Adjustment LH ________ RH ________

After Adjustment LH ________ RH ________

MAX FWD N1: (Target 95%)

Before Adjustment LH ________ RH ________

After Adjustment LH ________ RH ________

MAX REV N1: (Target 88%)

Before Adjustment LH ________ RH ________

After Adjustment LH ________ RH ________

MAX N2: (Target 1,700 RPM)*

Before Adjustment LH ________ RH ________

After AdjustmenT LH ________ RH ________

ENGINE OPERATING PARAMETER CHECK (Ref. Chart 4) Field Barometric Pressure: ________ OAT: ________ Torque from Graph 1: _________

Indicated LH ________ RH ________

Propeller Speed: (Target 1,700 RPM)*

Indicated LH ________ RH ________

Fuel Flow from Graph 2:_________

Indicated LH ________ RH ________

Engine N1 from Graph 3:_________ Indicated LH ________ RH ________ ITT from Graph 3: _________

Indicated LH ________ RH ________

Engine Start Comments: LH __________________________________________________________________ RH __________________________________________________________________ * For alternate propeller settings, see PROPELLER SETTINGS in Chapter 61-00-00.

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Figure 209 Chart 2 Torque Graph -1500 RPM

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Figure 210 CHART 3 - TORQUE GRAPH - 1500 RPM WITH PROPELLER TEST SWITCH OFF

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3600

FIELD BAROMETRIC PRESSURE (IN HG)

31

3500 29.92

3300

28

3200

27

3100

26

3000 2900

25

2800 24

2700 23

ENGINE FUEL FLOW (LB/HR)

2600

22

700

FIELD BAROMETRIC PRESSURE (IN HG)

ENGINE TORQUE (FT. LBS)

3400

29

2500

31 29.92

2400

28

600

26 24 23 22

500

400

104 102 100 98

820

96

800

94

780 INTERTURBINE

760 740 720 700 INSTALLATION ASSUMPTIONS

680

NO POWER EXTRACTION (ENVIR OFF) (GEN OFF)

660 640 620

-20

-10

0

10

20

30

40

50

60 UE76B 050428AA.AI

Figure 211 CHART 4 PT6A-65B OPERATING PARAMETER CHECK GRAPHS

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11. ENGINE POWER This procedure is to help the mechanic determine if the engine is developing sufficient power to be considered airworthy for safe operation at higher elevation airports. An engine that has deteriorated to the condition that it would not develop sufficient takeoff power in hot ambient conditions at 9000 feet of elevation may be completely sufficient in cooler ambient conditions at a lower elevation. Many factors, including the engine controls being out of rig and malfunctioning components, could make a good engine appear unairworthy. When using Chart 5 and Figure 212 to make an engine power check, refer to Chart 6 and the graph in Figure 213 for an example of how the conditions defined in the Chart are plotted on the graph to determine when engine ITT is within acceptable limits for safe operation at higher elevation airports.

A. Check (1) Using the STATIC TAKE-OFF POWER AT 1700 RPM Chart in the Pilot’s Operating Handbook and FAA Approved Airplane Flight Manual, determine the torque at which to set the engine according to the outside air temperature and pressure altitude. (2) Start the engine according to the Pilot’s Operating Handbook and FAA Approved Airplane Flight Manual. (3) Set the engine torque as determined in Step (1). (4) Allow the engine to stabilize for two minutes. (5) Record the torque, ITT, OAT and pressure altitude (PA) in Chart 5, Condition 1. (6) Reduce the engine torque by 500 foot-pounds and allow the engine to stabilize for two minutes. (7) Record the torque, ITT, OAT and PA in Chart 5, Condition 2. (8) Shutdown the engine according to the Pilot’s Operating Handbook and FAA Approved Airplane Flight Manual. (9) Using a copy of Figure 212, plot the data from Step (5) as indicated in the following Steps: (a) At the lower left corner of the Chart draw a vertical line up from the recorded torque to the recorded OAT. (b) From the recorded OAT draw a horizontal line to the right to the recorded PA. (c) From the PA draw a vertical line up. (d) At the upper left corner locate the recorded ITT and draw a vertical line down to the recorded OAT. (e) From the recorded OAT draw a horizontal line to the right. (f) The intersecting point of the lines drawn in Substeps (c) and (e) is point A. (10) Using the same copy of Figure 212, plot the data from Step (7) as indicated in the following Substeps: (a) At the lower left corner of the Figure 212, draw a vertical line up from the recorded torque to the recorded OAT.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (b) From the recorded OAT draw a horizontal line to the right to the recorded PA. (c) From the PA draw a vertical line up. (d) At the upper left corner locate the recorded ITT and draw a vertical line down to the recorded OAT. (e) From the recorded OAT draw a horizontal line to the right. (f) The intersecting point of the lines drawn in Substeps (c) and (e) is point B. (11) Draw a line through points A and B. (12) Using the STATIC TAKEOFF POWER AT 1700 RPM Chart in the Pilot’s Operating Handbook and FAA Approved Airplane Flight Manual, determine the takeoff torque setting required at the expected OAT and PA. (13) Using the same copy of Figure 212, enter the torque determined in Step (12) at the lower left corner of the Chart and draw a vertical line up to the expected OAT used in Step (12). (14) From the expected OAT point draw a horizontal line to the right until it intersects the expected PA. (15) From the expected PA point draw a vertical line up to the line drawn in Step (11) and call the intersection of the two lines point C. (16) From point C draw a horizontal line to the left to the expected OAT used in Step (12). (17) From the expected OAT point draw a vertical line up to the ITT scale at the upper left corner of the Chart. (18) If the ITT determined in Step (17) is less than the redline (820°C) for the engine, the engine is acceptable for the anticipated flight.

12. FUEL CUTOFF A. Check (1) Disconnect the engine fuel line at the fuel flow divider. (2) Place a container under the fuel line to catch any fuel that may be pumped out of the line during this check. (3) Motor the engine by placing the start control switch in the STARTER ONLY position. (4) Move the condition lever forward and confirm that fuel flow is present. (5) Place the condition lever in CUTOFF and confirm that the fuel flow has stopped. (6) Reconnect the fuel line to the fuel flow divider.

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13. LOW IDLE A. Adjustment (1) Loosen the fuel control unit lever clamping screw (Ref. Figure 206). Do not confuse the nut in the center of the fuel control unit with the clamping screw. (2) Loosen one of the two idle adjustment screws and tighten the other to obtain the proper idle adjustment setting (Ref. Figure 206). One flat on the allen screw will change the idle 4%. Loosen the top screw to adjust the idle up or loosen the bottom screw to adjust the idle down. (3) Tighten the fuel control unit lever clamping screw and safety wire the clamping screw and the idle adjust screws. (4) Any time low idle adjustments are changed, the maximum forward and reverse N1 schedule will shift up or down with idle speed and must be set again. (5) Go to the GROUND PERFORMANCE CHECK PROCEDURE, Step (7).

14. ENGINE TORQUE A. Adjustment (1) Shorten the beta cable to lower idle torque by adjusting the beta cable end at the cambox. Idle torque can be increased by lengthening the beta cable. Should insufficient threads remain in the clevis of the beta cable to allow lengthening of the cable, the adjustment should be made at the forward end of the cable. (2) Tighten and safety wire the jam nuts. (3) Go to the GROUND PERFORMANCE CHECK PROCEDURE, Step (8).

15. HIGH IDLE There are two possible procedures for adjusting the high idle speed of the engine. The first procedure should always be the procedure of choice unless the condition levers are split, not properly aligned together; then the second procedure would be the procedure of choice.

A. Adjustment (Condition Levers Not Split) (1) Locate the high idle roller assembly on the main input shaft of the fuel control unit (Ref. Figure 206). (2) Adjusting the roller out will increase high idle speed 1.5% per one complete turn of the adjusting nuts. Adjusting the roller in will decrease high idle speed 1.5% per one complete turn of the adjusting nuts. CAUTION: Do not set the roller so far out that the dead-band stop screw fails to ride against the dead-band stop. Should this happen, reset the high idle stop screw to allow more rotation of the condition lever shaft, then readjust the high idle roller. (3) Go to the GROUND PERFORMANCE CHECK PROCEDURE, Step (12).

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B. Adjustment (Condition Levers Split) (1) Locate the high idle stop adjusting screw (Ref. Figure 206). (2) Turning the screw in will increase the high idle 1.6% for each flat of the screw head. Turning the screw out will decrease the high idle speed 1.6% for each flat of the screw head. (3) Safety wire the adjusting screw after making this adjustment. (4) Go to the GROUND PERFORMANCE CHECK PROCEDURE, Step (12).

16. MAXIMUM FORWARD AND REVERSE N1 A. Adjustment (1) Locate the max forward N1 adjusting screw (Ref. Figure 206). (2) Turning this screw in will increase max N1 0.3% per flat of the screw. Turning this screw out will decrease max N1 0.3% per flat of the screw. (3) Locate the max reverse stop adjusting screw. (4) Turning this screw in will increase the max reverse N1 0.6% per flat of the screw. Turning this screw out will decrease the max reverse N1 by 0.6% per flat of the screw. (5) Safety wire the adjusting screws. (6) Go to the GROUND PERFORMANCE CHECK PROCEDURE, Step (16).

17. PROPELLER GOVERNOR A. Adjustment WARNING: Do not attempt to adjust the propeller governor while the engine is running. (1) Remove the safety wire from the propeller governor adjusting screw. (2) Turn the adjusting screw one turn for each 10 rpm of desired change. Turning the adjusting screw clockwise will effect a decrease in maximum N2. (3) After adjusting the governor, recheck the maximum propeller speed. More than one adjustment may be necessary in order to achieve the proper rpm. (4) Safety wire the adjusting screw.

18. POWER LEVERS A. Alignment (1) Place the power levers in IDLE. (2) Measure the angles of both cambox input levers. (3) Calculate the difference between the angles of the input levers by subtracting one from the other.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (4) Advance the power levers to the full-forward position and measure the angles of the cambox input levers again. (5) The difference between the angles of the input levers should be nearly the same as the difference measured when the power levers were in the idle position. (6) If the angular difference between the cambox input levers remained constant throughout the full travel of the power levers, and the power lever mismatch is minor, split the difference by adjusting the interconnect rod on each engine one-half of the amount required to realign the power levers. (7) If the angular difference between the cambox input levers did not remain constant, check the rigging of the power lever bellcranks in the pedestal. Refer to Step (1) in POWER LEVERS RIGGING. (8) Check the position of the dead band according to Step (15) in POWER LEVERS RIGGING. It may be necessary to widen the dead bands slightly in order to cover the full range between the idle and ground fine detents. (9) If the previous attempts to correct the power lever mismatch failed, the fault may be in either the fuel control unit or the engine. NOTE: When attempting to match a new engine to an old engine, it may be necessary to alter the adjustment of the rigging in the pedestal or to alter the engine-to-engine rigging of the camboxes in order to match the power levers. Fuel flows and ITT’s will not match between engines with differing amounts of use. (10) Perform the GROUND PERFORMANCE CHECK PROCEDURE.

19. CONTROL CABLE REPAIRS AND CHECKS Occasionally, should moisture be trapped within the control cable housing, freezing of the engine control could occur during operation in very cold conditions. This procedure will help the technician identify possible leaks and effect proper corrective action.

A. Equipment Required •

Vacuum Source

•

Dry Nitrogen

•

Dual Manifold and Gage Assembly

•

Snap Ring Pliers

•

Seal Kit (500-100-003, two required for each 3/16-inch cable)

•

Seal Kit (500-100-005, two required for each 1/4-inch cable)

B. Leak Check Procedure (1) Visually inspect cables for evidence of damage, such as crimps, cuts, unusually tight bends or abrasions. (2) Remove the snap ring washer and packing from the nacelle end of the suspect cable.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (3) Connect the manifold gage assembly to the control cable, and connect the vacuum source to one side of the manifold. (4) Apply vacuum (20 in. Hg.) to the cable, and close the valve which controls the vacuum side of the manifold, thereby trapping the vacuum inside the cable housing. (5) Monitor the vacuum gage; there should be no loss of vacuum pressure during a two minute period. (6) Any loss of vacuum pressure will indicate either a leaking packing at the pedestal end of the cable, or a leak in the cable housing cover. (7) Refer to the following procedure for repair of faults revealed in this procedure.

C. Repairing Control Cables NOTE: The following procedure applies to Teflon coated cables only and does not apply to stainless steel braided cables. No repairs are authorized on stainless steel braided assemblies. Chaff through the outer housing is justification for change. This procedure can be used to repair cuts or cracks in the outer teflon covering of the cable only. More extensive damage to the cable housing will necessitate replacing the cable. A repaired cable that exhibits any additional signs of leaking or freezing should be retested and purged. If freezing still persists, the cable should be replaced. (1) Visually examine the cable to determine the extent of the damage. The cable assembly must be discarded if the housing has been crushed, the binder or strand wires cut or if the binder or strand wires have been contaminated by petroleum products or dirt. (2) Clean the damaged area and the area within three inches either side of the damage, using a clean cloth which has been dampened with 100% alcohol or triethane. Dry the area with a clean dry cloth. (3) Wrap the area two inches either side of the damaged area with pressure sensitive teflon tape (0.005-inch-thick x 3/4-inch-wide extruded-type tape may be locally obtained). Wrapping should be performed so that the tape overlaps itself by one-half its width. (4) In areas subject to abrasion damage (clamp points and areas where the cable passes through bulkheads or firewalls), a second layer of teflon tape should be applied. This layer should be applied in the opposite direction of the first layer. (5) Heat shrinking teflon tubing may be applied over the teflon tape to provide additional protection. The recovered wall thickness of the heat shrinking teflon tubing should be approximately 0.010-inch. Maximum shrinkage may be obtained by using the largest size of tubing possible. The heat shrinking tubing should extend at least 0.25-inch beyond the tape wrapped area. (6) Heat the tubing above the crystalline melt point of the tubing (327 °F) with a forced air heat gun. As the tubing begins to shrink, remove the heat (20 to 30% shrinkage takes place as the tubing cools). (7) Apply teflon heat shrinking tubing to the high abrasion areas when installing new control cables. (8) After repairs have been effected, perform the following CABLE PURGING procedure.

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D. Cable Purging (1) Attach the dry nitrogen source to the gage manifold assembly. (2) Remove the snap ring, packing and washer from the pedestal end of the suspect control cable. (3) Apply 18 psig of dry nitrogen pressure to the cable for at least 15 minutes. (4) Install a new packing, washer and snap ring on the pedestal end of the cable. (5) Apply 10 psig of dry nitrogen pressure to the cable, trap the pressure in the cable, and monitor the pressure gage for a two minute period. There should be no observed pressure loss. (6) Remove the gage-manifold assembly and install a new packing, washer and snap ring on the nacelle end of the cable. (7) Connect the cable to the engine. CHART 5 POWER PREDICTION WORKSHEET Airplane S/N _________________

Engine S/N _________________

Date _________________

Condition 1

Condition 2

Expected Ambient Conditions

*Torque Ft. Lbs. ________

Torque Ft. Lbs. ________

*Torque Ft. Lbs. ________

ITT °C_________

ITT °C ________

*ITT °C ________

OAT °C ________

OAT °C ________

OAT °C ________

PA Ft. ________

PA Ft. ________

PA Ft. ________

* From Airplane Flight Manual. CHART 6 POWER PREDICTION WORKSHEET Airplane S/N XXXXXX

Engine S/N XXXXXX

Condition 1

Condition 2

Expected Ambient Conditions

*Torque Ft. Lbs. 3400

Torque Ft. Lbs. 2900

*Torque Ft. Lbs. 2870

ITT °C 793

ITT °C 753

* ITT °C 806

OAT °C 20.6

OAT °C 20.6

OAT °C 35

PA Ft. 5200

PA Ft. 5200

PA Ft. 7000

* From Airplane Flight Manual.

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Figure 212 Engine Temperature Check Chart

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This Page Intentionally Left Blank

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Figure 213 Engine Temperature Check Chart (Example)

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1. 2. 3. 4. 5. *6. 7. 8. 9. 10. 11. 12. 13. **14. 15. 16. 17. 18. 19. 20. 21.

PIN SPRING WASHER ARM ASSEMBLY WASHER NUT WASHER COTTER PIN PIN NUT WASHER WASHER PIN WASHER PIN POWER LEVER PIN COTTER PIN WASHER WASHER WASHER

A

16

2

4 1 5

3

*6

7

15 **14

8

13

C 12

DETAIL

11

9

B

B

10

21

20 19 17

* WITH KIT 129-5009

18

DETAIL

C DETAIL

** INSTALL WASHER WITH THE COATING FACING THE LEVER.

A

UC76B 100125AA.AI

Figure 214 Power Lever Assembly

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ENGINE CONTROLS POWER CONTROL MAINTENANCE PRACTICES

76-10-00 200200

1. CONDITION CONTROL CATCH GATE A. Removal (1) Set both the aileron and rudder trim tab adjusting knobs (7 and 10) so that the tab indicators (8) read 0. This will allow the adjusting knobs to be reinstalled in the proper position (Ref. Figure 202). (2) Loosen the set screw in the side of the trim adjusting knobs (7 and 10) and remove the knobs. (3) Remove the grip ring under each knob. (4) Remove the indicator dials (8). (5) Loosen the set screws and remove the friction lock knobs (3). (6) Remove the screws attaching the electroluminescent panel (2) that is installed over the POWER, PROP and CONDITION levers. (7) Position the electroluminescent panel (2) to gain access to the area where the condition control catch gate is located (Ref. Figure 201). (8) Peel back the rubber cover (1) in the area of the condition control catch gate. Two screw heads will become visible (Ref. Figure 202). (9) Remove the screws attaching the escutcheon (9). Lift the escutcheon and disconnect the wires to the electroluminescent panel (6). Remove the escutcheon. (10) Remove the screws attaching the flap switch gate (5). The gate may be allowed to hang from the flap switch lever. (11) Remove the screws attaching the flap switch cover plate (4) and remove the plate. (12) Gain access to the condition control catch gate and remove the two screws, washers and nuts attaching the catch to the pedestal. (13) Discard the old catch.

B. Installation (1) Apply lubricant (45, Table 1, Chapter 91-00-00) to the areas of the new catch gate that will rub against the condition lever. (2) Install the new catch gate with the screws, washers and nuts. Locate the catch gate at the low idle end of the lever travel (Ref. Figure 201). (3) Secure the rubber cover (1) that was peeled back during removal with adhesive (16, Table 1, Chapter 91-00-00) (Ref. Figure 202). (4) Install the flap switch cover plate (4) with screws.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (5) Install the flap switch gate (5) with screws. (6) Connect the wires to the electroluminescent panel (6) and install the escutcheon (9) with screws. (7) Install the friction lock knobs (3) and tighten the set screws. (8) Install the electroluminescent panel (2) with screws. (9) Install the indicator dials (8) so that they read 0. (10) Install the grip rings. (11) Install the trim tab adjusting knobs (7 and 10) and tighten the set screws.

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OUTBOARD

OUTBOARD

"CONDITION LEVERS" SHOWN IN LOW IDLE POSITION

"CATCH GATE" (TYPICAL) NOTE: DIFFERENT STYLES MAY BE INSTALLED - HOWEVER THEIR FUNCTION IS THE SAME.

PEDESTAL SHOWN WITH RUBBER SLOT SEAL REMOVED FOR CLARITY

UB76B 991585AA.AI

Figure 201 Condition Control Catch Gate

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1 2 3

3 4 5

10 6 7

8

8

9

1. RUBBER COVER 2. ELECTROLUMINESCENT PANEL 3. FRICTION LOCK KNOBS 4. FLAP SWITCH COVER PLATE 5. FLAP SWITCH GATE 6. ELECTROLUMINESCENT PANEL 7. RUDDER TRIM TAB ADJUSTING KNOB 8. TRIM TAB INDICATOR DIALS 9. ESCUTCHEON 10. AILERON TRIM TAB ADJUSTING KNOB UB76B 991586AA.AI

Figure 202 Condition Control Catch Gate Installation

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CHAPTER 77 - ENGINE INDICATING TABLE OF CONTENTS SUBJECT

PAGE

ENGINE INDICATING SYSTEM 77-00-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Troubleshooting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 101 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 101 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Interturbine Temperature (ITT) Indicating System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 ITT Thermocouple . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202 Loop Resistance Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202 Harness Insulation Resistance Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 203 ITT Indicator . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 203 Calibration Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 203 RPM Indicator Systems (N1 and N2) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 204 N1 and N2 Indicator System Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 204 Torque Indicating System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 205 Torque Indicator . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 205 Calibration Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 205 Fuel Flow Indicating System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 206 Fuel Flow Indicator (UC-2 thru UC-9, UC-11 thru UC-14, UC-17, UC-19, UC-20, UC-22 thru UC-26 Without SB 2275 Accomplished and Without Kit 114-9026 Installed; UA-1 and After; UB-1 thru UB-74 Without Kit 114-9026 Installed) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 206 Calibration Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 206 Fuel Flow Transmitter (UC-2 thru UC-9, UC-11 thru UC-14, UC-17, UC-19, UC-20, UC-22 thru UC-26 without SB 2275 Accomplished and without Kit 114-9026 Installed; UA-1 and After; UB-1 thru UB-74 without Kit 114-9026 Installed) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 206 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 206 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 207 Fuel Flow Transmitter (UC-1, UC-10, UC-15, UC-16, UC-18, UC-21 and UC-27 and After Without Kit 114-9026 Installed; UC-2 thru UC-9, UC-11 thru UC-14, UC-17, UC-19, UC-20 and UC-22 thru UC-26 with SB 2275 Accomplished, but Without Kit 114-9026 Installed) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 207 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 207 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 208 Fuel Flow Indicator (UA-1 and After; UB-1 and After; UC-1 thru UC-174 With Kit 114-9026 Installed) . . . . . . . . . . . . . . . . . . 209 Calibration Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 209 Fuel Flow Transmitter (UA-1 and After; UB-1 and After; UC-1 thru UC-174 With Kit 114-9026 Installed) . . . . . . . . . . . . . . . . . . 209 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 209 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 210 Running Jetcal Indicator . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 210 Checks . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 210

77-CONTENTS

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List of Effective Pages CH-SE-SU

PAGE

DATE

77-LOEP

1

May 1/10

77-CONTENTS

1

May 1/10

77-00-00

1 101 201 thru 220

Nov 1/09 Nov 1/09 May 1/10

C2

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

ENGINE INDICATING ENGINE INDICATING SYSTEM DESCRIPTION AND OPERATION

77-00-00 00

1. GENERAL Checks performed at Hawker Beechcraft facilities on engine indicating systems include the interturbine temperature indicating system (ITT), gas generator speed (N1) indicating system, torquemeter system, fuel flowmeter system, and the propeller speed (N2) indicator system. Various analyzers and equipment are available for the purpose of checking and calibrating the engine indicating systems; however, only the equipment used by Hawker Beechcraft facilities will be discussed in this Chapter. The JETCAL ANALYZER, a product of Howell Instruments, Inc. of Fort Worth, Texas, is used by Hawker Beechcraft facilities to check and calibrate the ITT indicator and N1 and N2 indicator systems. The torquemeter system is checked by a precision pressure regulator gage assembly, manufactured by the Heise Company and available through Hawker Beechcraft parts and service operations under the part number 101-000000/934. A mechanical flowrater is used in the check procedure for the fuel flowmeter system. Any precision flowrater with a scale up to at least 700 pph can be used for this check.

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ENGINE INDICATING ENGINE INDICATING SYSTEM TROUBLESHOOTING

100100

1. PROCEDURES The first indication that checks on the engine indicating systems may need to be performed would be that the airplane engine had a shift in engine performance parameters. Usually, when only one engine performance parameter is suspect, it can be assumed that a problem in the indicating system for that parameter is present and an immediate check of that indicating system must be performed. Normally, when the fault is in the engine, at least two of the performance parameters will be outside the expected limits. In either case, the engine should never be automatically condemned without a thorough investigation of the appropriate engine indicating systems. Engine rigging should never be ruled out as a causative factor when an engine failed to meet the minimum performance parameters.

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ENGINE INDICATING ENGINE INDICATING SYSTEM MAINTENANCE PRACTICES

77-00-00 200200

1. PROCEDURES The following checks on the various engine indicating systems are compatible with analyzers and equipment being used in Hawker Beechcraft facilities. Some modifications in these procedures may be necessary if different equipment is used. All pressures, resistances and voltages are standard values and will remain unaffected, regardless of the type of equipment being used. Use of these analyzers and equipment by Hawker Beechcraft facilities does not necessarily constitute an endorsement by Hawker Beechcraft Corporation, as other factors affecting the accuracy of the equipment, such as calibration, frequency of certification or misuse, must be taken into consideration before any equipment may be used to check or calibrate the engine indicating systems.

2. INTERTURBINE TEMPERATURE (ITT) INDICATING SYSTEM A series of thermal resistive temperature sensing probes, located in the interturbine gas flow between the first stage and second stage turbines, are interconnected by a harness to form a continuous loop. The combined resistance of the thermocouples, connected in parallel, and the harness is less than 8 ohms. The resistance of the ITT loop is compensated to 8 ohms ± 0.05 ohms by the resistor thermocouple compensator spool assembly located under the floor panel immediately inside the forward entry door. The compensator spool is connected electrically in series with the thermocouple loop (Ref. Figure 201). An adjustable ITT indicator, located in the instrument panel, completes the indicating system. ITT loop resistance can be adjusted to a lower resistance by removing wire from the compensator spool or increased, if excess wire has been left within the case and not separated from the spool, by winding wire back onto the spool. The compensator spool assembly contains two spools. Only one of the spools is used at a time. If wire has previously been removed from the spool and cut off, the alternate spool can be wired into the system. An adjusting screw on the back of the ITT indicator is used to calibrate the indicator to the same temperature as indicated on the JETCAL unit. The following resistance checks should be performed according to the instructions which are a part of the particular Jetcal unit being used. These checks should not be performed until at least 10 hours have elapsed since the engine was last run so that the internal engine components and thermocouples will have cooled to ambient temperature. The worksheet in Table 201 should be used to record the results of these checks and may be retained in the airplanes records for the purpose of documentation.

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Figure 201 Compensator Spool Assembly

3. ITT THERMOCOUPLE A. Loop Resistance Check (1) Place all switches on the Jetcal unit in the OFF position. (2) Connect the power cable BH499A to the power input receptacle and to the power source. (3) Connect the power interconnect cable between the two interconnect receptacles on the unit. (4) Connect the BH485A instrument cable to the instrument cable receptacle and to the resistance adapter BH823. (5) Remove the ITT indicator from the instrument panel. (6) Connect the airplane thermocouple harness leads to the terminal block on the resistance adapter observing proper polarity. (7) Place the resistance and A/C indicator check switch in the 8 ohm position. (8) Place the function select switch in the RES position and the master power switch in the ON position.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (9) Thermocouple loop resistance must be 8 ohms ± 0.05 ohms; if not, adjust the thermocouple spool as necessary to bring the loop resistance within tolerance. NOTE: Perform the ITT THERMOCOUPLE HARNESS INSULATION RESISTANCE CHECK prior to making any adjustments to the thermocouple compensator spool.

B. Harness Insulation Resistance Check If the thermocouple loop resistance is within acceptable limits, perform the following checks on the harness insulation resistance: (1) Place all switches on the Jetcal unit in the OFF position. (2) Remove the resistance adapter from the instrument cable and connect the insulation resistance adapter BH821 to the instrument cable. (3) Place the function select switch in the INSUL position. (4) Place the insulation check switch in the R X 1000 position. (5) Place the master power switch in the ON position. (6) Connect one lead of the insulation resistance adapter to one of the thermocouple harness leads and connect the other lead of the insulation resistance adapter to an airplane ground. (7) The insulation resistance meter should indicate 5,000 ohms or more. (8) Connect the insulation resistance adapter leads to the other thermocouple harness lead and to an airplane ground. (9) Resistance should read 5,000 ohms or more. If either resistance reading is not within tolerance, locate the short circuit and repair.

4. ITT INDICATOR A. Calibration Check The following check on the ITT indicator should be performed any time the indicator is suspected to be in error or during Jetcal checks of other indicating instruments. Should thermocouple harness resistances be found to be within acceptable limits, the fault will probably be found in the indicator. The indicator may be non-functioning or simply needing a recalibration adjustment due to seasoning of internal components or minor shock damage. An adjustment screw on the back of the indicator is used to align the temperature indications on the indicator with the temperature indications on the Jetcal unit. In the event the indicator has lost its linear response to changes in thermocouple resistance, no adjustment can be made to correct this condition and the indicator must be replaced. (1) Place all switches on the Jetcal unit in the OFF position. (2) Remove the insulation resistance adapter from the instrument cable. (3) Connect the BH822 ITT indicator adapter to the instrument cable. (4) Connect the ITT indicator adapter to the proper posts on the ITT indicator.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (5) Place the resistance and A/C indicator check switch in the 8-ohm position. (6) Place the function select switch in the A/C IND CHK position. (7) Place the temperature switch in the OPERATE position and turn the master power switch on. (8) Set the test temperature on the Jetcal unit with the A/C IND ADJ control. (9) The ITT indicator readings should match the temperature settings on the Jetcal unit within the tolerances listed in Table 201. (10) Adjust the ITT indicator as necessary. (11) Seal the adjusting screw with Glyptol after calibration.

5. RPM INDICATOR SYSTEMS (N1 AND N2) Gas generator speed (N1) is indicated as a percentage of the maximum allowable gas generator speed in rpm. The Jetcal readout is expressed in percentage as well. Propeller speed (N2) is indicated in rpm and will have to be converted to percentage in order to correlate to Jetcal readings. The Jetcal unit is capable of measuring rotating speeds within ± 0.1%; therefore, the Jetcal unit lends itself well to being used as a calibrating standard. The N1 and N2 indicator systems checks must be performed while the engine is running. The airplane inverter system is used to power the Jetcal unit; therefore, a power adapter (Ref. Figure 202) will have to be fabricated in the shop to allow using the AC test jack, located beneath the RH subpanel immediately adjacent to the pedestal. The N1 or N2 indicator has no adjustment and must be replaced in the event of a disagreement between the indicator and the Jetcal reading.

A. N1 and N2 Indicator System Check (1) Remove the interconnect cable from the Jetcal unit and lift the unit out of the carrier. (2) Place the Jetcal unit in a convenient location in the airplane and connect BH499A power cable to the POWER INTERCONNECT receptacle. (3) Connect the shop fabricated power adapter to the power cable and connect the unit to the black and blue AC test jacks under the RH subpanel. (4) Place all switches on the Jetcal unit in the OFF position. (5) Connect the BH485A instrument cable to the Jetcal unit. (6) Connect the rpm check adapter BH820B to the instrument cable. (7) Remove the N1 or N2 indicator from the instrument panel and connect the rpm check adapter in series between the indicator and the indicator harness. (8) Place the RPM switch in the N1 or N2 position.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (9) Start the engine and bring the generator on line. Refer to the Pilot's Operating Handbook. (10) Turn on the airplane inverter and the power switch on the Jetcal unit. (11) Compare the N1 or N2 reading on the indicator to the N1 or N2 reading from the Jetcal and record these on the worksheet in Table 201. When making N2 checks, refer to the conversion table in Table 201 for converting propeller rpm to percentage. (12) Place all switches in the OFF position before disconnecting any cables from the airplane or the Jetcal unit.

6. TORQUE INDICATING SYSTEM The torque transmitter, attached to the torque manifold located on the LH side of the reduction gearbox senses pressure inside the reduction gearbox and transmits a signal to the torque indicator (Ref. Figure 203). The torque indicator displays torque applied to the propeller shaft in foot-pound units. The torque transmitter and indicator is functionally tested by simulating gearbox pressure. Shop air pressure is applied to the torque manifold through a torque calibration unit. The torque calibration unit is a precision pressure regulator and valve assembly. The face of the instrument is calibrated in pounds per square-inch. While the Heise unit referenced in this Chapter is accuracy certified, any precision pressure regulator accurate to within ± 0.25 psi in the range of 0 psi to 100 psi may be used in calibrating the torque indicator system. The torque indicating system response to gearbox pressure is a linear one and translates to approximately 83.6 foot-pounds of torque to 1 psi of gearbox pressure. The worksheet in Table 201 should be used to record test data and may be retained in the airplanes files for the purpose of documentation. Adjustments to the torque indicating system may be accomplished by adjusting the adjustment screw which is under a plug in the case of the transmitter.

7. TORQUE INDICATOR A. Calibration Check (1) Disconnect the vent line from the rear of the torque transmitter (Ref. Figure 203). (2) Disconnect the lower AN fitting from the torque manifold and connect the torque calibration unit to the manifold. (3) Turn both valves on the torque calibration unit off and connect shop air to the unit. (4) Connect an APU to the airplane with an output of 28 vdc. (5) Turn the battery switch on and bring one of the inverters on the line. (6) Turn the pressure source valve on the torque calibration unit on. (7) Adjust the pressure valve to each of the values on the worksheet in Table 201 and record the torque readings. (8) Calibration adjustments to the torque transmitter can be made by adjusting the screw under the adjustment plug.

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8. FUEL FLOW INDICATING SYSTEM The fuel flow indicating system makes use of a fuel flow transmitter, located beneath the RH side of the engine just forward of the fuel control unit, and a precision indicator mounted in the instrument panel (Ref. Figure 204). The transmitter measures the volume of fuel actually flowing to the injectors. The volumetric measurement is converted to mass flow (pounds per hour, pph) based upon the weight of fuel compensated for temperature. As a rule of thumb; Jet A type fuel is approximately 6.74 pounds per gallon. There are no calibration adjustments provided for this system. If an indication error is found to exist after the calibration check, it is recommended that an alternate indicator, known to be accurate, be installed and the calibration check be performed again. Should the alternate indicator yield readings not within the specified tolerance, as noted in Table 201, the fault may be assumed to be in the transmitter.

9. FUEL FLOW INDICATOR (UC-2 THRU UC-9, UC-11 THRU UC-14, UC-17, UC-19, UC-20, UC-22 THRU UC-26 WITHOUT SB 2275 ACCOMPLISHED AND WITHOUT KIT 114-9026 INSTALLED; UA-1 AND AFTER; UB-1 THRU UB-74 WITHOUT KIT 114-9026 INSTALLED) A. Calibration Check (1) Connect the fuel flow check setup to the airplane (Ref. Figure 205). (2) Close the valve on the check setup. (3) Connect an APU with an output of 28 vdc to the airplane and turn the external power switch and the battery switch on. (4) Turn the appropriate fuel standby pump on. (5) Open the valve on the check setup and purge all the air from the setup. (6) Adjust the flow rate on the flowrater to each of the values listed on Table 201 and record the airplane flowmeter indication. (7) Any disagreement between the flowrater and the airplane flowmeter must not exceed +21/-25 pph.

10. FUEL FLOW TRANSMITTER (UC-2 THRU UC-9, UC-11 THRU UC-14, UC-17, UC-19, UC-20, UC-22 THRU UC-26 WITHOUT SB 2275 ACCOMPLISHED AND WITHOUT KIT 114-9026 INSTALLED; UA-1 AND AFTER; UB-1 THRU UB-74 WITHOUT KIT 114-9026 INSTALLED) A. Removal (1) Perform the REMOVING GROUND POWER procedure (Ref. Chapter 24-40-00). (2) Perform the BATTERY DISCONNECTION procedure (Ref. Chapter 24-31-00). (3) Perform the COWLING REMOVAL procedure (Ref. Chapter 71-10-00) to gain access to the fuel flow transmitter on the RH side of the engine, just forward of the fuel control unit. (4) Disconnect the electrical connector from the fuel flow transmitter (Ref. Figure 204).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL CAUTION: Do not remove the cylinder portion of the transmitter from the square housing as this will disturb the calibration of the transmitter. (5) Cut and remove the safety wire from the fuel flow transmitter inlet and outlet fuel line fittings, disconnect the fittings and remove the transmitter. Cap the open fuel lines to prevent system contamination.

B. Installation (1) Perform the REMOVING GROUND POWER procedure (Ref. Chapter 24-40-00). (2) Perform the BATTERY DISCONNECTION procedure (Ref. Chapter 24-31-00). (3) Remove the protective caps from the inlet and outlet fuel line fittings, position the fuel flow transmitter and tighten the fuel line fittings. Secure with safety wire (Ref. Figure 204). (4) Connect the electrical connector to the top of the transmitter. (5) Perform the BATTERY CONNECTION procedure (Ref. Chapter 24-31-00). (6) Place the BATT switch and the applicable STANDBY PUMP switch to the ON position. (7) Observe the appropriate FUEL PRESS LO annunciator light extinguishes. (8) Check the fuel flow transmitter for leaks. (9) If no leaks are observed, place the STANDBY PUMP and BATT switches to the OFF position. (10) Perform the COWLING INSTALLATION procedure (Ref. Chapter 71-10-00).

11. FUEL FLOW TRANSMITTER (UC-1, UC-10, UC-15, UC-16, UC-18, UC-21 AND UC-27 AND AFTER WITHOUT KIT 114-9026 INSTALLED; UC-2 THRU UC-9, UC-11 THRU UC-14, UC-17, UC-19, UC-20 AND UC-22 THRU UC-26 WITH SB 2275 ACCOMPLISHED, BUT WITHOUT KIT 114-9026 INSTALLED) A. Removal (1) Perform the REMOVING GROUND POWER procedure (Ref. Chapter 24-40-00). (2) Perform the BATTERY DISCONNECTION procedure (Ref. Chapter 24-31-00). (3) Perform the COWLING REMOVAL procedure (Ref. Chapter 71-10-00) to gain access to the fuel flow transmitter on the RH side of the engine, just forward of the fuel control unit. NOTE: The magnetic shield assembly consist of an inboard and outboard half. The outboard half may be identified by the flow direction decal affixed to it. If the decal is not present, mark the outboard half to facilitate installation. (4) Remove magnetic shields halves (6 and 7) attaching screws (5) and disconnect bonding jumper (4) from the magnetic shields (Ref. Figure 206). (5) Disconnect the bonding jumper (4) from the engine flange (1).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (6) Remove nut (14), screw (11), washer (12), spacer (13) and clamp (outboard half) (2) from the fuel flow transmitter fuel line and remove the outboard shield half (6). (7) Remove nut (14), screw (11), washer (12), spacer (13) and clamp (inboard half) (9) from the fuel flow transmitter fuel line and remove the inboard shield half (7). (8) Disconnect electrical connector (10) from the fuel flow transmitter (8). (9) Cut and remove the safety wire from the fuel flow transmitter inlet and outlet fuel line fittings, disconnect the fittings and remove the transmitter. Cap the open fuel lines to prevent system contamination.

B. Installation (1) Perform the REMOVING GROUND POWER procedure (Ref. Chapter 24-40-00). (2) Perform the BATTERY DISCONNECTION procedure (Ref. Chapter 24-31-00). (3) Remove the protective caps from the fuel flow transmitter inlet and outlet fuel lines, position the fuel flow transmitter (8) and tighten the fuel line fittings. Secure with safety wire (Ref. Figure 206). NOTE: The magnetic shield assembly consist of an inboard and outboard half. The outboard half may be identified by the flow direction decal affixed to it. The inboard half must be installed first. (4) Position the inboard magnetic shield (7) half aft of the fuel flow transmitter (8) and install the clamp (inboard half) (9), spacer (13), washer (12), screw (11) and nut (14). Do not tighten the clamp at this time. (5) Position the outboard magnetic shield (6) half forward of the fuel flow transmitter and install the clamp (outboard half) (2), spacer (13), washer (12) screw (11) and nut (14). Do not tighten the clamp at this time. (6) Connect bonding jumper (4) to magnetic shields halves (6 and 7) with attaching screws (5). Do not tighten the screws at this time. (7) Tighten the clamps installed in Steps (4) and (5). (8) Tighten the magnetic shield attachment screws installed in Step (6). (9) Attach bonding jumper (4) to engine flange (1) using engine flange nut (3) located at the 3 o’clock position; torque 85 to 95 inch-pounds. (10) Connect the electrical connector (10) to the fuel flow transmitter (8). (11) Perform the BATTERY CONNECTION procedure (Ref. Chapter 24-31-00). (12) Place the BATT and STANDBY PUMP switches to the ON position. (13) Observe the appropriate FUEL PRESS LO annunciator light extinguishes. (14) Check the fuel flow transmitter for leaks. (15) If no leaks are observed, place the STANDBY PUMP and BATT switches to the OFF position.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (16) Perform the COWLING INSTALLATION procedure (Ref. Chapter 71-10-00).

12. FUEL FLOW INDICATOR (UA-1 AND AFTER; UB-1 AND AFTER; UC-1 THRU UC-174 WITH KIT 114-9026 INSTALLED) A. Calibration Check (1) Perform the COWLING REMOVAL procedure (Ref. Chapter 71-10-00). (2) Disconnect the fuel flow transmitter outlet (7) line from the fuel flow transmitter (6) and cap the line (Ref. Figure 207). (3) Connect a test flowmeter (4) to the outlet port of the fuel flow transmitter. (4) Connect a flex hose (2) discharge line from the test flowmeter output that will return the fuel to the wing fuel tank (1) or a clean container. (5) Perform the APPLYING GROUND POWER procedure (Ref. Chapter 24-40-00). (6) Place the appropriate STANDBY PUMP switch to the ON position. (7) Open the needle valve (3) on flowmeter (4) and let the fuel flow at full flow rate until no bubbles are present in the flowmeter glass before starting the test (8) Adjust the flow rate on the flowmeter to each of the values listed on Table 201 and record the airplane fuel flow indicator indication. (9) Any disagreement between the flowmeter and the airplane fuel flow indicator must not exceed ± 2%. (10) Place the appropriate STANDBY PUMP switch to the OFF position. (11) Perform the REMOVING GROUND POWER procedure (Ref. Chapter 24-40-00). (12) Remove flowmeter (4) and connect the fuel flow transmitter outlet (7) line to the fuel flow transmitter (6). Tighten the fuel line fittings and secure with safety wire. (13) Perform Steps (1) thru (12) for the opposite fuel flow indicator. (14) Perform the COWLING INSTALLATION procedure (Ref. Chapter 71-10-00).

13. FUEL FLOW TRANSMITTER (UA-1 AND AFTER; UB-1 AND AFTER; UC-1 THRU UC-174 WITH KIT 114-9026 INSTALLED) A. Removal (1) Perform the REMOVING GROUND POWER procedure (Ref. Chapter 24-40-00). (2) Perform the BATTERY DISCONNECTION procedure (Ref. Chapter 24-31-00). (3) Perform the COWLING REMOVAL procedure (Ref. Chapter 71-10-00) to gain access to the fuel flow transmitter on the RH side of the engine, just forward of the fuel control unit.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (4) Disconnect electrical connector (2) from the fuel flow transmitter (1) (Ref. Figure 208). NOTE: Note the position of the flow direction arrow for installation. (5) Cut and remove the safety wire (3) from the fuel flow transmitter inlet and outlet fuel line fittings (4), disconnect the fittings and remove the transmitter. Cap the open fuel lines to prevent system contamination.

B. Installation (1) Perform the REMOVING GROUND POWER procedure (Ref. Chapter 24-40-00). (2) Perform the BATTERY DISCONNECTION procedure (Ref. Chapter 24-31-00). (3) Remove the protective caps from the inlet and outlet fuel line fittings (4), position the fuel flow transmitter (1) and tighten the fuel line fittings. Secure with safety wire (Ref. Figure 208). (4) Connect the electrical connector (2) to the fuel flow transmitter. (5) Perform the BATTERY CONNECTION procedure (Ref. Chapter 24-31-00). (6) Place the BATT switch and the applicable STANDBY PUMP switch to the ON position. (7) Observe the appropriate FUEL PRESS LO annunciator light extinguishes. (8) Check the fuel flow transmitter (1) for leaks. (9) If no leaks are observed, place the STANDBY PUMP and BATT switches to the OFF position. (10) Perform the COWLING INSTALLATION procedure (Ref. Chapter 71-10-00).

14. RUNNING JETCAL INDICATOR A. Checks Occasionally, vibration, flight loads or varying flight attitudes may induce faults of an intermittent nature which may not show up in routine indicator system checks. It may be desirable, at times to perform Jetcal checks while the airplane engines are running or while the airplane is in flight. Performance graphs in the appropriate flight manual may be used to set up engine power as a reference for in-flight checks. Operating parameter graphs in Chapter 76, Chart 4 of this manual, should be used for reference when performing static ground checks. The following procedure provides instructions for connecting the Jetcal unit to the airplane indicating systems for either static ground checks or the in-flight checks. (1) Place all switches on the Jetcal unit in the OFF position. (2) Remove the power interconnect cable, lift the Jetcal unit out of the carrier and place the unit in a convenient position in the airplane. (3) Connect the BH499A power cable to the POWER INTERCONNECT receptacle. (4) Connect the shop fabricated power adapter to the power cable and connect the unit to the black and blue AC test jacks under the RH subpanel (Ref. Figure 202).

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (5) Connect the BH485A instrument cable to the RPM INPUT INSTRUMENT CABLE receptacle on the Jetcal unit. (6) Connect the BH15572 rpm cable to the instrument cable. (7) Connect two rpm check adapters BH820B to the rpm cable. (8) Connect the two rpm check adapters in series with the N1 and N2 INDICATORS. A support bracket may be shop fabricated to support the indicators in the instrument panel (Ref. Figure 209). (9) Connect the BH450 check cable to the CHECK CABLE receptacle on the Jetcal unit. (10) Connect the check cable to the top receptacle of the BH134 switch box. (11) Remove the ITT indicator from the instrument panel and connect the indicator to the switch box. (12) Connect the airplanes ITT harness to the rear receptacle of the switch box and slip the switch box into the instrument panel ITT opening. (13) With the engines running and the battery and inverter on, turn on the power switch on the Jetcal unit. (14) Place the TEMPERATURE switch on the Jetcal unit in the OPERATE position. (15) Place the RPM switch in either the N1 or N2 position. (16) Refer to 76-00-00, ENGINE OPERATING PARAMETER CHECK procedure for instructions in setting up engine power. (17) A switch on the ITT switch box is used to select readings from either the indicator or the Jetcal unit. (18) Rpm readings occur simultaneously on the indicator and the Jetcal unit; however, it is necessary to select either N1 or N2 readings with the RPM switch on the Jetcal unit.

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Figure 202 Power Adapter

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Figure 203 Torque Manifold-Transmitter

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Figure 204 Fuel Flow Transmitter Without SB 2275 or Kit No. 114-9026 Installed

Figure 205 Fuel Flow Check Setup Without Kit No. 114-9026 Installed Page 214 May 1/10

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

1. ENGINE FLANGE 2. CLAMP (OUTBOARD HALF) 3. ENGINE FLANGE NUT 4. BONDING JUMPER 5. SHIELD ATTACHMENT SCREWS 6. SHIELD (OUTBOARD HALF) 7. SHIELD (INBOARD HALF)

8. FUEL FLOW TRANSMITTER 9. CLAMP (INBOARD HALF) 10. CONNECTOR 11. SCREW 12. WASHER 13. SPACER 14. NUT

1

A 3

8

5

4 10

11

7

12 13

2

14 14 9

13 12 11

DETAIL

A

6 UC77B 062080AA.AI

Figure 206 Fuel Flow Transmitter Installation With SB 2275 Installed

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1. FUEL TANK 2. FLEX HOSE 3. NEEDLE VALVE 4. FLOW METER 5. FLEX HOSE, FUEL FLOW TRANSMITTER OUTLET 6. FUEL FLOW TRANSMITTER 7. FUEL FLOW TRANSMITTER OUTLET

2

4

FLOW DIRECTION 3 1

5

6

FWD

7

UC77B 062084AA.AI

Figure 207 Fuel Flow Check Setup With Kit No. 114-9026 Installed Page 216 May 1/10

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

4 1

3

4 2 FWD

1. FUEL FLOW TRANSMITTER AND MAGNETIC SHIELD 2. ELECTRICAL CONNECTOR 3. SAFETY WIRE 4. FUEL LINE FITTINGS

RH ENGINE VIEW LOOKING INBOARD

UC77B 062083AA.AI

Figure 208 Fuel Flow Transmitter With Kit No. 114-9026 Installed

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Figure 209 Indicator Support Bracket

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Table 201 ENGINE INDICATING CHECK WORKSHEET DATE: __________ AIRPLANE S/N _____________

ENGINE S/N LH ________ RH ________

TECHNICIAN _____________________

COMMENTS: ______________________________

ITT INDICATING SYSTEM LOOP RESISTANCE: LH _____ RH _____ 8 ohms ± 0.05

HARNESS INSULATION: LH _____ RH _____ Greater than 5,000 ohms

ITT INDICATOR LEFT S/N RIGHT S/N JETCAL°C TOL

LH ITT

RH ITT

600 ± 25

________

________

800 ± 5

________

________

700 ± 15

________

________

900 ± 15

________

________

1000 ± 25

________

________

JETCAL°C TOL

LH ITT

RH ITT

59 ± 1

________

________

73 ± 1

________

________

87 ± 1

________

________

98 ± 1

________

________

JETCAL % TOL

LH N2 RPM

LH N2 %

RH N2 RPM

RH N2 %

77.3 ± 1

________

______

________

______

68.2 ± 1

________

______

________

______

59.1 ± 1

________

______

________

______

59.1 ± 1

________

______

________

______

RPM INDICATING SYSTEM N1 INDICATOR

N2 INDICATOR

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Table 201 ENGINE INDICATING CHECK WORKSHEET (Continued) TEST PRESSURE

TORQUE

TOL

LH IND TORQUE

RH IND TORQUE

10

836

735 to 935

_____

_____

20

1673

1605 to 1745

_____

_____

30

2509

2450 to 2570

_____

_____

40

3345

3310 to 3375

_____

_____

43.34

3625

3600 to 3650

_____

_____

50

4182

4130 to 4230

_____

_____

59.78

5000

4735 to 5065

_____

_____

FUEL FLOW INDICATING SYSTEM FLOWRATOR PPH

TOL

LH IND

RH IND

300

+ 21 - 25

________

________

400

+ 21 - 25

________

________

500

+ 21 - 25

________

________

600

+ 21 - 25

________

________

700

+ 21 - 25

________

________

800

+ 21 - 25

________

________

PROPELLER RPM CONVERSION TABLE PROP RPM

JETCAL %

1700

77.3

1650

75.0

1600

72.7

1550

70.5

1500

68.2

1450

65.9

1400

63.5

1350

61.4

1300

59.1

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CHAPTER 78 - EXHAUST TABLE OF CONTENTS SUBJECT

PAGE

EXHAUST SYSTEM 78-00-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Exhaust Stack . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Crack Repair . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Flange Crack Limits . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202 Flange Repair . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202

78-CONTENTS

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List of Effective Pages CH-SE-SU

PAGE

DATE

78-LOEP

1

Nov 1/09

78-CONTENTS

1

Nov 1/09

78-00-00

1 and 2 201 thru 204

Nov 1/09 Nov 1/09

78-LOEP

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EXHAUST EXHAUST SYSTEM DESCRIPTION AND OPERATION

78-00-00 00

1. GENERAL The exhaust system consists of the engine exhaust stacks and the engine air inlet anti-ice lip (Ref. Figure 1). The exhaust stack is an “L” shaped tube that is mounted directly to the engine and channels hot engine exhaust gases past the cowlings and away from the engine. The engine anti-ice lip is a continuous duct that channels hot engine exhaust gases from the engine's left exhaust stack around the front edge of the engine air inlet opening to the engine's right exhaust stack. The purpose of this lip is to prevent the formation of ice in and around the engine air intake during inclement weather.

78-00-00

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Figure 1 Engine Exhaust System

Page 2 Nov 1/09

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

EXHAUST EXHAUST SYSTEM MAINTENANCE PRACTICES

200200

1. EXHAUST STACK The exhaust stacks are mounted directly to the engine exhaust ports and can be removed for service or replacement by the following procedure:

A. Removal (1) Remove the upper forward, upper aft, upper mid and lower forward engine cowlings in that order (Ref. COWLING REMOVAL, Chapter 71-10-00). (2) Remove the brackets for the fire detection sensor cable and note their placement (Ref. Chapter 26-10-00). (3) Remove the twelve bolts that secure the exhaust stack to the engine exhaust port and remove the exhaust stack.

B. Installation NOTE: Existing exhaust stacks may be chafed to a maximum of 20% of original material thickness without repair. (1) Position the exhaust stack on the engine exhaust port. NOTE: The exhaust stacks must clear the cowling by a minimum of 1/2-inch. The cowling may be trimmed as required to provide clearance. (2) Secure the exhaust stacks with the twelve attaching bolts and torque the bolts to 60 ± 10 inch-pounds. (3) Install the brackets for the fire detection sensor cable (Ref. Chapter 26-10-00). (4) Install the lower forward, upper mid, upper aft and upper forward engine cowlings in that order (Ref. COWLING INSTALLATION, Chapter 71-10-00).

C. Crack Repair NOTE: If there are cracks in the mounting flange, refer to EXHAUST STACK FLANGE CRACK LIMITS, in this Chapter. Repair cracks that are three inches or less in length. Exhaust stacks having cracks greater than three inches in length must be replaced. Cracks in exhaust stacks may be repaired as follows: (1) Remove the exhaust stack from the engine. Refer to EXHAUST STACK REMOVAL, in this Chapter. (2) Remove all carbon from the interior and exterior surfaces using hot soapy water and a suitable brush. Allow the exhaust stack the dry thoroughly. (3) Wire brush both the inside and outside of the area requiring a weld, using a brush with stainless steel bristles.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (4) Rinse the area to be repaired with solvent (14, Table 1, Chapter 91-00-00) and allow to dry. (5) Prior to welding the crack, accomplish one of the following procedures: (a) Place a copper sheet, a minimum of 0.063-inch thick, against the inside surface of the crack. (b) Cap the openings in the exhaust stack with wooden blocks and seal the remaining small crevices with tape. Introduce an argon gas purge at 10 cubic feet per hour five minutes before welding; continue until the repair is complete. (6) Repair the crack using the tungsten inert gas (TIG) process. The welding rod type is determined by the exhaust stack material. Stainless Steel stacks use rod 347 (157, Table 1, Chapter 91-00-00), Incoloy 800 stacks use Incoloy 82 rod (160, Table 1, Chapter 91-00-00) and on the latest style stack made from Inconel 625 use rod 625 (156, Table 1, Chapter 91-00-00). (7) Install the exhaust stack. Refer to EXHAUST STACK INSTALLATION, in this Chapter.

D. Flange Crack Limits Airplanes may operate with cracked exhaust stack flanges which do not exceed the limitations listed below (Ref. Figure 201). If any crack exceeds these limits, the exhaust stack must be repaired or replaced. (1) No more than 8 of the 12 mounting holes may be cracked and the remaining 4 holes (non-cracked) must be 90° apart. (2) No more than 2 consecutive mounting holes may be cracked. (3) Cracks that radiate from a mounting hole to the inside diameter of the mounting flange or a crack that radiates from mounting hole to mounting hole is not acceptable.

E. Flange Repair Airplanes may operate with cracked exhaust stack flanges which do not exceed the limitations listed below (Ref. Figure 201). If any crack exceeds these limits, the exhaust stack must be repaired or replaced. (1) Remove the exhaust stack from the engine as instructed in EXHAUST STACK REMOVAL, in this Chapter. (2) Remove all carbon from the flange surfaces using hot soapy water and a suitable brush. Allow the exhaust stack to dry thoroughly. (3) Wire brush both sides of the area to be welded using a brush with fine stainless steel bristles. (4) Rinse the area to be repaired with solvent (14 or 30, Table 1, Chapter 91-00-00). (5) Place a copper sheet, a minimum of 0.063-inch thick, against the mating side of the flange. NOTE: Weld from the bolt head side of the flange. (6) Repair the crack using the tungsten inert gas (TIG) process and welding rod (157, Table 1, Chapter 91-00-00). (7) Install the exhaust stack. Refer to EXHAUST STACK INSTALLATION, in this Chapter.

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Figure 201 Exhaust Stack Crack Limits

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CHAPTER 79 - OIL TABLE OF CONTENTS SUBJECT

PAGE

OIL SYSTEM 79-00-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Procedures . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Servicing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Changing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201 Oil Filter . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202 Cleaning . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202 Magnetic Drain Plug Inspection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 204 Chip Detector (Annunciators Installed in Caution/Advisory Panel) . . . . . . . . . . . . . . . . . . . . . . . . . . . 204 Chip Detector (Annunciators not Installed in Caution/Advisory Panel) . . . . . . . . . . . . . . . . . . . . . . . . 204 Oil Cooler . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 206 Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 206 Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 206 Oil Pressure Switch and Transducer . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 206

OIL TANK 79-10-01 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Oil Tank Drain Shutoff Valve . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 201 201

OIL BREATHER (UA-1 THRU UA-3, UB-1 THRU UB-52 WITHOUT KIT 114-9006-1) 79-10-02 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Oil Breather Vent Hose . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

201 201 201 201

OIL BREATHER (UB-53 AND AFTER, UC-1 AND AFTER AND (UA-1 THRU UA-3, UB-1 THRU UB-52 WITH KIT 114-9006-1 INSTALLED)) 79-10-03 Maintenance Practices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Oil Breather Vent Hose (Aft Engine Enclosure) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Oil Breather Vent Tube (Aft Wheel Well) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Removal . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

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OIL OIL SYSTEM DESCRIPTION AND OPERATION

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1. GENERAL The turboprop engines used on this airplane have integral oil tanks. All components of the oil system, except the oil cooler, are mounted on or within the engine (Ref. Figure 1). The oil pressure relief valve is preset to maintain the proper oil pressure and normally will require no adjustment. The oil system has two scavenge pumps that return the lubricating oil to the oil cooling system. The oil cooler is mounted in the center of the lower air inlet duct below the engine. The oil cooler contains a thermostat type valve that bypasses oil around the cooler until proper operating temperature is reached. A bypass valve allows the oil to circulate around the cooler in the event the cooler becomes clogged. For a more detailed description of the engine oil system refer to the Pratt & Whitney PT6A-65B ENGINE MAINTENANCE MANUAL. Each engine oil pressure gage and temperature gage is internally lighted by 28 vdc from the triple fed buss through a 5-amp circuit breaker. The circuit breakers are identified LEFT and RIGHT OIL TEMP PRESS. Normal oil pressure is 90 to 135 psi green arc and the minimum oil pressure limit is 60 psi red line. Ref low pressure warning lights for the left and right engines, located in the warning annunciator panel, will illuminate if oil pressure drops below 60 psi. Normal oil temperature is 0° to 99°C and red line is 110°C. The oil pressure transmitter, the temperature bulb and the low pressure sensor are located on the engine accessory case at approximately the 4 o’clock position as viewed from the rear. A self-sealing chip detector is installed in the 6 o’clock position on the reduction gear case of each engine.

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Figure 1 Engine Lubrication System Schematic

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OIL OIL SYSTEM MAINTENANCE PRACTICES

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1. PROCEDURES CAUTION: Anytime the oil system is contaminated by metal particles, the oil cooler must be replaced and the oil system flushed to prevent engine damage. In addition, all airplane components and associated plumbing utilizing the engine oil system, such as propeller governors, must be flushed until free of contamination.

A. Servicing Servicing the engine oil system primarily involves maintaining the engine oil at the proper level, inspecting and cleaning the engine oil filter and changing the engine oil when the oil becomes contaminated. Refer to Chapter 12-10-00 for the oil system filling procedure. The oil filter should be cleaned and inspected every 200 hours and at each engine oil change, and replaced as required. Refer to the CHANGING THE ENGINE OIL and OIL FILTER CLEANING procedures.

B. Changing CAUTION: When changing to a different brand of oil, completely drain the airplane oil system as indicated in this procedure. Remove the oil filter and immerse it in the brand of oil to be used. Install the oil filter and drain plugs. Fill the system to the proper level, and ground run the engines for 20 minutes to thoroughly circulate the new brand of oil throughout the system. Completely drain the oil system and again remove the oil filter and immerse it in the new brand of oil. Refill the oil system as indicated below. This will thoroughly purge the system of the oil to prevent chemical interaction between it and the new brand. Remove the upper forward, upper aft and lower forward cowlings to gain access to the engine oil drains. Remove the small cover on the lower aft cowling to gain access to the oil cooler oil drain. (1) Provide suitable containers, funnels and drain hoses to facilitate oil drainage and to prevent unnecessary oil spillage. (2) Remove the dust cap from the oil tank drain line at the lower right side of the accessory section. (3) Attach an oil drain tube to the end of the engine oil tank drain line. (4) Remove the safety wire from the shutoff valve and turn it to the open position. (5) Remove the safety wire and drain plug from the oil cooler. (6) Remove the oil filter assembly. Refer to the OIL FILTER CLEANING procedure. (7) Remove the chip detector at the 6 o’clock position on the reduction gearbox front case. (8) Remove the drain plug from the oil-to-fuel heater.

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(9) With all the drain plugs removed, motor the engine over with the starter only (no ignition) to permit the scavenge pumps to clear the engine. CAUTION: Limit motoring to the time required to accomplish the above because of the limited lubrication available to the engine during this operation. To prevent damage to the fuel control unit, leave the condition lever in IDLE CUT-OFF while motoring the engine. (10) Install the oil filter assembly. Refer to the OIL FILTER CLEANING procedure. (11) Install new preformed packings on the drain plugs and the chip detector. (12) Install the chip detector on the reduction gearbox front case. Torque to 50 ± 5 in.-lbs and safety. (13) Install the drain plug on the oil cooler and safety. (14) Install the drain plug in the oil-to-fuel heater and torque to 70 ± 5 in.-lbs, then safety. (15) Remove the oil drain tube from the end of the engine oil tank drain line and install the dust cap, then safety wire the shutoff valve in the closed position. NOTE: The oil tank capacity is 2.5 gallons (10 quarts). An additional 4.4 quarts of oil is required to fill the lines and oil cooler, giving a total system capacity of 14.4 quarts; however, because of the residual oil trapped in the system, no more than 13 quarts should be added during an oil change. (16) Fill the engine with the correct amount and type of oil. CAUTION: Observe the starter operating limits of 30 seconds ON, 10 minutes OFF, 30 seconds ON, 30 minutes OFF, 30 seconds ON, 30 minutes OFF, 30 seconds ON, 30 minutes OFF, 30 seconds ON, 30 minutes OFF. (17) Motor the engine over, with the starter only, long enough to get an oil pressure reading. (18) Check engine for oil leaks. (19) Fill the engine to proper oil level. (20) Install the lower forward, upper aft and the upper forward cowlings. Install the small cover on the lower after cowling.

2. OIL FILTER A. Cleaning The engine oil filter contains a metal screen element. It is located under the square cover plate at the 3 o’clock position of the compressor inlet case and just behind the aft fire seal. The filter should be cleaned and inspected at each engine oil change and at each inspection interval specified in Chapter 5-20-00, CONTINUOUS INSPECTION SCHEDULE. (1) Remove the four self-locking nuts and washer securing the filter cover to the compressor inlet case and remove the cover. (2) Withdraw the filter from the filter housing, then agitate the filter screen for 5 minutes in clean unused solvent (31, Table 1, Chapter 91-00-00) (Ref. Figure 201). Page 202 Nov 1/09

79-00-00

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (3) Dry the filter element with clean, dry, compressed air. If such air is not available, let the filter stand until dry. (4) Inspect the filter element with a strong magnifying glass. If more than 5 percent of the visible passages are blocked, the filter element must be cleaned and inspected at an overhaul facility. (5) If dents or broken wires are found on the filter screen, the filter must be replaced. (6) Insert the filter element (perforated, flanged end first) into the filter housing. (7) Coat a new packing with engine oil and install the seal and cover on the engine. (8) Secure the filter cover with four plain washers and self-locking nuts. Torque the nuts to 34 ± 2 in.-lbs above the torque necessary to turn the nuts before seating.

Figure 201 Engine Oil Filter

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3. MAGNETIC DRAIN PLUG INSPECTION A magnetic drain plug (chip detector) is installed in each engine. The plug attracts ferrous particles which may be present in the sump area.

A. Chip Detector (Annunciators Installed in Caution/Advisory Panel) Inspect the magnetic drain plug as follows: (1) Remove the electrical connector. (2) Remove the magnetic drain plug. NOTE: When the drain plug is properly removed, the check valve in the drain plug fitting will prevent the oil from draining out. (3) Inspect the drain plug. Refer to the Pratt & Whitney PT6-65B ENGINE MAINTENANCE MANUAL. (4) Install the electrical connector finger tight. (5) Apply electrical power to the airplane. (6) To short the detector, position metallic object in the gap of the chip detector. (7) Check to see if the chip detector light in the cockpit is illuminated. If not, repeat the above procedures. (8) Remove the metallic object and check to ensure the light does not stay illuminated. (9) Remove the electrical connector from the chip detector. (10) Install new packing on the magnetic plug. (11) Install the magnetic plug and torque to 50 ± 5 in.-lbs. (12) Install the electrical connector to the chip detector. (13) Safety wire the chip detector.

B. Chip Detector (Annunciators not Installed in Caution/Advisory Panel) Inspect the magnetic chip detectors as follows: (1) Gain access to the terminal strip on the upper right side of the aft fireseal (Ref. Figure 202). (2) Check the wiring between the terminal strip and the chip detector for broken or otherwise damaged wires. (3) Check for continuity between terminals 1 and 2 of the terminal strip by using a continuity checker or ohmmeter. (4) If no continuity is found, terminate this procedure. If continuity is indicated, continue this procedure. (5) Remove the electrical connector.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (6) Remove the magnetic drain plug. NOTE: When the drain plug is properly removed, the check valve in the drain plug fitting will prevent the oil from draining out. (7) Inspect the drain plug. Refer to the Pratt & Whitney PT6-65B ENGINE MAINTENANCE MANUAL. (8) Install the new packing on the magnetic plug. (9) Install the magnetic plug and torque to 50 ± 5 in.-lbs. (10) Install the electrical connector. (11) Safety wire the chip detector.

Figure 202 Chip Detector Terminal Strip

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4. OIL COOLER A. Removal (1) Remove the cowling (Ref. COWLING REMOVAL, Chapter 71-10-00). (2) Place a suitable drip pan under the engine to catch oil spillage. (3) Unsafety and remove the oil cooler drain plug. (4) Disconnect the oil inlet and outlet hose assemblies at the cooler. Cover the open ports. (5) Remove the bolts attaching the cooler to the support brackets.

B. Installation (1) Position the oil cooler against the support bracket and install the attaching bolts. (2) Uncover open ports and connect the oil inlet and outlet hose assemblies on the cooler. (3) Install and safety the oil cooler drain plug. (4) Fill the oil system (Ref. Chapter 12-10-00). CAUTION: Observe the starter operating limits of 30 seconds ON, 10 minutes OFF, 30 seconds ON, 30 minutes OFF, 30 seconds ON, 30 minutes OFF, 30 seconds ON, 30 minutes OFF, 30 seconds ON, 30 minutes OFF. (5) Motor the engine over with the starter (no ignition) long enough to obtain an oil pressure reading. (6) Check in the vicinity of the oil cooler inlet and outlet hose assemblies for leakage. (7) Install the cowling (Ref. COWLING INSTALLATION, Chapter 71-10-00).

5. OIL PRESSURE SWITCH AND TRANSDUCER When installing the oil pressure switch, transducer or associated fittings, the maximum torque limit is 40 in.-lbs.

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OIL OIL TANK MAINTENANCE PRACTICES

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1. OIL TANK DRAIN SHUTOFF VALVE A. Removal (1) Perform the appropriate engine cowling removal procedure (Ref. Chapter 71-10-00). (2) Place a drip pan under the engine to catch any oil spillage. (3) Remove the dust cap (5) from the shutoff valve (4) (Ref. Figure 201) and connect an engine drain tube (Ref. Chapter 91-00-00, Figure 1, Sheet 3). (4) Remove the safety wire and open the shutoff valve (4) to drain the engine oil tank (Ref. Figure 201). (5) Remove the engine drain tube from the shutoff valve (4). (6) Disconnect the drain hose (1) from the shutoff valve (4). CAUTION: To prevent contamination of the oil system after the drain hose (1) is disconnected, cap the hose end. (7) Remove nut (2) from the shutoff valve (4) and remove the shutoff valve (4) and washers (3) from the bracket.

B. Installation (1) Install the shutoff valve (4) and washers (3) into the bracket (Ref. Figure 201). (2) Install nut (2) on the shutoff valve (4) to secure the shutoff valve to the bracket. (3) Connect the drain hose (1) to the shutoff valve (4). (4) Place the shutoff valve (4) in the closed position and safety wire. (5) Install the dust cap (5) on the shutoff valve (4). (6) Check the oil level and add oil as necessary to replace the oil drained from the tank (Ref. Chapter 12-10-00). (7) Motor the engine over with the starter only (no ignition) long enough to obtain an oil pressure reading. CAUTION: For continuous motoring without engine start, observe the starter operating limits of 20 seconds on, 5 minutes off. (8) Check the area around the shutoff valve for leaks. (9) If no leaks are found, perform the appropriate engine cowling installation procedure (Ref. Chapter 71-10-00).

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Figure 201 Oil Tank Drain Valve Removal/Installation

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OIL OIL BREATHER (UA-1 THRU UA-3, UB-1 THRU UB-52 WITHOUT KIT 114-9006-1) MAINTENANCE PRACTICES 1. OIL BREATHER VENT HOSE A. Removal (1) Remove the access panel from the right side of the engine cowling (Ref. Figure 201). (2) Remove the clamps which secure the oil breather vent hose (3) in the aft engine compartment. (3) Loosen hose clamps (2 and 4) and remove the oil breather vent hose (3) from the oil breather adapter (1) on the engine accessory gearbox housing and from the weld assembly (5) at the engine firewall. (4) Remove the oil breather vent hose (3) from the airplane.

B. Installation (1) Install the clamps to secure the oil breather vent hose in the aft engine compartment (Ref. Figure 201). (2) Install the oil breather vent hose (3) on the oil breather adapter (1) on the engine accessory gearbox and on the weld assembly (5) at the engine firewall and secure the hose with the attaching hose clamps (2 and 4). (3) Install the access panel on the right side of the engine cowling.

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Figure 201 Oil Breather Vent System Installation

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OIL OIL BREATHER (UB-53 AND AFTER, UC-1 AND AFTER AND (UA-1 THRU UA-3, UB-1 THRU UB-52 WITH KIT 114-9006-1 INSTALLED)) MAINTENANCE PRACTICES

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1. OIL BREATHER VENT HOSE (AFT ENGINE ENCLOSURE) A. Removal (1) Remove the access panel from the right side of the engine cowling (Ref. Figure 201). (2) Remove the clamps which secure the oil breather vent hose (3) in the aft engine compartment. (3) Loosen the hose clamps (2) and remove the oil breather vent hose (3) from the oil breather adapter (1) on the engine accessory gearbox housing and from the elbow on the aft engine firewall (7). (4) Remove the oil breather vent hose (3) from the airplane.

B. Installation (1) Install the clamps to secure the oil breather vent hose in the aft engine compartment (Ref. Figure 201). (2) Install the oil breather vent hose (3) on the oil breather adapter (1) on the engine accessory gearbox and on the elbow on the aft engine firewall (7) and secure the hose with the attaching hose clamps (2). (3) Install the access panel on the right side of the engine cowling.

2. OIL BREATHER VENT TUBE (AFT WHEEL WELL) A. Removal (1) Perform the appropriate NACELLE INNER FENDER REMOVAL procedure (Ref. Chapter 54-10-01). (2) Disconnect and remove the hose clamps (2) from the connecting hoses (6) between the oil breather vent tubes (4) (Ref. Figure 201). (3) Disconnect the plumbing fitting on the oil breather vent tube at the aft side of the firewall (7). (4) Disconnect and remove the clamps which secure the oil breather vent tubes to the structure in the wheel well. Slip the connecting hoses (6) from the tubes (4) and remove the tubes.

B. Installation (1) Position the oil breather vent tube (4) to the aft side of the firewall, and secure it with the plumbing fitting (Ref. Figure 201). (2) Connect each tube (4) section to the connecting hoses (6), and position each hose for proper clearance in the wheel well.

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL (3) Install the hose clamps (2) which secure the hoses to the oil breather vent tubes (4). (4) Install clamps that secure the oil breather vent tube to the structure in the wheel well. (5) Perform the appropriate NACELLE INNER FENDER INSTALLATION procedure (Ref. Chapter 54-10-01).

Page 202 Nov 1/10

79-10-03

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

1. 2. 3. 4. 5. 6. 7.

OIL BREATHER ADAPTOR HOSE CLAMP OIL BREATHER VENT HOSE OIL BREATHER VENT TUBE NACELLE SKIN HOSE FIREWALL

B

A

2 5

2

1

4

3 2

4 6

1

7

2

6

2 RIGHT HAND WHEEL WELL AREA DETAIL 3

B

4 2

4

2 2 4

7

2 6

5

6 LEFT HAND WHEEL WELL AREA DETAIL

A

UC79B 100364AB.AI

Figure 201 Oil Breather Vent System Installation

79-10-03

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

CHAPTER 80 - STARTING TABLE OF CONTENTS SUBJECT

PAGE

STARTING SYSTEM 80-00-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 General . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1

80-CONTENTS

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

STARTING STARTING SYSTEM DESCRIPTION AND OPERATION

80-00-00 00

1. GENERAL The starter is a starter/generator. The same unit serves both the function of starter during engine cranking and generator after the engine is started. Under adverse starting conditions, the unit requires a power source capable of supplying 1250 to 1400 amps for a short duration and 300 to 400 amps continuous at a maximum of 28 volts. The starter has operating limitations of 30 seconds ON, 10 minutes OFF, 30 seconds ON, 10 minutes OFF, 30 seconds ON, 30 minutes OFF, 30 seconds ON, 30 minutes OFF, 30 seconds ON, 30 minutes OFF. When operating as a generator, the unit is capable of delivering 300 amps at 28.25 ± 0.25 volts (Ref. Figure 1). For a more complete description and operation or for maintenance of the system refer to Chapter 24 or to the appropriate supplier publication.

80-00-00

Page 1 Nov 1/09

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 1 Simplified Start Circuit

Page 2 Nov 1/09

80-00-00

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

CHAPTER 91 - CHARTS TABLE OF CONTENTS SUBJECT

PAGE

CONSUMABLE MATERIALS/SPECIAL TOOLS AND EQUIPMENT 91-00-00 Description and Operation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Consumable Materials . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Tables . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Sealing Materials . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 21 Torquing for Coarse Thread Series Bolt-Nut Combination . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 23 Torquing for Fine Thread Series Bolt-Nut Combination . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 24 Flare Fitting Torque . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 25 Thread Lubricants for Fluid-line Fittings . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 26 Special Tools and Equipment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 27

91-CONTENTS

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1 thru 39

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

CHARTS CONSUMABLE MATERIALS/SPECIAL TOOLS AND EQUIPMENT DESCRIPTION AND OPERATION

91-00-00 00

The recommended materials listed in Table 1 as meeting federal, military or supplier specifications are provided for reference only and are not specifically recommended by Hawker Beechcraft Corporation. The products included in this Table have been tested and approved for aviation usage by Hawker Beechcraft Corporation, by the supplier or by compliance with the applicable specification. Generic or locally manufactured products which conform to the requirements of the specification listed may be used even though not included herein. Only the basic number of each product specification is listed. No attempt has been made to update the listing to the latest revision. It is the responsibility of the technician or mechanic to determine the current revision of the applicable specification prior to usage of the product listed. This can be done by contacting the supplier of the product to be used.

1. CONSUMABLE MATERIALS A. Tables NOTE: For the various engine oils approved by Pratt and Whitney, refer to the Pratt and Whitney Service Bulletin No. 13001. Table 1 Consumable Materials MATERIAL

SPECIFICATION

PRODUCTS

(1) Engine Oil (2) Solvent

SUPPLIER (Refer to NOTE above).

PD-680, Type III

Barton Solvents, Inc. 201 S. Cedar Street Valley Center, Kansas 67147

(3) Urethane Primer

83-Y-3 Urethane Primer

U.S. Paint, Inc. 831 S. 21st Street St. Louis, Missouri 63103

(4) Catalyst, Urethane Primer

83-C-13 Catalyst, Urethane Primer

U.S. Paint, Inc. 831 S. 21st Street St. Louis, Missouri 63103

(5) Primer, (Epoxy Polyamide)

MIL-P-23377

Sterling Lacquer Manufacturing Company 3150 Brannon Avenue St. Louis, Missouri 63139

(6) Urethane Paint

6160 Matterhorn White

U.S. Paint, Inc. 831 S. 21st Street St. Louis, Missouri 63103

(7) Sanding Surfacer

P900 Primer Surfacer Ameritex Chemical and Coatings Company, Inc. 801 E. Lee Street Irving, Texas 75060

91-00-00

Page 1 Nov 1/11

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Table 1 Consumable Materials (Continued) MATERIAL

SPECIFICATION

(8) Paint Stripper

(9) Desiccant

PRODUCTS Turco 4260 Epocon

MIL-D-3464

SUPPLIER Turco Products, Inc. 7300 Bolsa Avenue Westminster, California 92684 Obtain Locally

(10) Oil

MIL-D-22851

Grade 1100

Obtain Locally

(11) Corrosion Preventative Compound

MIL-C-16173, Grade 2

Petrotect Grade 2 (use only products qualified under products list 16173, such as LPS-3)

Obtain Locally

(12) Primer

MIL-P-8585

Zinc Chromate

Obtain Locally

(13) Preservative Hydraulic MIL-H-6083 Fluid

Avrex 904

Mobil Oil Corporation 3225 Gallows Rd. Fairfax, VA 22037

(14) Solvent

Methyl Propyl Ketone (MPK)

Obtain Locally

(15) Turco Metal-Glo No. 3

Turco Products, Inc. 7300 Bolsa Avenue Westminster, California 92684

(16) Adhesive

EC1300L

Minnesota Mining and Manufacturing Company 3M Center St. Paul, Minnesota 55144

(17) Solvent

(Refer to item 54.)

(18) Solvent

TT-T-548

Technical Toluene

Obtain Locally

(19) Sealer

AMS-S-8802 Type 2 Class A and B

PR-1425 B1/2 Class A and B

PRC DeSoto International, Inc. 5454 San Fernando Road P.O. Box 1800 Glendale, California 91209

Icex

B.F. Goodrich Company P.O. Box 5471 Akron, Ohio 44313

(20) Compound

(21) Trichloroethane

ASTM D4080 (Supersedes O-T-634)

(22) Cement

(23) Grease, Aircraft and Instrument, Gear and Actuator Screw

Page 2 Nov 1/11

MIL-PFR-23827

91-00-00

Obtain Locally A56B

B.F. Goodrich Company P.O. Box 5471 Akron, Ohio 44313

Super Mil Grease No. A72832

AMOCO, Inc. 2021 Spring Road Oakbrook, Illinois 60521

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Table 1 Consumable Materials (Continued) MATERIAL

SPECIFICATION MIL-PFR-23827 Type II Do Not mix with Type I.

(24) Acetone

ASTM D-329 (Supersedes O-A-51)

(25) Cement, General Purpose

MIL-A-1154

(26) Solvent, Propeller Slip Ring Cleaning

PRODUCTS

SUPPLIER

Aeroshell 33

Shell Oil Co., One Shell Plaza, P.O. Box 2463, Houston, TX 77001 Obtain Locally

EC1403

Minnesota Mining and Manufacturing Company 3M Center St. Paul, Minnesota 55144

CRC-2-26

CRC Industries 885 Louis Drive Warminster, Pennsylvania 18974

(27) Tape, Anti-Seize

MIL-T-27730

Obtain Locally

(28) Aviator's Breathing Oxygen

MIL-O-27210

Obtain Locally

(29) Rubber Protect Agent

MIL-P-11520E

Age Master No. 1

B.F. Goodrich Company P.O. Box 5471 Akron, Ohio 44313

(30) Solvent

TT-I-735

Isopropyl Alcohol

Obtain Locally

Varsol

Exxon Company, U.S.A. P.O. Box 2180 800 Bell Street Houston, Texas 77252

Grade 1010

Obtain Locally

Permatex No. 2

Permatex Company, Inc. Division of Loctite Corporation 705 N. Mountain Road Newington, Connecticut 06111

Thread Lubricant (Refer to item 76.)

Obtain Locally

(35) Oakite Solution

No. 6

Oakite Products, Inc. 50 Valley Road Berkeley Heights, New Jersey 07922

(36) Sealer

EC1675B1/2

Minnesota Mining and Manufacturing Company 3M Center St. Paul, Minnesota 55144

(31) Solvent

(32) Oil, Preserving

MIL-L-6081

(33) Gasket Cement

(34) Petrolated Graphite

MIL-T-5544

91-00-00

Page 3 Nov 1/11

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Table 1 Consumable Materials (Continued) MATERIAL

SPECIFICATION

(37) Sealer

PRODUCTS

SUPPLIER

EC1675A1/2

Minnesota Mining and Manufacturing Company 3M Center St. Paul, Minnesota 55144

(38) Sealer

AMS-S-8802 Type II

PR-1440B1/2

Courtaulds Aerospace, Inc. P.O. Box 1800 5454 San Fernando Road Glendale, California 91209

(39) Hydraulic Fluid

MIL-H-5606

Brayco 756

Castrol Industrial, N. America 1100 W. 31st st. Downers Grove, IL 60515

Mobil Aero HF

Mobil Oil Corporation 3225 Gallows Rd. Fairfax, VA 22037

Aeroshell 14

Shell Oil Co. One Shell Plaza Houston, TX 77210

Caltex Low Temp Oil

Caltex Oil Products Company 125 E. John W. Carpenter Freeway Irving, Texas 75062

(42) Aerodynamic Smoother

Whitestar

Fibre Glass - Evercoat Co. Inc. 6600 Cornell Road Cincinnati, Ohio 45242

(43) Translucent Adhesive

RTV-108

General Electric Company Silicone Products Department Hudson River Road Waterford, New York 12188

(45) Lubricant

Door Ease

American Grease Stick Company 2651 Hoyt Muskegon, Michigan 49443

(46) Lubricant

Silicon 4 Compound

Dow Corning P.O. Box 997 3901 S. Saginaw Road Midland, Michigan 48602

(40) Sealer

(Refer to item 166.)

(41) Lubricating Oil, General Purpose Low Temperature

MIL-L-7870

(44) Adhesive

(Refer to item 57.)

(47) Leak Detector Fluid, Oxygen System

Page 4 Nov 1/11

MIL-L-25567C, Type I

91-00-00

Obtain Locally

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Table 1 Consumable Materials (Continued) MATERIAL

SPECIFICATION

(48) Plexiglas Polish and Cleaner

Federal Specification P-P-560

(49) Kerosene

VV-K-211

(50) Plexiglas Scratch Remover

PRODUCTS

SUPPLIER

Tend Plastic Cleaner and Polish

Regal Plastic 329 N. Indiana Wichita, Kansas 67214 Obtain Locally

HP-100C

Micro-Mesh Cushioned Abrasives Micro-Surface Finishing Products P.O. Box 818 1217 W. Third Street Wilton, Iowa 52778 Polysand Cushioned Abrasives Fredrick B. Anthon Enterprises U.S. Distributor-Cope Plastics 4441 Industrial Drive Godfrey, Illinois 62035

(51) Hexane

J.T. Baker Chemical Company 1037 Lower Brownville Road Jackson, Tennessee 38301 Mallinckrodt, Inc. Science Products Division P.O. Box 5840 St. Louis, Missouri 63134

(52) Adhesive

Presstite No. 576

Inmont Corporation 3900 Chouateau Avenue St. Louis, Missouri 63110

(53) Sealing Tape

PF 5422 Weatherban

Minnesota Mining and Manufacturing Company 3M Center St. Paul, Minnesota 55144

Aliphatic Naphtha

Barton Solvents, Inc. 201 S. Cedar Valley Center, Kansas 67147

EA9304 PT

Dexter Hysol Aerospace, Inc. P.O. Box 312 2850 Willow Pass Road Pittsburg, California 94565

Dapcotac #3300 (Supersedes DOE A-4000)

D. Aircraft Products Company 1191 Hawk Circle Anaheim, California 92708

(54) Solvent

TT-N-95, Type II

(55) Epoxy Adhesive

(56) Sealer

(Refer to item 166.)

(57) Silicone Adhesive

(58) Sealer

(Refer to item 166.)

91-00-00

Page 5 Nov 1/11

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Table 1 Consumable Materials (Continued) MATERIAL

SPECIFICATION

PRODUCTS

SUPPLIER

(59) Thread Locking Compound

Loctite Sealer Grade A (for studs under 2 inches in dia.) Studloc (for studs over 2 inches in dia.)

Loctite Corp., 705 N. Mountain Road Newington, Connecticut 06111

(60) Sealer

Scotch Seal or Weatherban 2084

Minnesota Mining and Manufacturing Company 3M Center St. Paul, Minnesota 55144

Mobilgrease 28

Mobil Oil Corporation 3225 Gallows Rd. Fairfax, VA 22037

(61) Lubricating Grease, Wide Temperature Range

MIL-G-81322

Mobil Aviation Grease Mobil Oil Corporation SHC 100 3225 Gallows Rd. Fairfax, VA 22037

(62) Jet-Fuel Anti-icing Additive

MIL-DTL-85470 or ASTM D4171, Type III

Aeroshell Grease 22 (Refer to item 87.)

Shell Oil Company P.O. Box 2463 One Shell Plaza Houston, Texas 77001

Royco 22

Royal Lubricants Co. River Rd. Hanover, NJ 07936

UCAR Fuel Additive 500

Union Carbide Corporation Chemicals and Plastics Division 200 Cottontail Lane Somerset, New Jersey 08873

High Flash Prist

Prist Division of PPG Industries, Inc. 5926 F.M. 1960 West Houston, Texas 77069 Hoffman-Taff, Inc. P.O. Box 1246 Springfield, Missouri 65805

(63) Biocidal Agent

Page 6 Nov 1/11

91-00-00

Biobor® JF

U.S. Borax & Chemical Corp. 8600 W. Bryn Mawr, Suit 710, N. Chicago, IL 60631

Kathon® FP 1.5

Fuel Quantity Services Inc. P.O. Box 1380 Flowery Branch, GA 30542-0023

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Table 1 Consumable Materials (Continued) MATERIAL

SPECIFICATION

PRODUCTS

SUPPLIER

(64) Lubricating Grease

Plastilube No. 3

Thiem Corporation P.O. Box 93019 5151 Denison Avenue Cleveland, Ohio 44101

(65) Lubricant

CRC-3-36

CRC Industries 885 Louis Drive Westminster, Pennsylvania 18974

(66) Lubricant

LPS No. 1

LPS Laboratories 4647 Hugh Howell Road Tucker, Georgia 30084

(67) Spray Lubricant

WD-40

WD-40 Company 1061 Cudahy Place San Diego, California 92110 P.O. Box 80607 1061 Cudahy Place San Diego, California 92138

(68) Lubricating Grease

Lubriplate No. 130A or Lubriplate Aero

Fiske Brothers Refining Company 129 Lockwood Newark, New Jersey 07105

(69) Lubricant

Molykote M-77

Dow Corning P.O. Box 997 3901 S. Saginaw Road Midland, Michigan 48602 Dixie Bearing, Inc. 1408 Haines Avenue Jacksonville, Florida 32206

(70) Air-Conditioning, Compressor Oil

(71) Air-Conditioning Refrigerant

Suniso No. 5

Virginia Chemical, Inc. 801 Water Street Portsmouth, Virginia 23704

Capella WF-100

Texaco, Inc. 135 E. 42nd, New York, New York 10017

Racon 12

Racon, Inc. 6040 S. Ridge Road Wichita, Kansas 67215

Genetron 12

Allied Chemical P.O. Box 1021 101 Columbia Road Morristown, New Jersey 07960

91-00-00

Page 7 Nov 1/11

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Table 1 Consumable Materials (Continued) MATERIAL

SPECIFICATION

PRODUCTS

SUPPLIER

Freon 12

DuPont, Inc. Freon Products Division 1007 Market Street Wilmington, Delaware 19898

Exxon 2389

Exxon Company P.O. Box 2180 800 Bell Street Houston, Texas 77252

(73) Compressor Cleaning Fluid

Turco 4217

Turco Products, Inc. 7300 Bolsa Avenue Westminster, California 92684

(74) Tire Sealer

Aero Seal (Gal. or Pt.)

Hawker Beechcraft Service Centers

(75) Grease, High Temperature

Aeroshell Grease 5

Shell Oil Company P.O. Box 2463 One Shell Plaza Houston, Texas 77001

(72) Oil, Air Cycle Machine For Alternate: see item 171.

(76) Antiseize Thread Compound, Graphite Petrolatum

MIL-L-7808G

W.J. Ruscoe and Company P.O. Box 3858 483 Kenmore Boulevard Akron, Ohio 44314

MIL-PRF-83483 (Preferred) MIL-T-83483 (Acceptable)

(77) Air-Conditioning System Cleaner

Racon 11

Racon, Inc. 6040 S. Ridge Road Wichita, Kansas 67215

Genetron 11

Allied Chemical P.O. Box 1021 101 Columbia Road Morristown, New Jersey 07960

Freon 11

DuPont, Inc. Freon Products Division 1007 Market Street Wilmington, Delaware 19898

(78) Red Refrigerant Leak Detector Dye (R-12 Systems)

Trace

Highside Chemicals, Inc. P.O. Box 3748 Reichhold Road Gulfport, Mississippi 39505

(79) Grease

Aeroshell #17 (Refer to item 83.)

Shell Oil Company P.O. Box 2463 One Shell Plaza Houston, Texas 77001

Page 8 Nov 1/11

91-00-00

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Table 1 Consumable Materials (Continued) MATERIAL

SPECIFICATION

(80) Grease

(81) Grease

(82) Lubricant, Dry Film

(83) Grease

MIL-M-7866

MIL-L-23398

MIL-G-21164

PRODUCTS

SUPPLIER

Molykote G-n

Dow Corning P.O. Box 997 3901 S. Saginaw Road Midland, Michigan 48602

Molykote Z

Haskel, Inc. 100 E. Graham Place Burbank, CA 91502

Moly-Paul No. 4 K.S.

K.S. Paul Products Ltd. Corneliusstrasse 72-74 Postfach 622 Duesseldorf, Germany

Perma-Silk G

Everlube Products 100 Cooper Circle Peachtree City, GA 30269

Lubri-Bond 220

Everlube Products 100 Cooper Circle Peachtree City, GA 30269

Sandstrom #238

Sandstrom Products Company P.O. Box 547 224 S. Main St. Port Byron, IL 61275

Everlube 211-G Moly Grease

E/M Corporation P.O. Box 2200 Highway 52 N.W. West Lafayette, IN 47906

Royco 64

Royal Lubricants Co., Inc. P.O. Box 518 River Road East Hanover, NJ 07936

ARPOLUBE 21164

Arpol Petroleum Co. 225 Broadway New York, NY 10007

Braycote 664

Castrol, Inc. Specialty Products Division 16715 Von Karmen Ave. Suite 230 Irvine, CA 92714

Aeroshell Grease 17

Shell Oil Company P.O. Box 2463 One Shell Plaza Houston, Texas 77001

91-00-00

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Table 1 Consumable Materials (Continued) MATERIAL

SPECIFICATION

PRODUCTS Aeroshell 33MS grease

(84) Marvel Mystery Oil

SUPPLIER Shell Chemical CO. P.O. Box 2463 One Shell Plaza Houston, Texas 77001 Marvel Oil Company, Inc. 331-337 N. Main Street Port Chester, New York 10573

(85) Grease

MIL-G-10924

(86) Solvent

(Refer to item 54.)

(87) Wheel Bearing Grease

Shell A and A

Shell Oil Company P.O. Box 2463 One Shell Plaza Houston, Texas 77001

Mobilux EP No. 2 (Above -20°F)

Mobil Oil Corporation 3225 Gallows Rd. Fairfax, VA 22037

Chevron Polyurea EP No. 2 (Above -20°F)

Chevron U.S.A. 225 Bush Street San Francisco, California 94120

Rykon No. 2 EP (Above -20°F)

AMOCO, Inc. 021 Spring Road Oakbrook, Illinois 60521

Amdex No. 2 EP (Above -20°F)

AMOCO, Inc. 2021 Spring Road Oakbrook, Illinois 60521

AMSOIL (All temperatures)

AMSOIL 925 Tower Avenue Superior, Wisconsin 54880

AMS/OIL (GHD) (All temperatures)

AMSOIL 925 Tower Avenue Superior, Wisconsin 54880

Mobil Aviation Grease Mobil Oil Corporation SHC 100 3225 Gallows Rd. (All temperatures) Fairfax, VA 22037

(88) Chemical Conversion Coating for Aluminum and Aluminum Alloys

Page 10 Nov 1/11

MIL-C-81706 per MIL-C-5541

91-00-00

Mobilgrease 28

Mobil Oil Corporation 3225 Gallows Rd. Fairfax, VA 22037

Alodine 1200, 1200S, 1201, 1203, or 600

Parker and Amchem, Inc. 32100 Stephenson Highway Madison Heights, Michigan 48071

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Table 1 Consumable Materials (Continued) MATERIAL

SPECIFICATION

(89) Tape, Sealant, closed cell foam Polyvinylchloride (PVC), 0.125 inch thick, 1.25 inches wide, adhesive one side, grey (90) Lubricant, Petrolatum

(91) Retaining Compound

VV-P-236A

PRODUCTS

SUPPLIER

534 Tape

Norton Company One Sealants Park Grandville, New York 12832

Braycote 236

Castrol, Inc. 16715 Von Karman Avenue Irvine, California 92714

Loctite 680

Loctite Corporation 705 N. Mountain Road Newington, Connecticut 06111

(92) Primer

Loctite Corporation 705 N. Mountain Road Newington, Connecticut 06111

(93) Adhesive

EA9309NA

Dexter Hysol Aerospace, Inc. P.O. Box 312 2850 Willow Pass Road Pittsburg, California 94565

(94) Sealer, Fuel Tank

Pro Seal 860

Courtaulds Aerospace, Inc. P.O. Box 1800 5454 San Fernando Road Glendale, California 91209

Pro Seal 890

Courtaulds Aerospace, Inc. P.O. Box 1800 5454 San Fernando Road Glendale, California 91209

(95) Sealer, Fuel Tank

PR1435

Courtaulds Aerospace, Inc. P.O. Box 1800 5454 San Fernando Road Glendale, California 91209

(96) Cloth, Fiberglass

#181

Owens Corning Fiberglas Corporation Fiberglas Tower Toledo, Ohio 43604

(97) Cloth, Fiberglass

#1581

Owens Corning Fiberglas Corporation Fiberglas Tower Toledo, Ohio 43604

91-00-00

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Table 1 Consumable Materials (Continued) MATERIAL

SPECIFICATION

PRODUCTS

SUPPLIER

(98) Cloth, Fiberglass

#7781

Owens Corning Fiberglas Corporation Fiberglas Tower Toledo, Ohio 43604

(99) Core, Honeycomb, 1/8 inch cell, 3.0 pound density, 2 inches thick

HRH10-1/8 -3/-2

Hexcel Corporation P.O. Box 2197 20701 Nordhoff Chatsworth, California 91311

(100) Film, Peel Plies, 350°

Release Ply F

Airtech International, Inc. 2542 E. De Lamo Boulevard P.O. Box 6207 Carson, California 90749

Style 52006 code 51789

Precision Fabrics Group, Inc. 50 Chestnut Ridge Road Montvale, New Jersey 07645

Wrightlon 700

Airtech International, Inc. P.O. Box 6207 2542 E. De Lamo Boulevard Carson, California 90749

HS-8172-66

Richmond Division Dexico, Inc. P.O. Box 1129 Coltan and Opal Streets Redlands, California 92373

Frekote 33

Dexter Hysol Aerospace, Inc. P.O. Box 312 2850 Willow Pass Road Pittsburg, California 94565

Monocoat 63

Chem-Trend, Inc. 3205 E. Grand River, Howell, Michigan 48843

5162-2

Schnee-Morehead Chemicals 111 N. Nursery Road Irving, Texas 75060

Corseal 725

ADCO Industries 15 W. 6th Street, Suite 1210 Tulsa, Oklahoma 74119

Mobil Lubricant SHC 75W-90

Mobil Oil Corporation 3225 Gallows Rd. Fairfax, VA 22037

(101) Film, Vacuum Bag, Type I, 12-Hour, 350°

(102) Film, Release, Type IV, Non-Perforated, 1-mil thick

(103) Tape, Bag, Sealer, 350°

(104) Lubricant, 75 weight

Page 12 Nov 1/11

MIL-L-2105C (Supersedes MIL-L-10324 and MIL-O-6086)

91-00-00

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Table 1 Consumable Materials (Continued) MATERIAL

SPECIFICATION

(105) Fire Retardant, Fabric, Spray-on (106) Lubricant, Special Preservative

VV-L-800

(107) Lubricating Grease

PRODUCTS

SUPPLIER

2105 Grade 75 Weight

Southwest Petro-Chemical Division of WITCO 1400 S. Harrison Olathe, Kansas 66061

Inspecta-Shield

Kansas Fire Equipment 123 S. Osage Wichita, Kansas 67213

Brayco 300

Castrol, Inc. 16715 Von Karman Avenue Irvine, California 92714

Royco 308A

Royal Lubricants Company 6 Campus Drive Parsippany, New Jersey 07054

Lubriplate Aero

Fiske Brothers Refining Company 129 Lockwood Newark, New Jersey 07105

(108) Coating

MIL-I-46058

Conothane CE-1155

Conap, Inc. 1405 Buffalo Street Olean, New York 14760

(109) Coating

MIL-I-46058

Conothane CE-1164

Conap, Inc. 1405 Buffalo Street Olean, New York 14760

(110) Solvent

Product S8

Conap, Inc. 1405 Buffalo Street Olean, New York 14760

(111) Lubricant

LPS #2

LPS Laboratories 4647 Hugh Howell Road Tucker, Georgia 30084

(112) Tedlar Tape

838 Tape

Minnesota Mining and Manufacturing Company 3M Center St. Paul, Minnesota 55144

(113) Anti-Static Coating

528-104, 528-105 or 528-306 Flat Black Paint

DeSoto, Inc. Southwestern Plant Box 401268 Garland, Texas 75042

(114) Nylon Tie Straps

MS3367

Ty Wrap

Obtain Locally

(115) Fluid, Deicing/ Anti-icing, Aircraft

Type I per SAE AMS 1424 and ISO 11075

UCAR ADF Concentrate

Union Carbide Customer Center 10235 W. Little York Road Houston, Texas 77040

91-00-00

Page 13 Nov 1/11

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Table 1 Consumable Materials (Continued) MATERIAL

SPECIFICATION

PRODUCTS

Type I per SAE AMS 1424 and ISO 11075

UCAR ADF 50/50

Union Carbide Customer Center 10235 W. Little York Road Houston, Texas 77040

Type II per SAE AMS 1428 and ISO 11078

UCAR AAF UC5.1

Union Carbide Customer Center 10235 W. Little York Road Houston, Texas 77040

Type IV per SAE AMS 1428

UCAR ULTRA+

Union Carbide Customer Center 10235 W. Little York Road Houston, Texas 77040

Type I per SAE AMS 1424 and ISO 11075

Arcoplus

Arco Chemical Company 3801 West Chester Pike New Town Square, Pennsylvania 19073

Type I per SAE AMS 1424 and ISO 11075

Kilfrost DF

Kilfrost LTD Albion Works Haltwhistle Northumb NE49 0HJ UK

Type II per SAE AMS 1428 and ISO 11078

Kilfrost ABC-3

Kilfrost LTD Albion Works Haltwhistle Northumb NE49 0HJ UK

Type I per SAE AMS 1424 and ISO 11075

Hoechst VP1732

Hoechst Aktiengesellschaft Frankfurt am Main, Germany

Type II per SAE AMS 1428 and ISO 11078

Hoechst 1704 LTV

Hoechst Aktiengesellschaft Frankfurt am Main, Germany

Type III per SAE AMS 1428

Obtain Locally

Type IV per SAE AMS 1428

Hoechst Safewing MP Hoechst AG IV 1957 Werk Gendorf Burgkirchen, D-84508 Germany

Type IV per SAE AMS 1428

OCTAGON MAX FLIGHT TYPE IV

Octagon Process, Inc. 596 River Road Edgewater, New Jersey 07020

Type IV per SAE AMS 1428

UCAR ULTRA+

Union Carbide Customer Center 10235 W. Little York Road Houston, TX 77040

Type IV per SAE AMS 1428

Clariant Safewing MP IV 2001

Clariant Corporation 625 E. Catawba Ave. Mount Holly, NC 28120

R-134a

Obtain Locally

(116) Air-Conditioning Refrigerant

Page 14 Nov 1/11

SUPPLIER

91-00-00

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Table 1 Consumable Materials (Continued) MATERIAL (117) Air-Conditioning Oil

SPECIFICATION

SUPPLIER

Ester Oil

ICI Americas, Inc. P.O. Box 2376 Concord Pike and Murphy Road Wilmington, Delaware 19850

(118) Air-Conditioning System Cleaner

AC Flush

Castrol, Inc. Specialty Products Division 16715 Von Karman Avenue Irvine, California 92714

(119) Erosion Tape

8672

Minnesota Mining and Manufacturing Company 3M Center St. Paul, Minnesota 55144

(120) Sealer

Oyltite Stik

LA-CO Industries, Inc. Markal Company 250 N. Washtenaw Avenue Chicago, Illinois 60612

(122) Patch, Fuel Repair

Click-Patch

Courtaulds Aerospace, Inc. Semco Division P.O. Box 1800 5454 San Fernando Road Glendale, California 91209

(123) Sealer

Glyptol 1201

General Electric Company 3135 Easton Turnpike Fairfield, Connecticut 06431

(124) Sealant

PR1425

Courtaulds Aerospace, Inc. 5454 San Fernando Road P.O. Box 1800 Glendale, California 91209

(121) Item Deleted

(125) Solvent

RL 100s

PRODUCTS

(Refer to Item 120).

(Refer to item 30.)

(126) Deleted (127) Deleted (128) Aluminum Tubing

(129) Flouro-Lite Refrigerant Leak Detector Dye (R-134a System)

3/16 O.D. X 0.028 or 0.035 wall per Federal Spec WW-T-700/4 or WW-T-700/6

Obtain Locally

TP-3840

Tracer Products 956 Brush Hollow Rd. Westbury, NY 11590

91-00-00

Page 15 Nov 1/11

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Table 1 Consumable Materials (Continued) MATERIAL

SPECIFICATION

PRODUCTS

SUPPLIER

(130) Flouro-Lite Refrigerant Leak Detector Dye (R-12 System)

TP-3830

Tracer Products 956 Brush Hollow Rd. Westbury, NY 11590

(131) Glow Away Dye Cleaner/Remover

TP-9000

Tracer Products 956 Brush Hollow Rd. Westbury, NY 11590

(132) Red Refrigerant Leak Detector Dye (R-134a System)

Trace2

Highside Chemicals, Inc. 10-12 Colfax Ave. Clifton, NJ 07013

(133) Urethane Paint

Matterhorn White No. 6160

U.S. Paint, Lacquer and Chemical Co. St. Louis, MO

(134) Aluminum Oxide Cloth (240-grit)

P-C-451 Type 1

Obtain Locally

(135) Aluminum Oxide Cloth (320-grit)

P-C-451 Type 1

Obtain Locally

(136) Masking tape

MIL-T-23397

Obtain Locally

(137) Coating

MIL-C-81706 Class 1A Form III

Turco Products, Inc. 7300 Bolsa Ave. Westminster, CA 92684

(138) Cleanser

MIL-C-85570 Type II

Turco Products, Inc. 2000 Market St. Philadelphia, PA 19103

(139) Tie Wraps

MS33367-4-9

Obtain Locally

(140) Corrosion Inhibitive Compound

MIL-C-16173

Obtain Locally

(141) Oven Cleaner

Zep Oven Brite (PWC11-060)

(142) Fluorescent Liquid Penetrant

MIL-I-25135

(143) Tape, Conductive Adhesive Foil

MIL-T-47012

Page 16 Nov 1/11

91-00-00

Zep Manufacturing Co. of Canada 600 Lepine Avenue Dorval Quebec Canada H9P 1G2

NF/ZC-7B Zyglo Magnaflux Corp. Cleaner, 953 Westgate Drive - Suite 109 ZL-22A Zyglo St. Paul, Minnesota 55114 Penetrant, ZP-9 Zyglo Developer Minnesota Mining & Mfg. Co. 3M Center St. Paul, Mn. 55101

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Table 1 Consumable Materials (Continued) MATERIAL

SPECIFICATION

PRODUCTS

SUPPLIER

(144) Sealer

Technit 72-0002

Technical Wire Products Inc., 129 Dermondy St., Crawford, NJ. 07016

(145) Adhesive

Silastic 140

Dow Corning 3901 S. Saginaw Rd. Midland, MI. 48641

(146) Copper Foil Tape

3M1181

Minnesota Mining & Mfg. Co. 3M Center St. Paul, Mn. 55101

(147) Primer

SS-4179

General Electric Co. Silicone Products Waterford, NY 12188

(148) Electronic Grade Silicone rubber

RTV-162

General Electric Co. Silicone Products Waterford, NY 12188

(149) Cleanser

Zud

Obtain Locally

(150) Scrubbing Pad

Scotchbrite

Obtain Locally

(151) Primer

PR-142

PRC DeSoto International, Inc. 5454 San Fernando Road P.O. Box 1800 Glendale, California 91209

(152) Zephiran Chloride

00-2572

AVOX Systems (Formerly Scott Aviation) 225 Erie Street Lancaster NY. 14086-9502

(153) Neo-Novacite

00-736

AVOX Systems (Formerly Scott Aviation) 225 Erie Street Lancaster NY. 14086-9502

(154) Access Door Sealant

PR-1428

PRC DeSoto International, Inc. 5454 San Fernando Rd. P.O. Box 1800 Glendale, Ca. 91209

(155) Speed Tape

3M 345

Minnesota Mining & Mfg. Co. St. Paul, MN 55144

(156) Welding rod (Inconel)

Inconel 625

Hunington Alloys, Inc. 11750 Chesterdale Rd. Atkinson Square Cincinnati, OH, 45246

Type 347

Obtain locally

(157) Welding rod (Stainless Steel)

AWS A5.9 Class ER347

91-00-00

Page 17 Nov 1/11

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Table 1 Consumable Materials (Continued) MATERIAL

SPECIFICATION

PRODUCTS

SUPPLIER

(158) Sealer

PR-700

PRC DeSoto International, Inc. 5454 San Fernando Rd P.O. Box 1800 Glendale, Ca. 91209

(159) Cement

Bostic 1096M

United Shoe Machinery Corp. B.B. Chemical Div. 784 Memorial Dr. Cambridge, MA 02138

(160) Welding rod - Incoloy

Incoloy 82

Hunington Alloys, Inc. 11750 Chesterdale Rd. Atkinson Square Cincinnati, OH 45246

(161) Thread lock

Loctite 222

Loctite Corp., 705 W. Mountain Rd., Newington, CT 06111

(162) Sealant

Uralite 3149

Hawker Beechcraft Corporation Wichita, Kansas

(163) Sealant

LJF801A - 1/2

Le Joint Francais Bezons, France

(164) Abrasion Resistance Film

8591

Minnesota Mining & Mfg. Co., St. Paul MN

(165) Sealer

PR-1826

PRC DeSoto International, Inc. 5454 San Fernando Rd. P.O. Box 1800 Glendale, Ca. 91209

PR-1440B 1/2

PRC DeSoto International, Inc. 5454 San Fernando Road P.O. Box 1800 Glendale, California 91209

A-400

Silpak Inc. 169 Atlantic Street Pomona, California 91768

PTFE

Obtain locally

(169) Grease

MS-122XD (A Krytox lubricant) or MS-122DF

Miller-Stephenson Chemical Company, INC. 6348 Oakton St. Morton Grove, IL 60053

(170) Adhesive

Epibond 104

M and T Chemicals Inc. Furane Products Div. 5121 San Fernando Road West Los Angeles, CA 90039

(166) Sealer

AMS-S-8802

(167) Cabin Door Seal Adhesive (168) Tape, Anti-Seize

Page 18 Nov 1/11

A-A-58092

91-00-00

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Table 1 Consumable Materials (Continued) MATERIAL

SPECIFICATION

PRODUCTS

SUPPLIER

(171) Oil, Air Cycle Machine

Exxon 2380

Exxon Company, U.S.A. P.O. Box 2180 800 Bell Street Houston, Texas 77252

(172) Pin Hole Repair Product

74-451-AE

B.F. Goodrich Company P.O. Box 5471 Akron, Ohio 44313

(173) Emery Cloth

180-grit

Obtain Locally

(174) Emery Cloth

320-grit

Obtain Locally

(175) Filler

Devcon WR-1121

Devcon Corp., 59 Endicott St. Danvers, MA 01923

(176) Filler

Devcon WR-1141

Devcon Corp., 59 Endicott St. Danvers, MA 01923

(177) AgeMaster

Goodrich, 2730 West Tyvola Rd, Charolette, NC 28217

(178) Safety Wire

MS20995C32

0.032 - Corrosion Resistant

Obtain locally

(179) Safety Wire

MS9226-03

0.025 - Corrosion and Heat Resistant

Obtain locally

(180) Safety Wire

MS20995CU20

0.020 - Copper

Obtain locally

(181) Safety Wire

MS20995-020

0.020 - Corrosion and Heat Resistant

Obtain locally

(182) Tape 2-inch wide double-faced

No. 445

Minnesota Mining & Mfg. Co., St. Paul MN

(183) Adhesive

EC-847

Minnesota Mining & Mfg. Co., St. Paul MN

(184) Anti-skid Walkway Coating

SP 1593 Gray

Minnesota Mining & Mfg. Co., St. Paul MN

(185) Adhesive

RTV 737

Dow Corning 3901 S. Saginaw Road Midland, Michigan 48602

(186) Adhesive

5776-A/B

Ciba-Geigy Corporation 5121 San Fernando Rd. West Los Angeles, CA

(187) Aluminum Speed Tape

3M 425

Minnesota Mining & Mfg. Co. St. Paul, MN 55144

91-00-00

Page 19 Nov 1/11

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Table 1 Consumable Materials (Continued) MATERIAL

SPECIFICATION

PRODUCTS

SUPPLIER

(188) Sealer

RTV 106

General Electric OC., Silicone Products Dept., Hudson River Road, Waterford, NY 12188

(189) Epoxy

EA 9394

Union Chemicals Division, Union Oil Co. of California, 2100 E. 37th N., Wichita, KS 67208

(190) Silver-Filled Conductive Adhesive

610-1016

Dayton/Grainger P.O. Box 350550 Ft. Lauderdale, FL 33335

Cheese Cloth/ Rymple Cloth

American Fiber & Finishing 225 N. Depot St. Albemarle, NC 28001

(192) Grease

Molykote #33 Light Extreme Low Temperature or Dow Corning #33 Light Extreme Low Temperature

Dow Corning, S. Saginaw Rd., Midland, MI 48641

(193) Adhesive

RTV-157

General Electric Co. Silicone Plastics Group P.O. Box 12180 Waterford, NY 12188

(194) Retaining Compound

609

Loctite Corporation, 705 North Mountain Road, Newington, CT 06111

(191) Wiping Cloth/ Cheese cloth

Page 20 Nov 1/11

AMS 3819 Type 1 grade A or Type 2 grade A

91-00-00

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

B. Sealing Materials Table 2 Sealing Materials ITEM

APPLICATION

MATERIAL

REMARKS NOTE

Because the Model 1900 Airliner is a pressurized airplane, sealing the skin, bulkhead seams, windows, and doors is of prime importance. Control cables and torque shafts have removable rubber seals. When making a structural repair or modification which creates a break in the pressure vessel, the mating surfaces must be sealed with proper sealer described in Table 2. All components piercing the pressure vessel or attached to it must be sealed with the sealers described in Table 2. To assure effective bonding of the sealers, be sure to clean all mating surfaces, mating parts and rubber seals thoroughly. 1. Removing Old Methyl Propyl Ketone (MPK), Use Methyl Propyl Ketone to remove EC750, Sealer TT-T-548 Toluol EC847 and A4000. Use Toluol to remove EC776. All other sealers must be scraped off. TT-T-548 Toluol, 2. Cleaning Metal Naphtha is preferable since it will not damage Surface For Methyl Propyl Ketone (MPK), painted surfaces or plastic windows. When using Sealing TT-N-95 Naphtha, acetone or Methyl Propyl Ketone, remember that TT-I-735 Isopropyl Alcohol, they will remove paint and craze acrylic plastic. O-M-232D Methyl Alcohol, MIL-T-81533A Chlorothane 3. Cleaning Cabin TT-T-548 Toluol, Clean all surfaces to be bonded and if possible, Door and Methyl Propyl Ketone (MPK) roughen the surfaces with a wire brush or sandpaper. Escape Hatch Rubber Seal 4. Access Doors PR-1428 Replace sealer as necessary each time pressure doors are removed. 5. Escape Hatch EC847 Fit the rubber seal, then coat the seal with Seals Adhesive adhesive and allow to dry. Reactivate the adhesive with a cloth dampened in methyl propyl ketone and carefully install the seal. 6. Cabin Door Seal (Refer to item 24) Adhesive 7. Seal Inflation Silastic 140 Roughen surfaces with sandpaper; clean with Tube Methyl Propyl Ketone. 8. Windows Devcon F Tape width should extend equally on both sides of panes. Fill windshield frame gaps with Devcon F. 9. Skin Lap Joints *PR-1440B 1/2, Use 3M® 2084H only when an aluminum finish is ® 2084H required. Form a bead 1/8 to 3/16-inch high. 2 or 4 3M 10. Structural Fillet *PR-1440B 1/2, 2 or 4 Apply a bead at least three times the thickness, Seal packing the sealant around rubber seal. 11. Trim Tab Cable *PR-1440B 1/2, 2 or 4 Apply an 1/8-inch fillet of sealant around rubber Seals seal. 12. Gaps and Voids PR-1440B 1/2 Gaps and voids larger than 3/8 inch in two directions must have a mechanical closure.

91-00-00

Page 21 Nov 1/11

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Table 2 Sealing Materials (Continued) ITEM 13.

APPLICATION Rivets

14.

Unused Drill Holes Plumbing Fittings Electrical Fittings Bleed Air Ducts

15. 16. 17. 18.

19. 20. 21. 22. 23. 24.

MATERIAL EC776

PR-1440B 1/2 PR-1440B 1/2 PR-1440B 1/2, EC776 RTV-108 or RTV-732 and Silastic 140 Elmer's Glue

REMARKS Brush an area at least double the size of the rivet head or butt, working the sealer with the brush to assure that the rivet is completely sealed. Plug any unused drill holes with rivets or seal with PR-1440B 1/2. Seal around plumbing fittings with PR-1440B 1/2. Seal all electrical plugs with PR-1440B 1/2 and seal the mounting screws with EC776. Seal all electrical plugs with JFM 1239 and seal the mounting screws with EC776. Coat screw threads with glue and install in the plastic while the glue is wet. Wash off excess glue with Naphtha.

Storm Window Hinge and Latch Attaching Screws Windshield PR-1440 Apply to exposed portion of windshield interlayer Interlayer immediately inside retainer strips. Wet Wing PR-1440B1/2 Apply around anchor nuts. Access Plates Flap Flexible PR-1440B1/2 Coat exposed portions of the rubber seal Driveshaft overlapping the shaft assembly and root rib. Integral Fuel PR-1826 Sealing of leak areas, seams, anchor nuts and Cell Sealer repair areas. Firezone Cables PR-700 Potting of Firezone Cables in their connectors. Cabin Door Seal Dapcotac 3300, Silpak A-400 or Adhesive RTV-157 * Use PR-1440B 1/2 when assembly time is 30 minutes or less. Use PR-1440B2 or 4 when assembly time exceeds 30 minutes.

Page 22 Nov 1/11

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

C. Torquing for Coarse Thread Series Bolt-Nut Combination Table 3 Torquing for Coarse Thread Series Bolt-Nut Combination NOTE The following torque values may be used as a guide when specific torque values are not specified within this manual. The torque values apply to class 3 threads. The threads are cadmium-plated and nonlubricated. TORQUE LIMITS RECOMMENDED FOR INSTALLATION (INCH-POUNDS) Size

MAXIMUM ALLOWABLE TIGHTENING TORQUE (INCH-POUNDS)

Column 1

Column 2

Column 3

Column 4

MS20365 & AN310 Nuts (Tension)

MS20364 & AN320 Nuts (Shear)

MS20365 & AN310 Nuts (Tension)

MS20364 & AN320 Nuts (Shear)

8-32

12-15

7-9

20

12

10-24

20-25

12-15

35

21

1/4-20

40-50

25-30

75

45

5/16-18

80-90

48-55

160

100

3/8-16

160-185

95-110

275

170

7/16-14

235-255

140-155

475

280

1/2-13

400-480

240-290

880

520

9/16-12

500-700

300-420

1100

650

5/8-11

700-900

420-540

1500

900

3/4-10

1150-1600

700-950

2500

1500

7/8-9

2200-3000

1300-1800

4600

2700

1-8

3700-5000

2200-3000

7600

4500

1-1/8-8

5500-6500

3300-4000

12000

7200

1-1/4-8

6500-8000

4000-5000

16000

10000

91-00-00

Page 23 Nov 1/11

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

D. Torquing for Fine Thread Series Bolt-Nut Combination Table 4 Torquing for Fine Thread Series Bolt-Nut Combination NOTE The following torque values may be used as a guide when specific torque values are not specified within this manual. The torque values apply to class 3 threads. The threads are cadmium-plated and nonlubricated. TORQUE LIMITS RECOMMENDED FOR INSTALLATION (INCH-POUNDS) Size

Column 1

Column 2

MAXIMUM ALLOWABLE TIGHTENING TORQUE (INCH-POUNDS) Column 3

Column 4

MS20365 & AN310 Nuts (Tension)

MS21042 Dry-Film Lubed (Tension)

MS20364 & AN320 Nuts (Shear)

MS21245 Dry-Film Lubed (Shear)

MS20365 & AN310 Nuts (Tension)

MS21042 Dry-Film Lubed (Tension)

MS20364 & AN320 Nuts (Shear)

MS21245 Dry-Film Lubed (Shear)

8-36

12-15

*

7-9

----

20

*

12

----

10-32

20-25

15-19

12-15

----

40

30

25

----

1/4-28

50-70

37-47

30-40

----

100

70

60

----

5/16-24

100-140

56-78

60-85

----

225

135

140

----

3/8-24

160-190

72-108

95-110

----

390

200

240

----

7/16-20

450-500

----

270-300

----

840

----

500

----

1/2-20

480-690

----

290-410

210-230

1100

----

660

415

9/16-18

800-1000

----

480-600

310-430

1600

----

960

660

5/8-18

1100-1300

----

660-780

485-605

2400

----

1400

1060

3/4-16

2300-2500

----

1300-1500 1090-1250

5000

----

3000

2500

7/8-14

2500-3000

----

1800-2400 1640-2100

7000

----

4200

3740

* This is a course thread nut (8-32) with recommended torque limits of 9 to 11 inch-pounds. The maximum allowable tightening torque is 14 inch-pounds.

Page 24 Nov 1/11

91-00-00

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

E. Flare Fitting Torque Table 5 Flare Fitting Torque NOTE Refer to Table 6 for thread lubricants. TORQUE LIMITS RECOMMENDED (INCH-POUNDS) Hose Assemblies1 Tubing OD Inches

Aluminum-Alloy Tubing

Steel Tubing

Aluminum-Alloy Tubing (Double Flare) For Use On Oxygen Lines Only

MIN

MAX

MIN

MAX

MIN

MAX

MIN

MAX

1/8

20

30

75

85

* / **

* / **

---

---

3/16

25

35

95

105

* / **

* / **

---

---

1/4

50

65

135

150

* / **

* / **

---

---

5/16

70

90

170

200

* / **

* / **

100

125

3/8

110

130

270

300

* / **

* / **

200

250

1/2

230

260

450

500

* / **

* / **

300

400

5/82

330

360

650

700

* / **

* / **

---

---

3/4

460

500

900

1000

* / **

* / **

---

---

1

500

700

1200

1400

* / **

* / **

---

---

1-1/4

800

900

1520

1680

* / **

* / **

---

---

1-1/2

800

900

1900

2100

* / **

* / **

---

---

1-3/4

---

---

---

---

---

---

---

---

2

1800

2000

2660

2940

* / **

* / **

---

---

* When the hose fitting (nipple and nut) is aluminum, the min/max values for aluminum alloy tubing shall apply. ** When the hose fitting (nipple and nut) is steel, the min/max values for steel tubing shall apply. 1

MIL-H-8790, MIL-H-8795, MIL-H-25579, MIL-H-26666, MIL-H-38360, MIL-H-38390

2

Fuel system H fittings in the wheel wells are installed on 5/8-inch OD tubing and should be torqued 60 (min) to 72 (max) inch-pounds.

91-00-00

Page 25 Nov 1/11

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

F. Thread Lubricants for Fluid-line Fittings Table 6 Thread Lubricants For Fluid-line Fittings Type of Line

Functional Fluid

Brass, Steel and Aluminum Fittings

(1) Fuel and Fuel Pressure Line Fuel

Braycote 236 (VV-P-236)

(2) Lubricating Oil and Oil Pressure Line

Lubricating Oil

Braycote 236 (VV-P-236) or MIL-G-6032 Lubricating Grease (Gasoline and Oil Resistant)

Air (ambient)

None

(3) Automatic Pilot (a) (straight threads) (b) (tapered threads)

3M Tape #48 or #547 PTFE Tape (A-A-58092, Type III, Size 1 or 2)

(4) Pressurization Control (a) Straight threads

Breathable Air (ambient)

(b) Tapered threads

None 3M Tape #48 or #547 PTFE Tape (A-A-58092, Type III, Size 1 or 2)

(5) Pitot System (a) Straight threads

Air (ambient)

(b) Tapered threads

None Loctite PST 59231 Pipe Sealant

(6) Fire Extinguisher System

Trifluoro Bromo Methane

Loctite PST 59231 Pipe Sealant

(7) Air Conditioning System

Freon (Possible Refrigeration Oil)

Oil of System

Oxygen

Krytox 240AC Grease (MIL-G-27617 Type III)

(8) Oxygen System (a) Straight threads (b) Tapered threads

3M Tape #48 or #547 PTFE Tape (A-A-58092, Type III, Size 1 or 2)

(9) Bleed Air System (a) CRES (straight and tapered threads)

Air (Max. 750°F)

Dow Corning 77 (M-77)

(b) AL (straight and tapered threads)

Air

Loctite PST 59231 Pipe Sealant

(10) Hydraulic System (a) Straight threads

Hydraulic Fluid

(b) Tapered threads

Fluid of System Fluid of System or Loctite 545

(11) Vacuum

3M Tape #48 or #547 PTFE Tape (A-A-58092, Type III, Size 1 or 2) or Dow Corning 111 Valve & Lubricant Sealant

(12) Deicer

Loctite PST 59231 Pipe Sealant

(13) Gyro (Edo Air)

Page 26 Nov 1/11

Air

91-00-00

3M Tape #48 or #547 PTFE Tape (A-A-58092, Type III, Size 1 or 2)

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

G. Special Tools and Equipment Table 7 Special Tools and Equipment TOOL NAME

PART NUMBER

SUPPLIER

USE

(1) Ultra Violet Light

ZB-100

Magnaflux Corp. 953 Westgate Drive - Suite 109 St. Paul, MN 55114

Inspect for cracks in airframe

(2) Eddy Current Tester

ED520

Magnaflux Corp. 953 Westgate Drive - Suite 109 St. Paul, MN 55114

Perform eddy current test

(3) Pitot/Static Leak Tester

TK1783-1/939

Hawker Beechcraft Corp. 9709 E. Central Wichita, KS. 67201

Leak test Pitot/Static system

or Model 1811D or equivalent

Barfield Inc. 4101 N.W. 29th Miami, Fl 33142

Leak test Pitot/Static system

TK1783-3/939-1 or equivalent

Hawker Beechcraft Corp. 9709 E. Central Wichita, KS. 67201 For local manufacture refer to SPECIAL TOOLS.

Leak test Pitot/Static system

(4) Pitot/Static Adapter

(5) Oxygen Test Harness TK1738-5/939-2 or equivalent

Hawker Beechcraft Corp. Check oxygen system 9709 E. Central Wichita, KS. 67201 For local manufacture refer to the Chapter 35-00-00 OXYGEN SYSTEM TEST EQUIPMENT (Figure 201, Sheet 2) illustration in the Maintenance Practices section.

(6) Pressure Gauge (0-160 psi)

TK1738-6/939 or equivalent

Hawker Beechcraft Corp. 9709 E. Central Wichita, KS. 67201

Check oxygen system

(7) Jumper Hose and Adapter

Size 6

For local manufacture refer to the Chapter 35-00-00 ALTITUDE COMPENSATOR TEST EQUIPMENT illustration in the Maintenance Practices section.

Check oxygen system

(8) Vibration Analyzer

Model 8500, Model 8350, or Model 192A

Chadwick-Helmuth Aviation Products 4601 N. Arden Dr. El Monte, CA 91731

Propeller balancing

91-00-00

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Table 7 Special Tools and Equipment (Continued) TOOL NAME

PART NUMBER

SUPPLIER

USE

(9) Deleted, moved to Chapter 27-00-00. (10) Deleted, moved to Chapter 27-00-02. (11) Deleted, moved to Chapter 27-00-02. (12) Deleted, moved to Chapter 27-00-02. (13) Deleted, moved to Chapter 27-00-02. (14) Fuel Quantity Test Set

TK 2129/935

Hawker Beechcraft Corp. 9709 E. Central Wichita, KS. 67201

Fuel quantity indicating checks.

(15) Probe Selector Unit

114-389001-935

Hawker Beechcraft Corp. 9709 E. Central Wichita, KS. 67201

Fuel quantity indicating checks.

(16) Probe Adapter

101-00814

Barfield Inc. 4101 N.W. 29th Miami, Fl 33142

Bench checking fuel quantity probes.

(17) Tension Adjustment Wrench

02314-0023-0001

BF Goodrich Aircraft Sensors Div. 14300 Judicial Rd. Burnsville, MN 55306

Set windshield wiper spring tension.

(18) Deleted, moved to Chapter 27-00-00. (19) Deleted, moved to Chapter 27-00-00. (20) Deleted, moved to Chapter 27-00-00. (21) Deleted, moved to Chapter 27-00-02. (22) Deleted, moved to Chapter 27-00-02. (23) Deleted, moved to Chapter 27-00-02. (24) Deleted, moved to Chapter 27-00-00. (25) Deleted, moved to Chapter 27-00-00.

Page 28 Nov 1/11

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL Table 7 Special Tools and Equipment (Continued) TOOL NAME

PART NUMBER

SUPPLIER

USE

(26) Torque Adapter (Figure 1, Sheet 4 of 10)

10113

Hawker Beechcraft Corp. 9709 E. Central Wichita, KS. 67201

Apply torque to propeller mounting bolts.

(27) Torque Adapter 3

AST-2877

Hartzell Propeller Inc. Piqua, OH 45356

Apply torque to propeller mounting bolts.

(28) Megohmmeter

2471 F

Barfield Instrument Corporation 1478 Central Avenue, Atlanta, GA 39344

Check Static Dischargers.

(29) Control Surface Gust Lock Assembly

101-590016-5

Hawker Beechcraft Corp. 9709 E. Central Wichita, KS. 67201

Lock flight controls when parked.

(30) Main Landing Gear Switch Test Box (Figure 1, Sheet 10 of 10)

P/N TK1763-7/935

Hawker Beechcraft Corp. 9709 E. Central Wichita, KS. 67201

To check the main landing gear switches

(31) Milliohmmeter

580A

Keithley Instruments Inc. 28775 Aurora Road Cleveland, OH. 44139

Check static discharger base mounts.

91-00-00

Page 29 Nov 1/11

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 1 (Sheet 1 of 10) Special Tools

Page 30 Nov 1/11

91-00-00

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 1 (Sheet 2 of 10) Special Tools

91-00-00

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MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 1 (Sheet 3 of 10) Special Tools

Page 32 Nov 1/11

91-00-00

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 1 (Sheet 4 of 10) Special Tools

91-00-00

Page 33 Nov 1/11

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 1 (Sheet 5 of 10) Special Tools

Page 34 Nov 1/11

91-00-00

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 1 (Sheet 6 of 10) Special Tools

91-00-00

Page 35 Nov 1/11

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

A 36 INCHES MINIMUM 60 INCHES MINIMUM

8 INCHES MINIMUM

A

30°

DETAIL

A WHEEL RAMP

A RAMP MAY BE FABRICATED LOCALLY BY ATTACHING ONE 36 INCHES (MINIMUM) PIECE AND ONE 60 INCHES (MINIMUM) PIECE OF HIGH QUALITY LUMBER TOGETHER. BOTH PIECES SHOULD BE AT LEAST 8 INCHES WIDE, WITH A COMBINED THICKNESS BETWEEN 3 AND 4 INCHES AND HAVE A 30 DEGREE (MAXIMUM) TAPER RELATIVE TO THE GROUND ON EACH END.

UC91B 001873AB.AI

Figure 1 (Sheet 7 of 10) Special Tools

Page 36 Nov 1/11

91-00-00

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 1 (Sheet 8 of 10) Special Tools

91-00-00

Page 37 Nov 1/11

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 1 (Sheet 9 of 10) Special Tools

Page 38 Nov 1/11

91-00-00

MODEL 1900/1900C AIRLINER MAINTENANCE MANUAL

Figure 1 (Sheet 10 of 10) Special Tools

91-00-00

Page 39 Nov 1/11