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CITATION MUSTANG PILOT TRAINING MANUAL

“The best safety device in any aircraft is a well-trained crew.”™

CITATION MUSTANG PILOT TRAINING MANUAL

FIRST EDITION

REVISION 1.1

REVISION 1.1 FlightSafety International, Inc. Marine Air Terminal, LaGuardia Airport Flushing, New York 11371 (718) 565-4100 www.FlightSafety.com

F O R T R A I N I N G P U R P O S E S O N LY

NOTICE The material contained in this training manual is based on information obtained from the aircraft manufacturer’s Airplane Flight Manual and Maintenance Manuals. It is to be used for familiarization and training purposes only. At the time of printing it contained then-current information. In the event of conflict between data provided herein and that in publications issued by the manufacturer or the FAA, that of the manufacturer or the FAA shall take precedence. We at FlightSafety want you to have the best training possible. We welcome any suggestions you might have for improving this manual or any other aspect of our training p ­ rogram.

F O R T R A I N I N G P U R P O S E S O N LY

Courses for the Citation Mustang aircraft are taught at the following FlightSafety Learning Centers: Citation Learning Center FlightSafety International 1851 Airport Road P.O. Box 12323 Wichita, Kansas 67277 Phone: (316) 220-3100 Toll-Free: (800) 488-3214 Fax: (316) 220-3134 Orlando Learning Center 4105 Bear Road Orlando, Florida 32827-5001 (321) 281-3200 (800) 205-7494 FAX (321) 281-3299 Farnborough Training Center Farnborough Airport Farnborough, Hampshire GU14 6XA United Kingdom +44 (0) 1252 554 500 Fax: +44 (0) 1252 554 599

Copyright © 2014 FlightSafety International, Inc. Unauthorized reproduction or distribution is prohibited. All rights reserved.

INSERT LATEST REVISED PAGES, DESTROY SUPERSEDED PAGES LIST OF EFFECTIVE PAGES Dates of issue for original and changed pages are: Original......................... 0.0.............FEB 2007 Revision........................ 1.0............ DEC 2009 Revision........................ 1.1...........JULY 2014 NOTE: For printing purposes, revision numbers in footers occur at the bottom of every page that has changed in any way (grammatical or typographical revisions, reflow of pages, and other changes that do not necessarily affect the meaning of the manual). THIS PUBLICATION CONSISTS OF THE FOLLOWING: Page *Revision Page *Revision No. No. No. No. 5-i—5-iv.................................................. 0.0 Cover...................................................... 1.1 Copyright................................................ 1.1 5-1—5-2................................................. 1.0 i—vi......................................................... 1.1 5-3—5-3................................................. 0.0 1-i—1-iv.................................................. 0.0 5-4—5-8................................................. 1.0 1-1—1-5................................................. 0.0 5-9—5-9................................................. 1.1 1-6—1-6................................................. 1.0 5-10—5-10............................................. 1.0 1-7—1-10............................................... 0.0 5-11—5-12............................................. 0.0 1-11—1-11............................................. 1.1 6-i—6-ii................................................... 0.0 1-12—1-12............................................. 0.0 7-i—7-vi.................................................. 0.0 2-i—2-iv.................................................. 0.0 7-1—7-2................................................. 0.0 2-1—2-2................................................. 0.0 7-3—7-3................................................. 1.0 2-3—2-4................................................. 1.0 7-4—7-6................................................. 0.0 2-5—2-6................................................. 0.0 7-7—7-7................................................. 1.0 2-7—2-7................................................. 1.0 7-8—7-10............................................... 0.0 2-8—2-9................................................. 0.0 7-11—7-11............................................. 1.0 2-10—2-12............................................. 1.0 7-12—7-15............................................. 0.0 2-13—2-15............................................. 1.1 7-16—7-17............................................. 1.0 2-16—2-16............................................. 1.0 7-18—7-18............................................. 0.0 3-i—3-iv.................................................. 0.0 7-19—7-19............................................. 1.1 3-1—3-1................................................. 1.0 7-20—7-21............................................. 1.0 3-2—3-2................................................. 0.0 7-22—7-22............................................. 0.0 3-3—3-5................................................. 1.0 8-i—8-vi.................................................. 0.0 3-6—3-8................................................. 0.0 8-1—8-1................................................. 0.0 4-i—4-ii................................................... 0.0 8-2—8-2................................................. 1.0 4-1—4-4................................................. 1.0 8-5—8-6................................................. 1.1 4-5—4-5................................................. 0.0 8-7—8-7................................................. 1.0 4-6—4-6................................................. 1.0 8-8—8-8................................................. 1.0 4-7—4-7................................................. 0.0 9-i—9-ii................................................... 0.0 4-8—4-8................................................. 1.0 9-1—9-1................................................. 0.0 4-9—4-9................................................. 1.1 9-2—9-2................................................. 1.0 4-10—4-10............................................. 0.0 9-3—9-6................................................. 0.0 *Zero in this column indicates an original page. Revision 1.1

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CONTENTS Chapter 1

AIRCRAFT GENERAL

Chapter 2

ELECTRICAL POWER SYSTEMS

Chapter 3

LIGHTING

Chapter 4

MASTER WARNING SYSTEM

Chapter 5

FUEL SYSTEM

Chapter 6

AUXILIARY POWER SYSTEM

Chapter 7

POWERPLANT

Chapter 8

FIRE PROTECTION

Chapter 9

PNEUMATICS

Chapter 10

ICE AND RAIN PROTECTION

Chapter 11

AIR CONDITIONING

Chapter 12

PRESSURIZATION

Chapter 13

HYDRAULIC POWER SYSTEM

Chapter 14

LANDING GEAR AND BRAKES

Chapter 15

FLIGHT CONTROLS

Chapter 16

AVIONICS

Chapter 17

OXYGEN SYSTEM

Chapter 18

MANEUVERS AND PROCEDURES

Chapter 19

WEIGHT AND BALANCE

Chapter 20

FLIGHT PLANNING AND PERFORMANCE

Chapter 21

CREW RESOURCE MANAGEMENT

WALKAROUND APPENDIX A

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CHAPTER 1 AIRCRAFT GENERAL CONTENTS INTRODUCTION.................................................................................................................. 1-1 Dimensions...................................................................................................................... 1-2 Weight Limitations.......................................................................................................... 1-2 STRUCTURES....................................................................................................................... 1-2 Entrance Door................................................................................................................. 1-4 Emergency Exit............................................................................................................... 1-6 Cabin............................................................................................................................... 1-7 Flight Compartment........................................................................................................ 1-7 Tailcone Compartment.................................................................................................... 1-7 Wing................................................................................................................................ 1-7 Empennage...................................................................................................................... 1-7 Nose Section.................................................................................................................... 1-8 SYSTEMS.............................................................................................................................. 1-8 Electrical System............................................................................................................. 1-8 Fuel System..................................................................................................................... 1-9 Engines............................................................................................................................ 1-9 Ice Protection System...................................................................................................... 1-9 Hydraulic System............................................................................................................ 1-9 Flight Controls................................................................................................................. 1-9 Environmental System.................................................................................................... 1-9 Avionics......................................................................................................................... 1-10 PUBLICATIONS.................................................................................................................. 1-10

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SINGLE-PILOT OPERATION............................................................................................. 1-10 LIMITATIONS...................................................................................................................... 1-10 EMERGENCY/ABNORMAL.............................................................................................. 1-10 QUESTIONS........................................................................................................................ 1-12

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ILLUSTRATIONS Figure Title Page 1-1.

Cessna Citation Mustang......................................................................................... 1-2

1-2.

Citation Mustang Dimensions................................................................................. 1-3

1-3.

Braking Taxi Turning Distance................................................................................ 1-4

1-4.

Engine Hazard Areas............................................................................................... 1-4

1-5.

Entrance Door, Interior Handle, and Latch Release................................................ 1-5

1-6.

Hinged Panel............................................................................................................ 1-5

1-7.

Door Pin Indicator................................................................................................... 1-6

1-8.

Emergency Exit........................................................................................................ 1-6

1-9.

Tailcone Baggage Door........................................................................................... 1-7

1-10.

Wing Trailing Edge.................................................................................................. 1-7

1-11.

Stall Strips................................................................................................................ 1-8

1-12. Empennage.............................................................................................................. 1-8 1-13.

Nose Storage Compartment..................................................................................... 1-8

1-14.

Nose Baggage Light................................................................................................ 1-8

TABLES Table Title Page 1-1. CAS MESSAGES.................................................................................................. 1-11

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CHAPTER 1 AIRCRAFT GENERAL

INTRODUCTION This manual provides a description of the major airframe and engine systems in the Cessna Citation Mustang (Figure 1-1). The information contained herein is intended only as an instructional aid. This material does not supersede, nor is it meant to substitute for, any of the manufacturer’s maintenance or flight manuals. The material presented has been prepared from current design data. Chapter 1 covers the structural makeup of the airplane and gives an overview of the systems.

GENERAL The Citation Mustang is certified in accordance with Title 14 of the Code of Federal Regulations (14 CFR 23) Part 23, including day, night, visual flight rules (VFR), instrument flight rules (IFR), single pilot, and flight into known icing conditions. Takeoff and landing performance and other special condition certification requirements are similar to 14 CFR Part 25. The Mustang meets 14 CFR

Part 36 noise standards, and meets 14 CFR Part 34 fuel venting and exhaust emission standards. It combines systems simplicity with ease of access to reduce maintenance requirements. Low takeoff and landing speeds permit operation at small airports. Medium bypass turbofan engines contribute to overall operating efficiency and performance.

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CITATION MUSTANG PILOT TRAINING MANUAL

Figure 1-1.   Cessna Citation Mustang

The Citation Mustang is equipped with a Garmin G-1000 integrated avionics system. This threepanel display system integrates flight instruments, flight guidance, navigation, and communication systems. Also integrated into the avionics system are the master warning and master caution systems and hazard avoidance systems. The engine indication and crew alerting system (EICAS) is a two-column display on the left side of the center multifunction display (MFD). The crew alerting system (CAS) messages are displayed on the lower left of the MFD to alert the crew of system emergencies, abnormal situations, or changes in system operation. EICAS and CAS will be referred to often in the following chapters.

DIMENSIONS Figure 1-2 shows a three-view drawing of the Citation Mustang containing the approximate exterior and cabin dimensions.

WEIGHT LIMITATIONS • Maximum ramp weight ....... 8,730 pounds • Maximum takeoff weight ...... 8,645 pounds • Maximum landing weight ...... 8,000 pounds

1-2

STRUCTURES The Citation Mustang (Figure 1-1) is an all-metal, pressurized, low-wing monoplane with a swept T-tail. The interior has provisions for up to four passengers and two crewmembers. It has one cabin entry door and one emergency exit. The aircraft has baggage compartments in the nose and tail cone. Two pylon-mounted Pratt & Whitney PW615F turbofan engines are on the rear fuselage. Figure 1-3 shows braking taxi turning distance, and Figure 1-4 is a diagram of engine hazard areas. The aircraft has five doors: • Entrance door • Emergency exit (escape hatch) • Left nose baggage compartment door • Right nose baggage compartment door • Aft (tail cone) compartment door Each door (except the emergency exit) has a monitoring system, which provides a specific CAS message for that door if it is not properly closed. However, if the monitoring system for any door fails to pass a test on the ground, stops operating, or does not indicate normal operating condition, the CHECK DOORS CAS message appears.

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43.17 FT (13.0 M)

11.79 FT (3.6 M)

55.00 IN. (1,397 MM)

54.00 IN. (1,372 MM)

34.24 IN. (870 MM)

FORWARD DIVIDER (0.53 IN. THICK) FORWARD (13 MM THICK) PRESSURE BULKHEAD FS 144.00 IN. (3,658 MM)

FS 202.76 IN. (5,150 MM) 24.00 IN. (610 MM)

BAGGAGE DOORS

FS 321.00 IN. (8,153 MM)

AFT PRESSURE BULKHEAD 13.10 FT (3.99 M)

46.00 IN. 54.00 IN. (1,168 MM) (1,372 MM)

RUDDER TRIM TAB TAIL CONE ACCESS DOOR (L SIDE ONLY)

FLIGHT COMPARTMENT 58.76 IN. (1,493 MM)

PASSENGER COMPARTMENT 117.71 IN. (2,990 MM)

14.35 FT (4.37 M)

40.56 FT (12.36 M)

Figure 1-2.   Citation Mustang Dimensions

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WING-TIP LIGHT TO WING TIP LIGHT CURB TO CURB 27.32 FEET 54.97 FEET (8.33 M) (16.75 M)

11.79 FEET (3.59 M)

15.53 FEET (4.734 M)

Figure 1-3.   Braking Taxi Turning Distance

27 FEET (8 M)

18 FEET (5.50 M) DISTANCE METERS

DISTANCE FEET

0

9

18

27

36

45

0

30

60

90

120

150

Figure 1-4.   Engine Hazard Areas

ENTRANCE DOOR The cabin entrance door is on the forward left side of the fuselage (Figure 1-5). The entrance door opens outboard and forward. It is secured in the closed position with eight locking pins attached to a handle. The door can be opened from inside or outside of the airplane. The exterior handle can be secured with a key.

1-4

Ensure that the key is removed from the entrance door prior to flight to prevent possible ingestion of the key into an engine. An adjustable stop prevents the door from opening too far. Once the door is fully open, a hook locks the door into position. To unlatch the hook and let the door close, a release button inside the cabin (inside left of door opening) must be pushed (Figure 1-5). This lets the door move freely.

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Figure 1-5.   Entrance Door, Interior Handle, and Latch Release

CAUTION The locking pins will contact and damage the painted surface of the fuselage if an attempt is made to shut the door with the handle in the closed (up) position. The seal system operates passively as the cabin is pressurized. The seal engages and disengages when the door opens and closes. A hinged panel at the main cabin door threshold is used as a water barrier during ditching (Figure 1-6). It hinges up to prevent water from entering the aircraft, and enables the use of the entrance door as an exit during ditching.

WARNING Water barrier must be raised and latched into position prior to ditching.

Figure 1-6.   Hinged Panel

NOTE The water barrier is installed at the main cabin door threshold. Crew members should be familiar with its location and operation; and passengers should be briefed prior to flights over water.

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A monitoring system checks for safe condition of the door (closed, pins secure, door latched). There are eight view ports on the inside panel of the entry door to verify that the eight locking pins are in the closed position. With the door closed and handle latched, the pilot should be able to see the white and black indicators in each port. (Figure 1-7). Three proximity switches verify that the door is closed, pins secure, and door latched by sensing the position of targets on the closing mechanism of the door and signals the condition to the CAS: • T he proximity switch on the doorway surround structure senses the door after it is closed. • W hen the door handle is latched, this moves a bracket on the left and right side of the door outward pushing the eight locking pins into position. A proximity switch senses that this locking bracket has moved into the locked position (pilot must inspect the eight door pin indicators) (Figure 1-7).

EXTERIOR

• A s the handle catch is engaged, a proximity switch on the inner handle assembly senses a flag. If one of the proximity switches does not sense its target, the CABIN DOOR CAS message appears.

EMERGENCY EXIT A plug-type emergency exit (escape hatch) is on the aft right side of the cabin, above the wing. It opens inboard. The emergency exit door can be opened from outside or inside the airplane (Figure 1-8).

INTERIOR

Figure 1-8.   Emergency Exit

The D-shaped inner door handle is recessed behind a magnetic cover. The flush-mounted outer handle is located at the top of the door. The outer handle is not directly connected to the inner handle. The outer handle has a black indicator to show when the door is latched.

Figure 1-7.   Door Pin Indicator

1-6

Because no other provisions are provided for securing the escape hatch when the airplane is unattended, a safety pin with a REMOVE BEFORE FLIGHT streamer is placed on the inside of the hatch. The pilot must ensure this pin is removed prior to flight. The emergency exit hatch is not connected to the door warning circuit.

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CABIN The cabin extends from the forward pressure bulkhead to the aft pressure bulkhead and measures approximately 14 feet in length, 4.7 feet in width, and 4.5 feet in height. Figure 1-2 shows the interior arrangements and dimensions. The standard interior arrangement consists of two aft-facing and two forward-facing passenger seats. There is a toilet on the right side of the fuselage, abeam the cabin entry door. The toilet is not equipped with a safety belt and cannot be occupied during taxi, takeoff, or landing. The cabin area has dropout, constant-flow oxygen masks for emergency use. The cabin overhead panels contain individual air vent outlets and seat lighting for passenger comfort.

FLIGHT COMPARTMENT The airplane is equipped with dual controls, including control yokes, brakes, and rudder pedals at each crew seat. There are two adjustable seats with seat belts and shoulder harnesses.

Figure 1-9.   Tailcone Baggage Door

WING The wing assembly attaches to the bottom of the fuselage and is constructed of aluminum. Each wing is also a fuel tank. Electromechanical speedbrakes and flaps, and hydraulically actuated main landing gear are attached to each wing (Figure 1-10).

TAILCONE COMPARTMENT The tailcone compartment is an unpressurized area and contains major components of the environmental, electrical distribution, flight controls, and engine fire extinguishing systems. Access is through the tail cone baggage door on the left side of the fuselage below the engine. This door opens the tail cone baggage compartment (Figure 1-9), which holds 300 pounds.

Figure 1-10.   Wing Trailing Edge

An aileron fence is attached to the inboard side of each aileron.

The tailcone compartment door is secured at the aft side by mechanical latches and a key lock and is hinged at the left forward edge. The door is secured by a key lock, which is monitored by the CAS. The AFT DOOR CAS message appears if the door is unlocked.

The wing leading edges are deiced by inflatable deice boots, which are inflated by regulated engine bleed-air. Vortex generators and stall strips are attached to the leading edge boots (Figure 1-11).

A light switch on the right side of the door opening is powered from the battery bus and provides illumination of the tail cone area for preflight inspection purposes. If the manual switch is left on, a microswitch in the door track extinguishes the light when the door is closed.

The empennage consists of a vertical stabilizer with T-tail mounted horizontal stabilizers (Figure 1-12). The leading edges of the horizontal and vertical stabilizers are deiced by inflatable deice boots.

EMPENNAGE

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Figure 1-11.   Stall Strips

Figure 1-13.   Nose Storage Compartment

Figure 1-12.   Empennage

NOSE SECTION The nose section is an unpressurized storage area. Various hydraulics components, pneumatic bottles, oxygen bottle, fresh-air duct, and radar antenna are located in this compartment (Figure 1-13). The nose storage compartment holds up to 20-cubic feet (320-pounds) of baggage. It has two swing-up doors (left and right). Each door has a mechanical lock. Each door has a key-operated cam lock, forward pin latch, and two independent paddle latches. The pin latch shows orange when not latched. Each latch has a switch and indicates the latch position with the NOSE DOOR L-R CAS message. A manual light switch is in the compartment (Figure 1-14). If the manual light switch is left on, a microswitch at the left and right storage door assembly extinguishes the storage compartment light when the doors are closed.

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Figure 1-14.   Nose Baggage Light

An over-center gas spring on each door holds the door in the full open position until the door is closed manually. Ensure that the keys are removed from both nose compartment doors prior to flight to prevent possible ingestion of a key into an engine.

SYSTEMS ELECTRICAL SYSTEM The Mustang is an all-DC aircraft. The 28-VDC electrical power is supplied by two starter-generators and one 24-volt, 28 amp-hour sealed lead acid battery. An optional battery is a 24-volt, 28 amphour NiCad battery. An external power receptacle is below the right engine pylon.

FOR TRAINING PURPOSES ONLY

For convenience of pilot and passengers, two DC power outlets are provided in the cabin, powered by the DC system through a converter. One DC outlet is in the cabinet behind the copilot seat and the other is in the aft center console.

FUEL SYSTEM The fuel system has two distinct, identical halves. Each wing tank stores and supplies the fuel to its respective engine. Fuel transfer capability is provided. Fuel is heated through an oil-to-fuel heat exchanger (PRIST is not required).

ENGINES Two pylon-mounted Pratt & Whitney PW615F turbofan engines are on the rear fuselage, and each produces approximately 1,460 pounds of thrust (sea level ISA + 0°C). To improve automation and efficiency, the engines are controlled by dual-channel full authority digital engine controls (FADECs). Engines are started with electrical starter-generators, which are powered by the onboard battery or a ground power unit (GPU). Ice-protection, fire-detection, and fire-extinguishing systems are provided for each engine. The engine pylons have ram-air inlets and exhausts to provide cooling airflow through the cabin air heat exchangers.

ICE PROTECTION SYSTEM Anti-ice protection is provided to the engine inlets, and deice protection is provided to the wings, and empennage by engine bleed air. Engine bleed air directly heats the engine inlets and generator cooling inlets. The wings, vertical tail, and horizontal stabilizers are deiced by boots inflated by engine bleed air regulated to 20 psig (service air). The windshields are electrically anti-iced and defogged. Electric heat also anti-ices the pitot-static systems, stall-warning vane, and engine inlet-mounted T2 sensors. Ice detection lights on the glareshield help the pilot detect icing on the windshield. A light on the outside left fuselage helps the pilot detect icing on the wings.

HYDRAULIC SYSTEM A single electrically driven hydraulic pump supplies pressure for operation of the landing gear and wheel brakes through a closed center system. The main gear are equipped with hydraulically operated antiskid-controlled wheel brakes. Pneumatic backup is available for emergency landing gear extension and braking.

FLIGHT CONTROLS Primary flight control is accomplished through conventional cable-operated surfaces. An aileronrudder interconnect provides improved lateral stability. Trimming is provided by aileron, elevator, and rudder tabs. The elevator trim is both mechanically and electrically actuated. Aileron and rudder trim are electrically activated. The flaps are electrically actuated and are on the trailing edges of the wing. Electrically powered speedbrakes are on the upper and lower wing surfaces. Nosewheel steering is mechanically controlled by the rudder pedals through steering bungees.

ENVIRONMENTAL SYSTEM The aircraft has a two-zone automatic temperature control system that is divided into cabin and cockpit. An independent vapor cycle air-conditioning system provides cooling to the cabin and cockpit. Conditioned engine bleed air is used for cabin pressurization and temperature control. Cabin pressurization is controlled by an autoscheduling pressurization system. The crew need only to adjust destination elevation any time prior to or during flight and the controller automatically controls cabin pressure for operation at the highest practical differential pressure with minimum rates and changes. A 22-cubic-foot oxygen bottle (40-cubic-foot optional) supplies oxygen to the quick-donning masks for the crew and automatic dropout masks for each passenger. If cabin altitude becomes excessive, passenger oxygen masks deploy automatically (utilizing an electrically actuated solenoid) and can be deployed manually upon pilot command.

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AVIONICS The Mustang uses a Garmin G1000 three-display “glass cockpit” to present most indications for flight instrumentation, navigation, avionics, and aircraft systems. The displays include two 10.4inch primary flight displays (for pilot and copilot) and a 15-inch multifunction display. The standard factory-installed avionics package includes the fully integrated flight instruments, flight guidance, communications, and navigation systems. The navigation system includes GPS, ground-based navigation, and is WAAS-capable. An integrated engine indicating and crew alerting system (EICAS) is included. The Garmin G1000 system manages the instrument and engine displays, the autopilot, flight guidance systems, and the flight director. Terrain and traffic avoidance systems and color radar are standard equipment. Data link weather capability is available with subscription.

Other publications that are not required to be in the aircraft include: • Operating Manual • G armin G1000 Pilot’s Guide for the Cessna Citation Mustang • FAA-approved Weight and Balance Manual

SINGLE-PILOT OPERATION The following are required when the airplane is operated with a crew of one pilot, per applicable operating rules: 1. Operable GFC-700 Autopilot 2. Headset with microphone (must be worn) 3. FAA-approved Pilots’ Abbreviated Normal Procedures Checklist (as revised)

PUBLICATIONS

4. FAA-approved Pilots’ Abbreviated Emergency and Abnormal Procedures Checklist (as revised)

The following publications must be immediately available to the flight crew:

5. Provisions for storage and retention of navigation charts, accessible to the pilot from the pilot station

• FAA-approved AFM contains the limitations, data pertinent to takeoffs and landings, and weight and balance data. Information in the AFM always takes precedence over any other publication. • F AA-approved Citation Mustang Abbreviated Checklist—Normal Procedures contains abbreviated normal operating procedures and abbreviated performance data. If any doubt exists or if the conditions are not covered by the checklist, the AFM must be consulted. • FAA-approved Citation Mustang Abbreviated Checklist—Emergency/Abnormal Procedures contains emergency and abnormal procedures. If any doubt exists or if the conditions are not covered by the checklist, the AFM must be consulted.

LIMITATIONS For specific information on limitations, refer to the FAA-approved AFM.

EMERGENCY/ ABNORMAL For specific information on Emergency/Abnormal procedures, refer to the FAA-approved AFM.

• G armin G1000 Cockpit Reference Guide for the Citation Mustang

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1 AIRCRAFT GENERAL

CITATION MUSTANG PILOT TRAINING MANUAL

Table 1-1.   CAS MESSAGES AFT DOOR DESCRIPTION INHIBITS

Tailcone baggage door is not fully secured. EMER

CABIN DOOR DESCRIPTION INHIBITS

Indicates the cabin door is not fully secured. EMER

CHECK DOORS DESCRIPTION INHIBITS

Indicates a door monitor has not been properly tested or has failed. EMER, IN AIR, LOPI, TOPI

NOSE DOOR L-R DESCRIPTION INHIBITS

Revision 1.1

One or both of the nose baggage doors are not fully secured. EMER

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1 AIRCRAFT GENERAL

CITATION MUSTANG PILOT TRAINING MANUAL

QUESTIONS 1. The maximum weight allowed in the nose baggage compartment is: A. 320 pounds B. 200 pounds C. 150 pounds D. 500 pounds

6. Minimum equipment to operate the Citation Mustang as a single pilot: A. Operable autopilot B. Headset with microphone C. Approved checklist and navigation charts D. All the above

2. The maximum takeoff weight for the Citation Mustang is: A. 8,730 pounds B. 8,645 pounds C. 8,000 pounds D. 6,854 pounds

7. The following flight controls are electrically actuated: A. Nosewheel steering, aileron, brakes B. Elevator, aileron, flaps, speedbrakes C. Elevator trim, flaps, rudder D. Aileron trim, rudder trim, speedbrakes, flaps

3. During single-pilot operation, the maximum number of passenger seats, excluding the pilot seat, is: A. 7 B. 6 C. 5 D. 4 4. Regarding the emergency exit escape hatch: A. The hatch is pushed outside the aircraft after the D handle is pulled. B. After the D handle is unlatched, rotate the hatch down into the cabin. C. If the hatch is not properly latched and locked, an amber CAS message will display. D. The hatch cannot be opened from the outside. 5. The amber AFT DOOR CAS message is activated by: A. Aft door in the full open position B. Aft door unlocked C. Aft baggage compartment light being left on D. Key left in the aft door

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CITATION MUSTANG PILOT TRAINING MANUAL

CHAPTER 2 ELECTRICAL POWER SYSTEM INTRODUCTION.................................................................................................................. 2-1 GENERAL................................................................................................................................2-1 DESCRIPTION....................................................................................................................... 2-3 COMPONENTS..................................................................................................................... 2-3 Battery............................................................................................................................. 2-3 Standby Battery............................................................................................................... 2-4 Starter-Generators........................................................................................................... 2-4 Ground Power Unit.......................................................................................................... 2-5 Distribution...................................................................................................................... 2-5 System Protection............................................................................................................ 2-7 CONTROLS AND INDICATIONS........................................................................................ 2-9 Battery Switch................................................................................................................. 2-9 Battery Disconnect switch............................................................................................. 2-10 Interior Disconnect Switch............................................................................................ 2-10 Avionics Standby Instrument Switch............................................................................ 2-10 Generator Switches....................................................................................................... 2-11 Engine Start Buttons..................................................................................................... 2-11 Indications..................................................................................................................... 2-11 OPERATION........................................................................................................................ 2-12 Preflight......................................................................................................................... 2-12 Starting (First Engine)................................................................................................... 2-12 Starting (Second Engine).............................................................................................. 2-13

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2 ELECTRICAL POWER SYSTEMS

CONTENTS

CITATION MUSTANG PILOT TRAINING MANUAL

Starting (In Flight)......................................................................................................... 2-13 Starting (Assisted by External Power Unit)................................................................... 2-13 LIMITATIONS...................................................................................................................... 2-13 EMERGENCY/ABNORMAL.............................................................................................. 2-13 2 ELECTRICAL POWER SYSTEMS

QUESTIONS........................................................................................................................ 2-15

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ILLUSTRATIONS Figure Title Page 2-1.

Electrical System Schematic................................................................................... 2-2

2-3.

GPU Receptacle....................................................................................................... 2-5

2-4.

Aft J-Box................................................................................................................. 2-6

2-5.

CB Panels................................................................................................................. 2-8

2-7.

BATT Switch......................................................................................................... 2-10

2-6.

BATTERY DISCONNECT Switch....................................................................... 2-10

2-8.

INTERIOR DISCONNECT Switch...................................................................... 2-10

2-9.

AVIONICS Standby Instrument Switch................................................................ 2-11

2-10.

Generator Switches................................................................................................ 2-11

2-11.

ENGINE START Buttons...................................................................................... 2-11

2-12.

Electric Display (Normal)...................................................................................... 2-12

TABLES Table Title Page 2-1.

EMERGENCY BUS ITEMS................................................................................... 2-7

2-2.

CAS MESSAGES.................................................................................................. 2-14

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2 ELECTRICAL POWER SYSTEMS

2-2. Battery..................................................................................................................... 2-4

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2 ELECTRICAL POWER SYSTEMS

INTENTIONALLY LEFT BLANK

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2 ELECTRICAL POWER SYSTEMS

CHAPTER 2 ELECTRICAL POWER SYSTEMS

INTRODUCTION This chapter provides a description of the electrical power system on Citation Mustang aircraft (Figure 2-1). The DC system consists of storage, generation, distribution, and system monitoring. Provision is also made for a limited supply of power during emergency conditions in flight and connection of a ground power unit (GPU) while on the ground.

GENERAL Direct current provides the principal electrical power for the Citation Mustang. Normal aircraft system voltage is 28.5 VDC. Two generators are the primary power sources (one generator is capable of supplying all standard requirements). Secondary sources (battery or external power) may also be used.

The battery and emergency buses normally tie to the main system, but they may isolate to only the battery or external power sources. When the aircraft is on the ground, an external DC power unit may supply electrical power to all buses.

Normal distribution of DC power is via a left and right feed bus connected by a crossfeed bus. This arrangement allows either generator to power the entire system or, working in parallel, to share the system load.

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2-1

2-2

L DC AMPS

V

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A

# 2

GROUND

INT DISC

INTERIOR BUS

CROSSFEED BUS

R AVN #2 SSR

R AVIONICS #2

200 A MP

R ELE #1

R FEED BUS #1

R FEED BUS #2

R ELE #2

# 1

V

STANDBY BATTERY PACK

BATT TEST

OFF

STBY INST

100 AMP A/C COMP

R DC AMPS

100 AMP W/S DEICE

A

V

BATT VOLTS

AFT J-BOX CIRCUIT BREAKERS

R GEN VOLTS

R AVN EMERG

# 2

S T A R T

R

B U S

C B

P W R

E M E R G

R STARTERGENERATOR

START RELAY

STANDBY INST SWITCH

B U S

S H U N T

R GEN RELAY R

L BOOST SSR

R SSR #2 BUS BAR

R SSR #1 BUS BAR

R AVN #1 SSR

R AVIONICS #1

R ELE EMERG

NORM

MASTER INTERIOR SSR

Figure 2-1.   Electrical System Schematic

L ELE #2

200 A MP

L SSR #1 BUS BAR

L AVN SSR

L AVIONICS

L ELE EMERG

L FEED BUS #2

L ELE #1

L FEED BUS #1

R BOOST SSR

100 AMP W/S DEICE

B U S

S H U N T

L GEN L RELAY

# 1

S T A R T

BATT POWER RELAY

A

S T A R T L AVN EMERG

BATT DISC RELAY

BATT AMPS

R

AVN EMER SSR LEFT

BATTERY

BATT T TEMP SENSOR

L

BATTERY POWER

LEGEND

50 AMP HYD PUMP

50 AMP FLAPS

L GEN VOLTS

S T A R T

L

START RELAY

L STARTERGENERATOR

EXT PWR CONNECTOR

EMER BUS RELAY

BATTERY BUS

2 ELECTRICAL POWER SYSTEMS

EXT PWR RELAY

CITATION MUSTANG PILOT TRAINING MANUAL

CITATION MUSTANG PILOT TRAINING MANUAL

The Mustang electrical system primarily provides 28-VDC power to operate electrical devices throughout the aircraft. A starter-generator is used to start its respective engine, with starting power coming from the battery or from a GPU. Additionally, the startergenerator of a functioning engine (with battery assistance) can be used to start the opposite side engine. One generator is capable of supplying all standard electrical requirements if a generator fails. DC power is routed from each J-box feed bus through individual circuit breakers to each of the circuit-breaker buses in the cockpit CB panels. Cockpit circuit breakers control power to individual systems. Battery power is supplied to a hot battery bus and then through the battery relay to the crossfeed bus and the left and right feed buses. When the BATT switch is in the EMER position, emergency DC power is supplied from the battery bus through the emergency power relay to the emergency bus circuit breakers on each cockpit CB panel. If the battery switch is in the BATT position, generator power is supplied through the battery relay to the hot battery bus to charge the battery and from the crossfeed bus through the emergency relay to the emergency power buses. The external power receptacle is underneath the right engine nacelle. First engine start is performed from the battery unless using external power. The second engine start may be powered three different ways: • With external power (if the first generator switch is OFF) • From the battery (if the first generator switch is OFF and external power is not connected) • From the battery with assistance from the first generator (ground only) if the first generator is online

power to the main bus system in the tail cone. This bus system and its associated relays provide connections and power management for the battery and provide for connection to a GPU. This bus system also allows either starter-generator to assist the other during starting and allows the two starter-generators to operate “in parallel” to share the electrical load evenly. From the main bus system in the tail cone, power is distributed through circuit breakers in the tail cone directly to a few electrical devices in or near the tail cone. More power is routed forward from the main buses through feeder cables to the cockpit buses. Buses on each side of the cockpit (behind the CB panels) supply power through the cockpit circuit breakers and panel controls to most of the aircraft electrical devices. Cockpit indicators monitor electrical system status and performance. Cockpit panel controls allow the crew to directly manage the generation and distribution of electrical power. Relays, solid state relays (SSRs), circuit breakers, current limiters, and generator control units (GCUs) protect the electrical system, and assist the crew in managing the supply and flow of electrical power.

COMPONENTS BATTERY A standard lead acid battery provides 24 volts rated at 28 amp hours. An optional NiCad battery provides 24 volts rated at 28 amp hours. The battery is in the tail cone compartment (Figure 2-2). It has a manual quick-disconnect, and is accessible through the tail cone door. The battery connects to the battery bus. A battery disconnect relay between the battery and its ground provides an electrical disconnect during certain conditions. A BATTERY disconnect switch (Figure 2-3) is in the cockpit on the left side console panel. This switch opens the battery disconnect relay. Use this switch in case of a battery overheat or stuck start relay.

Normally, when both engines are operating, the starter-generator in each engine provides 28-VDC

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2 ELECTRICAL POWER SYSTEMS

DESCRIPTION

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STANDBY BATTERY The standby instrument battery is a 24 V, 1.2 amp hour NiCad battery. The standby battery is controlled with the AVIONICS STBY INST switch.

2 ELECTRICAL POWER SYSTEMS

The standby instrument battery is in the radome on the avionics shelf assembly. The battery automatically supplies electrical power to the standby airspeed, attitude, and altitude instruments and the lighting for the magnetic compass when normal electrical power is not available. Figure 2-2.   Battery

NOTE The optional NiCad battery is susceptible to, and must be protected from, overheat due to excessive charging or discharging. During an external power start cycle, to prevent battery discharge, the battery disconnect relay automatically disconnects the battery from its ground. A GPU start is not a battery start. Starting the engines with an external power source is recommended practice to prolong the life of the batteries and conserve battery power for times when battery starts must be accomplished. When it is anticipated the aircraft will be idle for more than 2 days, it is advisable to disconnect the battery to prevent frequency memory circuits, or other equipment that may be powered by the battery bus, from draining the batteries. A battery in good condition supplies power to all buses for a minimum of 10 minutes with maximum load. If powering only the battery and emergency buses, battery life should be a minimum of 30 minutes. An INTERIOR DISCONNECT switch is on the pilot side console panel (Figure 2-7). When the switch is selected to the up position, the master interior relay opens, shutting off all electrical power in the cabin. When the switch is in the NORM position, the master interior relay closes, and electrical power flows to the cabin normally (when DC power is available and the other electrical controls are in the appropriate positions).

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STARTER-GENERATORS Two engine-driven DC starter-generators (one on each engine accessory gearbox) are the primary source of aircraft electrical power and supply power to all DC buses. Each generator is air cooled, rated at 29 VDC, and regulated to 28.5 volts. The generators are engine-starting motors that revert to generators at the completion of the start cycle. Each generator system operates independently, but power distributes evenly through bus systems that are in parallel except during fault conditions. DC power from the engine-driven generators distributes to two feed buses (see Figure 2-1). Each feed bus connects through a 200-amp current limiter to the crossfeed bus, allowing each feed bus to parallel the other. During normal operation, the generators share loads equally (within ±10% of the total load) via the crossfeed bus. Each starter-generator is regulated by its own generator control unit (GCU). Generator power routes from the crossfeed bus through the battery relay (when it is closed) to the battery bus. This provides power to charge the battery, and during generator operation powers the items on the battery bus. Normally (with the battery switch set to BATT), generator power routes from the crossfeed bus through the emergency relay to power the cockpit emergency buses and through the battery relay to power the emergency bus in the J-box. The battery and emergency relays are operated with the battery switch.

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A GPU can be connected to the aircraft DC system through a receptacle in the fuselage below the right engine nacelle (Figure 2-4). External power is routed through the external power relay to the battery bus. The battery charges from the GPU, regardless of the battery switch position. A GPU providing a maximum voltage of 29 VDC may be used. The left and right start controllers monitor GPU voltage and open the external power relay to disconnect the GPU from the aircraft if voltage exceeds approximately 32.5 VDC. Before connecting a GPU, ensure that the voltage of the GPU is regulated to 28–29 volts the amperage output between 800 and 1,100 amps. When using external power for prolonged ground operation (over 30 minutes), disconnect the battery to preclude overheating the battery. Do not use the battery disconnect switch.

CAUTION Some GPUs do not have reverse-current protection. If the GPU is powered off while connected to the aircraft, the battery may be rapidly discharged and/or damaged. Always disconnect the GPU from the aircraft when not in use. Connecting the external power source energizes the external power relay, which connects the external power source to the battery bus.

Setting the battery switch to the BATT position energizes the battery relay, which allows the connection of external (or battery) power from the battery bus to the emergency buses, and through the crossfeed bus to the left and right feed buses. When either the left or right generator power relay closes, the external power relay deenergizes to remove external power from the battery bus. This prevents the aircraft generators and the GPU from simultaneously applying power to the aircraft buses.

CAUTION If the battery is charged using the GPU, it must be monitored. Current from most GPUs is not regulated and a battery overheat may occur.

DISTRIBUTION DC power is distributed throughout the aircraft though several buses (see Figure 2-1) via the main junction box (aft J-box) and cockpit buses (behind CB panels).

Main Junction Box (Aft J-Box) The main junction box (aft J-box) (Figure 2-5) in the tail cone compartment contains: • Two feed buses: °° Left feed bus No. 1

°° Right feed bus No. 1 • Two start buses: °° Left start bus No. 1 °° Right start bus No. 1 • Two shunt buses: °° Left shunt bus °° Right shunt bus • Crossfeed bus °° Battery bus

Figure 2-3.   GPU Receptacle

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2 ELECTRICAL POWER SYSTEMS

GROUND POWER UNIT

CITATION MUSTANG PILOT TRAINING MANUAL

Crossfeed Bus The crossfeed bus functions solely as a bus tie connecting the battery bus, the emergency buses, and the two main feed buses into one integral system.

2 ELECTRICAL POWER SYSTEMS

If the battery switch is selected to BATT, the battery power relay closes, which connects the battery bus to the crossfeed bus. Power extends from the crossfeed bus through 200-amp current limiters to each main feed bus. Power also extends from the crossfeed bus to the cockpit emergency power circuit-breaker buses and (through the avionics power switch) to the avionics buses. Figure 2-4.   Aft J-Box

Battery Bus

Main Feed Buses Each generator (left and right) normally supplies power through its respective generator relay to its respective main feed bus (left feed bus No. 1 and right feed bus No. 1). These buses are tied together through the crossfeed bus.

Start Buses The left and right start buses provide power to the left and right start No. 2 buses and those provide power to the controllers and related systems. In order for the start relay to close, the battery switch must be in the BATT position. When the respective start relay closes, the start bus is connected to the battery bus, which supplies power from the battery or a GPU.

Shunt Buses The left and right shunt buses connect the startergenerators to the electrical system. The GCUs and starter-generators manage the connection of the start buses through the left and right start relay and the left and right generator relay. With the exception of a cross-generator start, normally only one relay at a time (either start relay or generator relay) is closed on each side to connect the corresponding start bus to the electrical system.

2-6

The battery bus is connected directly to the battery. It may receive power from a GPU and during normal operation, receives its power from either or both generators.

Cockpit Distribution and CB Panels Various feed-extension buses, avionics buses, emergency buses, and the interior buses are in the cockpit.

Feed Extension Buses From each main feed bus (left feed No. 1 and right feed No. 1) in the tail cone, various extension feed buses distribute power to components through controls and circuit breakers in the cockpit. The main left and right feed-extension buses are behind the pilot and copilot CB panels respectively (Figure 2-6). Other feed-extension buses are also behind the corresponding CB panels and are powered through 25-amp circuit breakers.

Avionics Buses Avionics buses (left and right) are powered from the respective main feed buses through solid-state relays (SSRs) when the AVN MASTER switch is selected ON. These buses provide power to the aircraft avionics, except the avionics that are on the emergency buses.

FOR TRAINING PURPOSES ONLY

Emergency Buses Emergency bus items are listed in Table 2-1. The aircraft has five emergency buses: • Emergency power circuit-breaker bus • Left electrical emergency bus • Right electrical emergency bus • Left avionics emergency bus • Right avionics emergency bus (with standby battery) The emergency power circuit-breaker bus is directly connected to the battery bus at all times. Other buses are powered from either the crossfeed bus, the battery bus, or the standby battery. With the battery switch in the BATT position, power to the emergency buses is from the crossfeed bus. Because the crossfeed bus normally feeds the emergency buses, the pilot must use the battery switch to energize the emergency power relay to the EMER position, which switches all emergency buses from the disabled crossfeed bus to the battery bus. Table 2-1.   EMERGENCY BUS ITEMS • PFD 1 - Reversion mode • COM 1 • NAV 1 (including marker beacon) • GPS 1 • ADC 1 • AHRS 1 • Pilot and copilto audio panel • L and R N1, N2, and ITT indications • L Oil Temperature • L Fuel flow • Fuel Temperature • Magnetic compass light

Revision 1.0

• Battery voltage indication • Autopilot control panel (HDG, CRS, ALT, knobs only) • Cabin altitude and differential pressure indications • Cabin dump system • Cockpit flood light • Pilot pitot-static heat • Landing gear indicator lights • Avionics audio warnings • Standby instruments (airspeed, altitude, attitude) • ELT GPS position interface

With the battery switch in the EMER position, the following aural warnings are available: • Terrain awareness and warning system (TAWS) alert • Autopilot disconnect • Check altitude • Barometric minimum descent altitude/decision height • Vertical track • Marker beacon

CAUTION With the battery switch in the EMER position, some aural warnings are NOT available, including: • Stall warning • Landing gear • Overspeed • Traffic alerts TIS, TAS, TCAS

SYSTEM PROTECTION Generator Control Units Two GCUs regulate, parallel, and protect the generators. The GCUs are in the tail cone, with one unit dedicated to each starter-generator. Each GCU controls a field and generator relay. Each generator relay connects the generator to its feed bus. The GCU permits the generator relay to close when the cockpit generator switch is in GEN and the generator output is within 0.5 volts of normal system voltage (28.5 VDC). When the GCU senses an internal feeder fault (short circuit) or an overvoltage, the respective side generator and field relays open. These relays also open when the ENGINE FIRE switchlight is selected.

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2 ELECTRICAL POWER SYSTEMS

CITATION MUSTANG PILOT TRAINING MANUAL

CITATION MUSTANG PILOT TRAINING MANUAL

2 ELECTRICAL POWER SYSTEMS

Figure 2-5.   CB Panels

A reverse current (10% of total load) or undervoltage opens only the generator relay, removing the generator from the system but leaving the field relay closed. The GCU utilizes software and solid-state circuitry to perform the following operations: • Control voltage regulation • Load sharing • Overvoltage/overexcitation protection • Automatic generator line contactor control 2-8

• • • • • • • • •

everse current protection R Overload protection Overspeed protection Open ground protection Open shunt protection Open point of regulation (POR) protection Starter cutoff Field weakening Ground fault protection

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CITATION MUSTANG PILOT TRAINING MANUAL

Parallel feeder cables (between each DC feed bus in the tail cone and the corresponding feed-extension buses in the cockpit) receive protection from circuit breakers. Various other circuit breakers on the feed buses in the tail cone protect against overload. Current limiters, also known as fuse limiters, are provided to protect against major electrical overload. A list of the protected buses and components is provided below: • Left feed bus No. 1—200 amps • Right feed bus No. 1—200 amps • Air-conditioner compressor—100 amps • Windshield heat (left)—100 amps • Windshield heat (right)—100 amps • Flaps—50 amps • Hydraulic powerpack—50 amps • HF radio (optional)—50 amps

Solid-State Relays Solid-state relays (SSRs) serve as a combination circuit breaker and relay for numerous components. SSRs are individually controlled by cockpit system switches or, in some instances, by remotely mounted printed circuit boards (PCBs). SSRs are installed in either a 25- or 10-amp size; however, they are resistor-adjustable for lower amperage trip points. The following buses and components are SSR-protected and controlled: • Left avionics bus • Right avionics bus (No. 1) • Right avionics bus (No. 2) • Left avionics emergency bus • Master interior bus • Cockpit fan • Cabin fan • Condenser fan • Left fuel boost pump • Right fuel boost pump

• • • •

eft ignitor No. 1 L Right ignitor No. 1 Left ignitor No. 2 Right ignitor No. 2

Relays and Engine Starting For generator-assisted second engine starts, the battery power relay opens to prevent high-current flow from the crossfeed bus to the battery bus and protects the 200-amp current limiters. This causes starting current from the online generator and battery to flow through the two starter relays and battery bus to the starter. A blown 200-amp current limiter splits the feed buses, preventing generator paralleling. Pressing the starter button for GPU starts, first opens the battery disconnect relay to prevent the battery cycles, then closes the start relay. If GPU voltage is excessive, an overvoltage sensor opens the external power relay and breaks the circuit to the battery bus. External power disable relays also disconnect the GPU from the battery bus whenever a generator relay closes, bringing a generator online.

CONTROLS AND INDICATIONS Control of DC power is maintained with a battery switch and two generator switches (Figure 2-7).

BATTERY SWITCH The battery switch is on the pilot DC POWER subpanel and has three positions: BATT, OFF, and EMER. If the battery switch is in the OFF position, the battery bus isolates from all other buses in the system with the exception of the emergency power circuitbreaker bus.

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2 ELECTRICAL POWER SYSTEMS

Circuit Breakers and Current Limiters

CITATION MUSTANG PILOT TRAINING MANUAL

NOTE The battery switch on the DC POWER subpanel must be in the BATT position for the battery disconnect switch to operate.

2 ELECTRICAL POWER SYSTEMS

If the battery ground is open, the battery cannot supply electrical power to the aircraft or receive a charge from the generators.

CAUTION

Figure 2-7.   BATT Switch

When the battery switch is in the BATT position, the battery power relay closes, completing a circuit to the crossfeed bus. The emergency relay deenergizes while the battery relay is in the BATT position and completes a circuit to the emergency buses from the crossfeed bus. In the EMER position, only the emergency power relay energizes, which connects the emergency buses to the battery bus. These buses receive power from the battery or external power. When external power is not applied to the aircraft and the generators are online, placing the battery switch in EMER or OFF isolates the battery from any charging source.

BATTERY DISCONNECT SWITCH A guarded battery disconnect switch (see Figure 2-6) is above the pilot armrest on the left side console panel. The switch has two positions: BATTERY (disconnect) and NORM. It disconnects the battery and is used only for abnormal operations involving stuck start relay or battery overtemperature. Activating this switch uses battery power to open the battery disconnect relay on the ground side of the battery.

Do not use the battery disconnect switch for an extended time. The battery disconnect relay will continue to draw a small current from the battery until the battery is discharged. The battery disconnect relay will then close, resulting in a very high charge rate and probable overheat.

INTERIOR DISCONNECT SWITCH The interior disconnect switch is above the pilots armrest on the left console panel by the battery disconnect switch. The interior disconnect switch disconnects the cabin lights (except for the emergency exit lights operated by the pax safety switch), the cabin DC-DC converters, and the cabin XM radio.

Figure 2-8.   INTERIOR DISCONNECT Switch

AVIONICS STANDBY INSTRUMENT SWITCH

Figure 2-6.   BATTERY DISCONNECT Switch

2-10

The avionics standby instrument switch is on the AVIONICS pilot switch panel in the cockpit. The switch can be set to the STBY INST, OFF, or BATT TEST position. The switch supplies power to the right avionics emergency bus.

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CITATION MUSTANG PILOT TRAINING MANUAL

Placing the switch in the spring-loaded RESET position rebuilds field voltage to provide a means of resetting a generator that has tripped as a result of a fault condition.

Figure 2-9.   AVIONICS Standby Instrument Switch

When the standby battery is powering the standby instruments, the amber light adjacent to the switch illuminates. Selecting BATT TEST performs a capacity check on the standby battery. A successful test is indicated by a green light adjacent to the switch.

Two engine start buttons (L and R) (see Figure 2-7) on the pilot ENGINE START subpanel activate a circuit to close the associated start relay and allow starting current to flow from the battery bus to the starter. A starter disengage (DISENG) button between the starter buttons opens the start circuit if manual termination of the start sequence is desired (see Figure 2-11).

GENERATOR SWITCHES Two generator switches (L GEN and R GEN) are on the pilot DC POWER subpanel (see Figure 2-10). The generator switches have three positions: L (or R) GEN, OFF, and RESET. Setting the switch to L GEN or R GEN allows the GCU to close the generator relay and connects the generator to its feed bus. The ammeter indicates the generator output to the feed buses. With the switch in the OFF position, the generator relay opens and the ammeter shows no generator load to the feed buses.

Figure 2-11.   ENGINE START Buttons

Pushing the engine start button illuminates a white light in the starter button as a direct indication that the start relay is closed.

INDICATIONS The DC electrical system is monitored by: • Crew alerting system (CAS) messages • Engine indicating and crew alerting system (EICAS) display window °° DC AMPS display °° DC VOLTS display °° BATTERY AMPS display °° BATTERY VOLTS display • ENGINE START button lights

Figure 2-10.   Generator Switches

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2 ELECTRICAL POWER SYSTEMS

ENGINE START BUTTONS

CITATION MUSTANG PILOT TRAINING MANUAL

EICAS Display Window

BATTERY–AMPS Display

The DC window of the EICAS display provides dual generator indications (left and right) for both voltage (VOLTS) and current (AMPS), and also BATTERY voltage and current.

The BATTERY–AMPS display is a digital display on the bottom of the DC window (Figure 2-12). The display indicates current into or from the battery. Positive amperage indicates battery charging. Negative amperage indicates battery discharge.

2 ELECTRICAL POWER SYSTEMS

OPERATION PREFLIGHT During the exterior preflight, visually check the battery for signs of deterioration or corrosion. Do not connect external power until completing these checks. During the interior preflight, place the generator switches to GEN if the intention is a battery start or to the OFF position if external power is desired. Place the battery switch to BATT and verify the voltage display is at or above 24 volts minimum (22 volts minimum for a NiCad). After checking lights and pitot heat, turn the battery switch to the OFF position. Figure 2-12.   Electric Display (Normal)

VOLTS Display The left and right generator VOLTS displays are on the upper-left area of the DC window (Figure 2-12). Each VOLTS display indicates voltage at its respective generator. In reversionary mode, only the digits are displayed.

AMPS Display The left and right generator AMPS displays are on the upper-right area of the DC window (Figure 2-12). Each display indicates current flow from its respective generator to its respective DC feed bus. During normal operation, the indication should be parallel within ±10% of total load. Amperage between the starter-generator and the battery bus is not reflected on the AMPS displays. In reversionary mode, only the digits are displayed.

BATTERY–VOLTS Display The BATTERY–VOLTS display is a digital display on the bottom-center area of the DC window. The display indicates voltage on the battery bus.

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STARTING (FIRST ENGINE) Before starting the engines, recheck the generator switches for proper position and verify battery voltage. Ensure that the battery switch is in the BATT position. Depressing the L or R ENGINE START button: • Closes the respective start relay • Activates the electric fuel boost pump Closure of the start relay (indicated by illumination of the start button white light) connects battery bus power to the starter for engine rotation. At approximately 8% turbine rpm (N2): • FADEC commands fuel flow to the start nozzles • Ignition is activated by the full-authority digital engine control (FADEC) • A green IGN appears on the multifunction display (MFD) at the upper interturbine temperature (ITT) scale and indicates current to one or both exciter boxes

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Within 10 seconds, combustion should occur as evidenced by rising ITT.

associated start relay closes and the boost pump on that side activates.

As the engine accelerates through 48.6% N2: • The GCU starter overspeed sensor automatically terminates the start sequence. • The start relay opens. • The electric boost pump is deenergized. • The GEN OFF message disappears from the CAS window(GEN switch ON). • The green IGN indication extinguishes. • N2 digits change from white to green.

The only difference between an in-flight start and a ground start with one generator online, is that the start relay on the same side as the operating generator does not close and the battery power relay opens. This isolation of the start circuit from the operating generator and buses in flight is through left squat switch logic and is required by certification regulations.

STARTING (SECOND ENGINE)

The protection circuit for the 200-amp current limiter is the same as previously described. Refer to the “Airstart Envelope” graph in “Limitations” of the Airplane Flight Manual (AFM).

Before starting the second engine, the operating engine must be INCREASED to 10% above ground idle N2.

STARTING (ASSISTED BY EXTERNAL POWER UNIT)

CAUTION If the operating generator drops off-line during a cross-generator start (GEN OFF L-R), an ENG CTRL SYS L or R CAS message posts, of ITT indication is lost at any time during the start sequence, abort the start immediately by bringing the throttle to CUTOFF to reduce the possibility of a hot or hung engine start. For a second engine start on the ground, the operating generator assists the battery in providing current to the starter.

A GPU can be used for engine starts. Check for voltage regulation to a maximum of 29 VDC and 800/1,100 amps. When external power starts are planned, the generator switches remain in the OFF position until the removal of external power from the aircraft. Otherwise, when the first generator comes online, the external power relay opens and the GPU automatically disconnects from the battery bus. The second engine start becomes a generator-assist battery start.

LIMITATIONS

When the remaining start button activates, both start relays close and the white light in each starter button illuminates.

For specific limitations, refer to the FAA-approved AFM.

When one generator relay closes and the other energizes as a starter, the battery disable relay causes the battery relay to open the circuit between the crossfeed bus and the battery bus in order to protect the 200-amp current limiter.

EMERGENCY/ ABNORMAL

STARTING (IN FLIGHT)

For specific information on emergency/abnormal procedures, refer to the appropriate checklist or FAA-approved AFM.

An engine start in flight using the start button is a battery start only. The squat switch disables generator-assist capability when airborne. Only the

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Table 2-2.   CAS MESSAGES BATTERY O’TEMP DESCRIPTION

2 ELECTRICAL POWER SYSTEMS

INHIBITS

BATT TEMP FAIL

With the optional NiCad battery installed, a battery overtemperature warning system warns the pilot of abnormally high battery temperatures. If the temperature reaches 71°C (160°F) a red BATT O’TEMP message displays and the MASTER WARNING lights flash. NONE

DESCRIPTION INHIBITS

The battery temperature sensor has failed. LOPI, TOPI

GEN OFF L-R DESCRIPTION INHIBITS

One electrical generator has gone offline. NONE

GEN OFF L-R DESCRIPTION INHIBITS

Both electrical generators have gone offline. NONE

AFT JBOX CB L-R DESCRIPTION INHIBITS

Start control circuit breaker (located in the aft j-box) is tripped) EMER

AFT JBOX LMT L-R DESCRIPTION

INHIBITS

The right and left current limiter circuit breakers are located in the aft j-box. When blown normal power will be available to both busses, but the cross-feed bus will not supply power form the battery or from the opposite generator in the event of a generator failure. EMER

BATTERY O’TEMP DESCRIPTION

INHIBITS

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With the optional NiCad battery installed, a battery overtemperature warning system warns the pilot of abnormally high battery temperatures. An internal temperature of 63–70°C (145–156°F) displays an amber BATT O’TEMP message and steady MASTER CAUTION lights. NONE

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1. What action should be taken if the engine starter fails to disengage? A. E N G I N E S TA R T D I S E N G switch—PRESS B. FADEC RESET switch—Select affected side and release C. IGNITION switch—NORM D. AP/TRIM DISC button—PRESS 2. What is indicated by illumination of the light in the start buttons? A. Engine has not reached a stabilized idle B. Start relay is closed C. Engine has been started and the generator is online D. Generator relay is closed 3. If the AFT JBOX LMT L CAS message is displayed in the CAS window, what bus or busses would be lost if the L generator fails? A. L FEED BUS and L SHUNT BUS B. R START BUS C. L AVIONICS EMER BUS D. R EXTENSION BUS 4. With the battery disconnect switch in the battery disconnect position: A. Do not use the switch for an extended period of time B. The aircraft battery is connected to the airframe C. The standby battery is connected to the airframe D. The keep alive for the avionics is disconnected from the airframe 5. With the battery switch in EMER and the generators OFF, battery life should be a minimum of: A. 5 minutes B. 30 minutes C. 60 minutes D. 90 minutes

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6. With the battery switch in the EMER position and the generators OFF, there is no: A. Cockpit floodlight B. Landing gear indication lights C. Exterior lights D. EICAS messages 7. The wing and tail deice systems are inoperative with the battery switch in the: A. EMER position B. NORM position C. ON position D. INTERIOR DISCONNECT position 8. For engine start with a sealed lead-acid battery, minimum battery voltage is: A. 18 VDC B. 20 VDC C. 24 VDC D. 28 VDC 9. The maximum electrical load per generator on the ground at idle is limited to: A. 50 amps B. 100 amps C. 150 amps D. 200 amps 10. Battery starts are limited to: A. One engine start per hour B. Two engine starts per hour C. Three engine starts per hour D. Four engine starts per hour 11. If a ground power unit is being used to start the engines, the generator switches should be in which position? A. ON B. OFF C. Reset D. Manual

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2 ELECTRICAL POWER SYSTEMS

12. With the STBY INST switch in the BATT TEST position, the pilot should see the following indication: A. A green STBY INST message in the CAS window B. The green test light illuminated C. An amber BATT TEST message in the CAS window D. The red LED test light illuminated

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CHAPTER 3 LIGHTING CONTENTS INTRODUCTION.................................................................................................................. 3-1 GENERAL................................................................................................................................3-1 INTERIOR LIGHTING.......................................................................................................... 3-2 Flight Compartment Lighting......................................................................................... 3-2 Magnetic Compass Light................................................................................................ 3-3 Entry/Exit Lighting and Entry Light Switch................................................................... 3-3

Passenger Safety Light System....................................................................................... 3-4 Baggage Compartment Lighting..................................................................................... 3-5 EXTERIOR LIGHTING......................................................................................................... 3-5 Landing/Recognition/Taxi Lights.................................................................................... 3-5 Beacon............................................................................................................................. 3-6 Anticollision Lights......................................................................................................... 3-6 Navigation Lights............................................................................................................ 3-6 Wing Inspection Light..................................................................................................... 3-6 LIMITATIONS........................................................................................................................ 3-6 EMERGENCY/ABNORMAL................................................................................................ 3-6 QUESTIONS.......................................................................................................................... 3-7

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ILLUSTRATIONS 3-1.

Flight Compartment Lighting Controls................................................................... 3-2

3-2.

Cockpit Overheat Lights and Controls.................................................................... 3-3

3-3.

Compass Light......................................................................................................... 3-3

3-4.

Entry Lights Switch................................................................................................. 3-3

3-5.

Cabin Lighting......................................................................................................... 3-4

3-6.

Cabin Lighting Controls.......................................................................................... 3-4

3-7.

Emergency Exit Light.............................................................................................. 3-4

3-8.

No Smoking/Fasten Seat Belt Sign.......................................................................... 3-4

3-9.

Nose Baggage Light and Switch.............................................................................. 3-5

3-10.

Aft Baggage Light Switch....................................................................................... 3-5

3-11.

Landing Lights......................................................................................................... 3-6

3-12.

Beacon Light............................................................................................................ 3-6

3-13.

Position Lights......................................................................................................... 3-6

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CHAPTER 3 LIGHTING

INTRODUCTION This chapter provides information on lighting for the Citation Mustang. Interior lighting illuminates the flight compartment area, all flight instruments, and the passenger cabin. Exterior lighting provides necessary illumination for day or night operation.

GENERAL Interior lighting is provided for the flight compartment, cabin, windshield ice detection, and passenger safety. Most instruments are internally lighted. For general illumination, map lights, and a floodlight are above the pilot and copilot. There are standard passenger advisory lights in the cabin area, and emergency

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exit lights that consist of a light over the main cabin entry door and a table light over the emergency exit door. Reading and table lights are also available in the cabin area. The exterior lighting consists of wingtip lights (navigation/anticollision lights), landing/ recognition lights, wing inspection light, and beacon light.

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INTERIOR LIGHTING The aircraft interior lights are DC powered. Interior light circuit breakers are on the right CB panel. The breakers are grouped within the LIGHTS category. Panel back lighting is provided by light emitting diodes (LEDs). Overhead lights are halogen and the displays are liquid crystal displays (LCDs).

The panel dimmer knob controls: • Switchlights • Oxygen gauge lighting • Magnetic compass light • Three standby instruments • Landing gear position lights • Audio panels and display bezels

FLIGHT COMPARTMENT LIGHTING

A DISPLAYS dimmer knob controls the dimming of the Garmin avionics. The dimmer knob is on the LIGHTING panel (Figure 3-1).

Flight compartment panel lighting is provided by LED panels (Table 3-1).

Rotating the DISPLAYS dimmer knob clockwise increases the intensity of the Garmin displays. Rotating it counterclockwise dims the displays; rotating the knob fully counterclockwise (to the DAY position) causes the intensity to be set automatically in response to photocell sensors.

Light intensity is controlled by a PANEL dimmer knob (Figure 3-1).

3 LIGHTING

Figure 3-1.   Flight Compartment Lighting Controls Table 3-1.   COCKPIT LIGHTING PANEL DIMMER KNOB POSITION LIGHT

DAY

JUST OUT OF DAY

FULL CW

VARIABLE?

BATTERY EMER. POSITION PANEL LIGHT DIM CB-HF021

LED PANELS X 19

OFF

VERY DIM

BRIGHT

YES

OFF

ICE DETECT LIGHTS

ON

ON

ON

NO

OFF

OXYGEN GAUGE

OFF

VERY DIM

BRIGHT

YES

OFF

STANDBY AIRSPEED

OFF

VERY DIM

BRIGHT

YES

OFF

STANDBY ATTITUDE

OFF

VERY DIM

BRIGHT

YES

OFF

STANDBY ALTITUDE

OFF

VERY DIM

BRIGHT

YES

OFF

MAGNETIC COMPASS

OFF

VERY DIM

BRIGHT

YES

OFF

GEAR LIGHTS

BRIGHT

VERY DIM

VERY DIM

NO

VERY DIM

DUMP SWITCH

BRIGHT

VERY DIM

DIM

YES

OFF

MASTER CAUTION

BRIGHT

DIM

DIM

NO

DIM

R ENGINE FIRE/ L ENGINE FIRE

BRIGHT

DIM

DIM

NO

DIM

BOTTLE ARMED PUSH

BRIGHT

BRIGHT

BRIGHT

NO

NO EFFECT

STANDBY PLACARD LIGHT

ON

ON

ON

NO

ON

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The DISPLAYS dimmer knob can only control a Garmin display if AUTO brightness mode is selected. This mode is the default lighting mode but can be changed. Refer to the Garmin Mustang G-1000 Cockpit Reference Guide. A cockpit floodlight and two map lights are overhead near the center of the aircraft (Figure 3-2). A center rheostat controls the floodlight and separate rheostats to the left and right of the floodlight rheostat control each of the two map lights.

ENTRY/EXIT LIGHTING AND ENTRY LIGHT SWITCH To activate the entry/exit lighting, use the entry light pushbutton switch on the aft side of the left divider cabinet (Figure 3-4). When the entry door is opened, a green backlight illuminates a symbol on the entry light switch. After a 10-minute delay, the entry lights (if turned on) and the switch backlighting extinguish. The entry light switch turns on the fixed light above the entry door and the fixed light above the toilet.

3 LIGHTING

Figure 3-2.   Cockpit Overheat Lights and Controls

MAGNETIC COMPASS LIGHT The magnetic compass is on the windshield center post (Figure 3-3). LED backlighting is provided for night operation when the PANEL dimmer knob is in any position other than DAY.

Figure 3-4.   Entry Lights Switch

CABIN LIGHTING Cabin lighting consists of (Figure 3-5): • Two table lights • Four reading lights • Two entry lights • Passenger safety light There are oval-shaped light assemblies above the passenger seats on the cabin headliner. Each assembly consists of (Figure 3-6): • Controllable air duct outlet • Inboard button that controls the respective table light • Outboard button that controls the respective cabin light • Light assembly

Figure 3-3.   Compass Light

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Figure 3-7.   Emergency Exit Light

Figure 3-5.   Cabin Lighting

NOTE

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If activated by the G-switch, these lights remain on until deactivated by maintenance, regardless of battery switch or cabin door status. The interior disconnect switch (above the pilot armrest on the left side console panel) does not disable the passenger safety lights controlled by the PAX SAFETY switch.

Figure 3-6.   Cabin Lighting Controls

PASSENGER SAFETY LIGHT SYSTEM

The cabin lights are powered through the CABIN LIGHTS circuit breaker on the aft J-box.

The Mustang is equipped with a passenger safety light system. The pilot can activate the light above the entry door and the right table light over the emergency exit (Figure 3-7) with the PAX SAFETY light switch on the light switch grouping below the multifunction display (MFD) (see Figure 3-1). These same two lights can also be activated by a 5-G switch. Anytime 5-G is exceeded (such as in an emergency landing), the exit lights illuminate. These two safety lights are powered from the emergency power CB bus.

No Smoking/Fasten Seat Belt Sign and Switches A no smoking/fasten seat belt sign is above the table on the left side of the cabin (Figure 3-8). The no smoking placard is always visible. The PAX SAFETY–SEAT BELT switch (on the LIGHTING panel under the MFD) controls the fasten-seat-belt light. The light illuminates when the switch is set to SEAT BELT or PAX SAFETY.

NOTE The passenger safety lights illuminate for emergency lighting when the Gswitch is tripped or when the passenger safety switch on the cockpit LIGHTS switch panel is in the PAX SAFETY position.



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Figure 3-8.   No Smoking/Fasten Seat Belt Sign

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BAGGAGE COMPARTMENT LIGHTING Baggage compartment lighting includes the nose baggage compartment light and the tail cone compartment light. They are wired directly to the emergency power bus and do not require the battery switch to be in the BATT or EMER position for operation.

LIGHT

Nose Baggage Compartment

When both nose baggage doors are closed, a microswitch on each nose baggage door hinge turns the light off regardless of rocker switch position.

BAGGAGE LIGHT SWITCH

Figure 3-10.   Aft Baggage Light Switch 3 LIGHTING

The manual switch assembly of the baggage light system is an illuminated two-position rocker switch. The switch is in the baggage compartment overhead and adjacent to the light assembly (Figure 3-9). The manual switch applies DC power to the light. During daylight hours or when the light is not desired, turn the manual switch OFF, which disconnects power from the light. When the switch is OFF, it illuminates so it is easy to locate at night.

EXTERIOR LIGHTING LANDING/RECOGNITION/ TAXI LIGHTS The aircraft is equipped with two lamps that illuminate for landing and taxi purposes. The landing lights consist of two 50-watt sealed high intensity discharge (HID) lamps in the belly fairing, forward of the forward wing spar. These lamps are protected behind tempered glass covers. They are situated so the flight compartment is shielded from glare (Figure 3-11).

Figure 3-9.   Nose Baggage Light and Switch

Aft Baggage Compartment The toggle switch for the aft baggage light system is on the right side door opening. The baggage light is aft of the baggage door on the upper right side. When the switch is in the ON position, DC power is applied to the light.

A LANDING–RECOG TAXI switch on the LIGHTING panel controls the landing lights. The LANDING position provides the brightest illumination for landing. The RECOG TAXI position dims the lights to a lower intensity. The landing lights receive power through the respective LAND/REC LIGHTS circuit breakers on the aft J-box.

When the aft baggage door is closed, a microswitch on the door turns the light off regardless of the manual toggle switch position (Figure 3-10).

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NAVIGATION LIGHTS Navigation lights are in assemblies behind clear tempered glass covers. The lights are located as follows (Figure 3-13): • Red forward light—Left wingtip • Green forward light—Right wingtip • White rear light—Both wingtips

Figure 3-11.   Landing Lights

The navigation lights are controlled by a NAV switch on the LIGHTING panel. The lights are powered by the circuit breakers in the aft J-box.

3 LIGHTING

BEACON

WING INSPECTION LIGHT

The aircraft is equipped with a beacon. The beacon assembly is on the top of the vertical stabilizer for optimum line of sight visibility (Figure 3-12).

A wing inspection light is on the left side of the fuselage, above and forward of the wing leading edge. The inspection light assembly includes a halogen bulb that illuminates the outboard leading edge of the left wing. Aircraft crew utilize the light to detect wing ice accumulation during nighttime flight in icing conditions. A WING INSP switch on the LIGHTING panel supplies power to the lamp. The WING INSP light circuit breaker is on the left CB panel.

The beacon consists of a red LED assembly with a strobe rate of 50 flashes per minute. The beacon is controlled by the BEACON switch on the LIGHTING panel.

LEFT (RED) NAVIGATION LIGHT

ANTICOLLISION LIGHT

AFT (WHITE) NAVIGATION LIGHT

Figure 3-13.   Position Lights

Figure 3-12.   Beacon Light

ANTICOLLISION LIGHTS In addition to the navigation lights, each wingtip assembly contains an anticollision strobe light. The anticollision lights flash at a rate of 50 flashes per minute. The lights are controlled by the ANTI COLL switch on the LIGHTING panel. The anticollision lights are powered through the L and R ANTI-COLLISION LT circuit breakers on the aft J-box. 3-6

LIMITATIONS For specific limitations, refer to the FAA-approved AFM.

EMERGENCY/ ABNORMAL For specific information on emergency/abnormal procedures, refer to the appropriate checklist or FAA-approved AFM.

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QUESTIONS 1. When the LANDING–RECOG TAXI switch is in the LANDING position, it provides: A. The dimmest lighting available for taxiing B. Significantly more lamp life C. The brightest lighting available for landing D. The best lighting available for deplaning 2. When the nose baggage compartment doors are closed, the: A. Compartment light extinguishes regardless of the rocker switch position B. Compartment light remains illuminated C. White CAS message remains in the steady ON state D. White CAS message flashes in reverse video 3 LIGHTING

3. In the event of activation, the 5-G switch sends power to: A. The light above the cabin door B. The right table light above the emergency exit door C. The cockpit floodlight D. Both A and B 4. If left illuminated, the cabin entry light: A. Remains illuminated B. Should extinguish after 10 minutes C. Begins to flash D. Changes its shading to red

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CHAPTER 4 MASTER WARNING SYSTEM CONTENTS INTRODUCTION.................................................................................................................. 4-1 GENERAL................................................................................................................................4-1 DESCRIPTION....................................................................................................................... 4-2 CONTROLS AND INDICATIONS........................................................................................ 4-2 CAS Messages................................................................................................................. 4-2 Master Indicator Lights................................................................................................... 4-5 Rotary TEST Knob.......................................................................................................... 4-6 LIMITATIONS........................................................................................................................ 4-7 EMERGENCY/ABNORMAL................................................................................................ 4-7 QUESTIONS.......................................................................................................................... 4-9

ILLUSTRATIONS 4-1.

CAS Window on MFD in Normal Display Mode................................................... 4-3

4-2.

CAS Window on PFD in Reversionary Display Mode............................................ 4-3

4-3.

CAS Scroll Buttons................................................................................................. 4-4

4-4.

Master Indicator Lights........................................................................................... 4-6

4-5.

Rotary TEST Knob.................................................................................................. 4-6

TABLES Table Title Page 4-1.

ROTARY TEST INDICATIONS (AIRCRAFT CONFIGURATION AG).............. 4-8

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CHAPTER 4 MASTER WARNING SYSTEM

This chapter describes the master warning systems on the Citation Mustang aircraft. The master warning systems provide a warning of aircraft system malfunctions, indications of unsafe operating conditions requiring immediate attention, and indications that some specific systems are in operation. Audio warnings provide further indications.

GENERAL The master warning system includes a pair of MASTER WARNING and MASTER CAUTION lights, and the crew alerting system (CAS) messages. CAS messages are displayed by the G1000 engine indicating and crew alerting system (EICAS). These lights and messages provide a visual indication to the pilots of certain faults, functions, and/or conditions of selected systems. Each CAS message

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indicates an individual system situation, or simultaneous situations on both sides (left and right) of a dual system. Additionally, an audio warning system provides indication of some situations. CAS messages appear on the EICAS, which is on the left side of the multifunction display (MFD). In the event of a PFD or MFD failure, the pilot can press the red DISPLAY BACKUP button on

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each audio panel. This selects reversionary mode which combines displays from the PFD and MFD (including CAS messages).

CONTROLS AND INDICATIONS

The combined displays appear on both the selected PFD and the MFD. When the MFD and a PFD are in reversionary mode, the EICAS displays the CAS messages on the upper-right section of both display panels.

The Mustang master warning system includes: • CAS messages • Master indicators: • MASTER WARNING lights • MASTER CAUTION lights • A ural warnings—Various audio warnings are incorporated into aircraft systems that warn of specific conditions and malfunctions.

NOTE Crew alerts associated with the autopilot, avionics, and engine fire warning/ suppression are displayed elsewhere in the displays or cockpit. The crew alerts are described in the corresponding chapters in this manual. The abnormal and emergency procedures in this section are keyed, where applicable, to the appropriate CAS messages. The rotary TEST knob provides testing for the master warning system.

DESCRIPTION The Mustang master warning system uses cockpit indications (visual and aural) to advise the crew of important warnings, cautions, and advisory information about the aircraft and its systems. 4 MASTER WARNING SYSTEM

Pressure sensors, temperature sensors, switches, and other devices detect conditions in the aircraft and its systems. This information is provided as analog signals or discrete digital signals to the G1000 system. These signals are received at the left and right Garmin integrated avionics unit (GIAs), and at the left and right Garmin engine/airframe interface units (GEAs). A detailed description of G1000 system architecture is provided in Chapter 16—“Avionics.” With the battery switch selected to EMER, only the left GIA and left GEA are powered; CAS messages requiring input from the right GIA or right GEA are inhibited and do not appear, regardless of aircraft or system condition.

Control of the system is provided by: • CAS scroll softkeys (up and/or down) • Rotary TEST knob

CAS MESSAGES The Mustang master warning system is primarily a function of the G1000 avionics system. The CAS displays warning, caution, and advisory messages in response to data from aircraft sensors and FADECs.

CAS Message Window All CAS messages are displayed in the CAS message window. The window is on either the MFD, or on one or both of the PFDs: • N ormal display mode—CAS message window is in the bottom left corner of the EICAS display, which appears on the left side of the MFD (Figure 4-1). • R eversionary display mode—CAS message window is on the right side of the PFD display (Figure 4-2).

CAS Message Types CAS messages are in one of three colors: red, amber, or white.

Red (Warning) Message Red indicates a warning (hazardous situation that requires immediate pilot corrective action).

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NOTE If a red CAS message is displayed and is not acknowledged by the crew, the message DOES NOT extinguish even if the problem is corrected by the system itself or by the crew. Red CAS messages must be acknowledged.

Amber (Caution) Message

Figure 4-1.   CAS Window on MFD in Normal Display Mode

Amber indicates a caution (abnormal or special situation that includes a possible need for future pilot corrective action). When an amber CAS message displays, it flashes in reverse video and the MASTER CAUTION light illuminates steady. Pressing the MASTER CAUTION light acknowledges the message, extinguishes the MASTER CAUTION light, and changes the CAS message to a steady ON state. Amber CAS messages are displayed until the situation is corrected. All amber CAS messages are grouped together below any red CAS messages in the CAS display window. Any new amber CAS message displays at the top of the amber CAS group.

If an amber CAS message is displayed, the CAS message extinguishes once the problem is corrected by the system or by the crew (regardless if the crew acknowledges the CAS by pressing the MASTER CAUTION light).

White (Advisory) Message Figure 4-2.   CAS Window on PFD in Reversionary Display Mode

When a red CAS message displays, it flashes in reverse video and the MASTER WARNING light flashes. Pressing the MASTER WARNING light acknowledges the message, extinguishes the MASTER WARNING light, and changes the CAS message to a steady ON state. Red CAS messages are displayed until the situation is corrected. All red CAS messages are grouped together at the top of the CAS display window. Any new red CAS message displays at the top of the red CAS group.

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White indicates an advisory (advisory in nature and denote items which are considered normal during operation of the aircraft or do not normally require any pilot action). When a white CAS message displays, it appears in a steady ON state. All white CAS messages are grouped together below any red or amber CAS messages in the CAS display window. Any new white CAS messages display at the top of the white CAS group.

CAS Message Type Priority Some CAS messages may appear in different colors at different times, indicating different conditions or levels of severity (warning, caution, or advisory).

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For instance, the CABIN ALT message may appear in red, amber, or white, depending upon various causes and conditions. However, the same CAS message does not appear in two colors at the same time for the same side. Each CAS message appears in its highest-priority color appropriate at the time.

CAS Message Display Sequence CAS messages are sorted by color, and then by the order in which they have been caused: • Red (warning) CAS messages display at the top of the CAS message window, with the most recent red message at the top of the red message group. • Amber (caution) CAS messages display after any red CAS messages in the CAS message window, with the most recent amber message at the top of the amber message group. • White (advisory) CAS messages display after any red and/or amber CAS messages in the CAS message window, with the most recent white message at the top of the white message group. If a message changes color (due to changing severity of the condition/indication), the appropriate master indicators will appear and the message moves to the top of the list of the appropriate color group. 4 MASTER WARNING SYSTEM

CAS Message Scrolling The CAS message window can only display 14 messages. To ensure all valid CAS messages can be seen when appropriate, the CAS message window uses a “scrolling list” of all the current valid CAS messages. If more than 14 messages are valid, a scroll bar appears on the right side of the CAS message window (Figure 4-3). A slider in the scroll bar indicates which portion of the list is currently visible.Red warning messages always stay at the top of the CAS messages and are unaffected by scrolling. Amber caution messages can run off of the display. If an amber message scrolls off, or if there are too many messages to display all active amber messages, the scrollbar track changes color to amber to indicate to the pilot that an amber CAS message is currently off the viewable portion of the CAS window. 4-4

Figure 4-3.   CAS Scroll Buttons

At the same time, CAS-scroll buttons (one with an up-arrow and one with a down-arrow) appear on the soft keys at the bottom of the display, allowing the pilot to scroll up or down through the current list of CAS messages. When the top of the list is visible, only the CAS down-arrow scroll button appears among the soft keys. When the bottom of the list is displayed, only the CAS up-arrow scroll button appears.

CAS Message Inhibits Many CAS messages are inhibited (prevented) from displaying during certain phases of aircraft operation, regardless of whether the CAS message would otherwise be valid or not. These inhibits reduce pilot workload or prevent invalid indications during certain phases of aircraft operation. Different CAS messages are inhibited by different phases of aircraft operation. There are six phases of aircraft operations that inhibit various CAS messages: • Engine start inhibit (ESI) • Takeoff operational phase inhibit (TOPI) • On ground/in flight (GROUND/ AIR) • Landing operational phase inhibit (LOPI) • Engine shutdown inhibit (ESDI) • Emergency power mode (EMER)

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When an engine is starting, the ESI for that engine inhibits some CAS messages for that engine and related systems. This inhibit is triggered by the fullauthority digital engine control (FADEC). Each FADEC only inhibits messages from its side.

Takeoff Operational Phase Inhibit The TOPI prevents messages from distracting the flightcrew during takeoff. TOPI becomes active when any of the following is true: • The aircraft transitions from on ground to in air. • Either indicated airspeed transitions from less than 50 knots to more than 50 knots. The TOPI becomes inactive when any of the following is true: • The aircraft has been in the air for more than 25 seconds. • Pressure altitude is more than 400 feet above the field elevation. The field elevation is the pressure altitude captured when the aircraft transitions from on ground to in air. • Either airspeed indication is less than 40 knots. • Both airspeed indications are invalid. • The TOPI has been active for more than 90 seconds.

On Ground/In Flight Messages with the GROUND inhibit do not appear when the aircraft is on the ground, and messages with the AIR inhibit do not appear when the aircraft is in flight. The inhibits are normally controlled by squat switch indications, but if that fails then GPS ground speed and air data computer (ADC) are alternate sources used by the CAS for setting these inhibits.

Landing Operational Phase Inhibit LOPI activates when any of the following occurs: • The aircraft transitions from in air to on ground.

• G lobal positioning system (GPS) altitude transitions below 400 feet above field elevation stored in the terrain database. LOPI becomes inactive when any of the following is true: • The aircraft has been on the ground for more than 25 seconds. • Either indicated airspeed is less than 40 knots. • Both indicated airspeeds are invalid. • The LOPI has been active for more than 90 seconds. • Global positioning system (GPS) altitude is greater than 600 feet above the field elevation stored in the terrain database.

Engine Shutdown Inhibit When either engine is shut down, or its FADEC determines that the engine has failed, the ESDI for that engine inhibits some CAS messages for that engine and related systems. This inhibit is triggered by FADEC (shutdown indication). Each FADEC only inhibits messages from its side.

Emergency Power Mode To prevent nuisance messages in emergency power mode, selected messages are inhibited when the emergency bus is powering the aircraft. Emergency bus status occurs when: • The battery switch is in the EMER position, and • Both generator bus input discretes have been lost. The left GIA and the left GEA receive power in emergency bus mode. However, most of the sensors required to indicate CAS messages are not powered, so most messages have the EMER inhibit.

MASTER INDICATOR LIGHTS The master indicator lights (red MASTER WARNING and amber MASTER CAUTION lights) illuminate to direct pilot attention to new CAS messages. A MASTER WARNING and MASTER CAUTION light is on the instrument panel above each PFD (directly in front of the pilot and copilot) (Figure 4-4). Each master indicator light has

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Engine Start Inhibit

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Amber MASTER CAUTION Lights When a new amber message appears in the CAS window, the amber MASTER CAUTION lights both illuminate steady. Figure 4-4.   Master Indicator Lights

an integral momentary-contact pushbutton switch. Pressing the light resets the light and acknowledges the CAS message.

NOTE When the aircraft is first powered on, any CAS messages that immediately appear do not cause the master indicator lights to illuminate. The CAS considers these messages to be already acknowledged.

Red MASTER WARNING Lights When a new red CAS message appears on the MFD or PFD(s), both red MASTER WARNING lights illuminate flashing. Illumination of the L–R ENGINE FIRE light(s) also triggers the MASTER WARNING lights.

4 MASTER WARNING SYSTEM

Pressing either red MASTER WARNING light closes the switch momentarily, resetting the light. This extinguishes the MASTER WARNING light and stops the CAS message flashing (the CAS message remains on until the condition that caused it is corrected).

Each MASTER CAUTION light is an integral momentary-contact pushbutton switch. Pressing either MASTER CAUTION light closes the switch momentarily, resetting the light, which extinguishes. If a new amber CAS message appears in the window, the MASTER CAUTION light illuminates again, and continues illuminated until reset. When a MASTER CAUTION light is reset, both MASTER CAUTION lights extinguish, and the flashing amber CAS message illuminates steady until its indicated condition stops or is corrected. If multiple amber CAS messages are flashing at the same time, pressing a MASTER CAUTION light once acknowledges them all; all stop flashing and illuminate steady. If all amber CAS message problems are solved before the MASTER CAUTION lights are reset, the MASTER CAUTION lights automatically extinguish.

ROTARY TEST KNOB The rotary TEST knob (Figure 4-6) is above the copilot PFD. Positioning the knob to ANNU causes the MASTER CAUTION lights, MASTER WARNING lights, and other lights to illuminate (Table 4-1). Illumination verifies only light emitting diodes (LEDs) lamp integrity.

If a new red message appears in the CAS window, the MASTER WARNING light flashes again, and continues flashing until reset. If two or more red CAS messages are flashing at the same time, pressing the MASTER WARNING light once acknowledges them all at the same time. If the condition that caused the CAS message is corrected or stops before resetting the MASTER WARNING lights, the CAS message and the red MASTER WARNING lights continue to flash until acknowledged by resetting the MASTER WARNING lights. This is to ensure the important CAS message is seen by the pilot, even if it is only indicating an intermittent condition. 4-6

Figure 4-5.   Rotary TEST Knob

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During rotary test, audio warnings are also tested and some other associated system indications appear.

LIMITATIONS For specific limitations, refer to the FAA-approved AFM.

EMERGENCY/ ABNORMAL

4 MASTER WARNING SYSTEM

For specific information on emergency/abnormal procedures, refer to the appropriate checklist or FAA-approved AFM.

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Table 4-1.   ROTARY TEST INDICATIONS (AIRCRAFT CONFIGURATION AG) POSITION FIRE WARN

INDICATIONS • Red L ENGINE FIRE and R ENGINE FIRE lights illuminate. • MASTER WARNING lights illuminate.

LANDING GEAR

• Three green gear downlock lights illuminate. • Red gear UNLOCK light illuminates. • Gear warning horn sounds. Alternates between pilot and copilot speakers.

CABIN ALT

• Red CABIN ALT message appears. • Amber CABIN ALT message appears. • MASTER WARNING/CAUTION lights illuminate.

STALL

• Amber STALL WARN FAIL message appears. • Stall warning tone sounds and alternates between pilot and copilot speakers. • Amber STALL WARN HTR message appears. • White STALL WARN HI message appears. • MASTER CAUTION lights illuminate.

FLAPS

• The flap indicator on the MFD is replaced with a red X for 3 seconds. • Amber FLAPS FAIL message appears. • Amber STALL WARN FAIL message appears for 3 seconds. • MASTER CAUTION lights illuminate. • The overspeed warning tone sounds and alternates between pilot and copilot speakers.

ANTI SKID

• Amber ANTISKID FAIL message appears for 6 seconds. • White NO TIRE SPINDOWN message appears for 6 seconds. • MASTER CAUTION lights illuminate.

4 MASTER WARNING SYSTEM

OVERSPEED

ANNU

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ILLUSTRATIONS

• MASTER CAUTION illuminates and cannot be cancelled. • MASTER WARNING illuminates and cannot be cancelled. • Autopilot mode control panel indicators illuminate. • Audio panel indicators illuminate. • Red DUMP illuminates on Cabin Dump switch. • Test audio tone sounds. • Amber STANDBY BATTERY DISCHARGE light illuminates (near STBY INST switch) (aircraft configuration AF).

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QUESTIONS 1. When a new red CAS message appears: A. The MASTER WARNING horn sounds B. The CAS message flashes in reverse video C. MASTER WARNING lights flash D. B and C

6. Red CAS messages cannot be: A. Scrolled B. Extinguished C. Reset D. Cancelled

2. After pressing the MASTER WARNING reset button: A. The MASTER WARNING lights extinguish B. The CAS message reverts to steady on state C. The MASTER WARNING horn silences D. A and B

7. If two or more red CAS messages post at the same time, depressing the MASTER WARNING light will: A. Cause the MASTER WARNING light to continue to flash for each red CAS message posted until all have been acknowledged B. Acknowledge them all at the same time C. Cause the MASTER WARNING horn to sound D. B and C

4. If the condition that caused a red CAS message is corrected before resetting the MASTER WARNING light: A. The MASTER WARNING horn silences B. The CAS message reverts to steady on state C. The MASTER CAUTION lights continue to flash D. The CAS message and MASTER WARNING lights continue to flash

8. If more than 14 CAS messages are posted in the MFD display: A. The Garmin system automatically deletes the less critical messages B. The MASTER WARNING lights continue to flash C. A scroll bar appears on the right side of the display window D. The MASTER CAUTION lights continue to flash 4 MASTER WARNING SYSTEM

3. If all amber CAS messages are resolved before the MASTER CAUTION lights are reset: A. The MASTER CAUTION lights remain illuminated B. The amber CAS message remains in reverse video C. The MASTER CAUTION lights automatically extinguish D. The MASTER WARNING lights continue to flash

5. When in reversionary mode, the CAS messages are displayed: A. On the right section of the PFD or MFD in reversionary mode B. On the MFD only C. On the copilot PFD only D. On the pilot PFD only

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CHAPTER 5 FUEL SYSTEM CONTENTS INTRODUCTION.................................................................................................................. 5-1 GENERAL................................................................................................................................5-1 FUEL STORAGE................................................................................................................... 5-2 Components..................................................................................................................... 5-2 FUEL DISTRIBUTION.......................................................................................................... 5-4 Components..................................................................................................................... 5-4 Controls and Indications.................................................................................................. 5-6 Operation......................................................................................................................... 5-7 LIMITATIONS........................................................................................................................ 5-9 EMERGENCY/ABNORMAL................................................................................................ 5-9

5 FUEL SYSTEM

QUESTIONS........................................................................................................................ 5-10

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ILLUSTRATIONS Figure Title Page 5-1.

Fuel Tank System..................................................................................................... 5-3

5-2.

NACA Scoop Fuel Vent........................................................................................... 5-4

5-3.

Fuel Tank Filler........................................................................................................ 5-4

5-4.

Sump Drain Valves.................................................................................................. 5-4

5-5.

Primary Ejector Pump............................................................................................. 5-5

5-6.

Fuel Display on MFD (EICAS Normal Mode)....................................................... 5-6

5-7.

Alternate Fuel Display (EICAS Reversionary Mode.............................................. 5-6

5-8.

Fuel Controls........................................................................................................... 5-7

5-9.

ENGINE FIRE and BOTTLE ARMED Lights....................................................... 5-7

5-10.

Fuel Tank Servicing................................................................................................. 5-7

5-11.

Grounding Point....................................................................................................... 5-7

TABLES Table Title Page FUEL SYSTEM CAPACITY.................................................................................. 5-2

5-2.

CAS MESSAGES.................................................................................................... 5-9

5 FUEL SYSTEM

5-1.

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CHAPTER 5 FUEL SYSTEM

INTRODUCTION This chapter presents information on the fuel system of the Citation Mustang. Integral fuel tanks in the left and right wing provide fuel storage. The fuel distribution system provides fuel to each engine from the corresponding wing tank. The fuel transfer system allows fuel to be transferred from one tank to the other. Crew alerting system (CAS) messages alert the pilot to fuel system emergency and abnormal situations. Information in this chapter is provided for the airframe fuel system upstream of the high-pressure engine driven fuel pump. For description and operation of the engine fuel system, refer to Chapter 7—“Powerplant.”

The Citation Mustang fuel system includes two integral wing fuel tanks. Each wing tank has a passive capacitance type fuel quantity system.

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The system consists of: • Independent dual channel digital signal conditioner (in the left aft wing fairing) • Five fuel probes in each wing tank • EICAS displays

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5 FUEL SYSTEM

GENERAL

CITATION MUSTANG PILOT TRAINING MANUAL

At the wing tip, outboard of each main wing tank is a smaller surge tank. This surge tank allows for fuel expansion and venting. Located at each wing tank root is an integral engine feed bay. Each engine feed bay houses a primary ejector pump and an electrically powered boost pump that provides fuel under pressure to the respective engine. The right side wing fuel system is identical to the left side, except for a transfer valve in the right engine feed bay and a fuel temperature probe in the left engine feed bay. Switches on the lower pilot tilt panel and CAS messages on the multifunction display (MFD) control and indicate fuel operation and transfer.

FUEL STORAGE COMPONENTS Main Tank Cavity Each main tank cavity (one in each wing, between the forward and rear wing spars) is integral to the wing. Holes in the main spar and ribs allow fuel flow through the wing (Figure 5-1). Flapper valves, attached to spar and rib holes, allow fuel flow inboard while inhibiting flow outboard. The left and right tanks each have a fuel capacity of 192.5 gallons (728.7 liters) for a total combined fuel capacity per aircraft of 385 gallons (1,457 liters). Refer to Table 5-1 for approximate volume and weights, or refer to the Airplane flight Manual (AFM) for current data. Due to the fuel oil heater (Refer to Chapter 7—“Powerplant”) the Citation Mustang fuel system does not require the use of anti-icing additive.

5 FUEL SYSTEM

Each wing tank system (left wing and right wing) includes: • Main tank cavity • Engine feed bay • Venting system • Tank filler • Sump drain valve • Scavenge pumps • Fuel probes • Flapper valves

5-2

Table 5-1.   FUEL SYSTEM CAPACITY STANDARD (U.S.)

METRIC

Total Capacity

Weight

Volume

Weight

Volume

Each Tank

1,290 pounds

192.5 gallons

585 kg

728.7 liters

Both Tanks

2,580 pounds

385 gallons

1,170 kg

1,457 liters

Fuel Gauging The fuel system has a capacitance probe quantity indication system that compensates for changes in density caused by temperature variations. Each fuel tank has five fuel quantity probes and a signal conditioner. One of the fuel quantity probes is in the engine feed bay and the other four extend outward from the engine feed bay to the outer tip of the fuel cell. These probes supply quantity information (in pounds) to the signal conditioner. The fuel quantity signal conditioners are in the wing fairing. They receive the quantity measurement from all five probes on the respective side and total the values. The total fuel quantity is then displayed to the pilot on the EICAS display.

Venting System A NACA scoop under each wing is the only ventilation source and allows air to enter or exit its respective vent surge tank (Figure 5-2). A floatcontrolled vent valve is connected to the vent surge tank. When the fuel level is full enough to raise the float, the valve closes preventing fuel from overflowing into the surge tank. When the fuel level is low, the valve opens to provide venting. The surge tank allows for fuel expansion and holds some fuel. As the fuel level decreases a flapper valve permits fuel to drain from the surge tank back into the main tank. If the vent surge tank fills, another vent allows spillage overboard through the NACA scoop. The NACA scoop does not require anti-icing.

Tank Filler The aircraft has one fuel tank filler assembly on the upper surface of each wing, between the main spar and the aileron (Figure 5-3). The filler assembly consists of a flush type cap and a standpipe.

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5 FUEL SYSTEM

GRAVITY SUCTION

VENT LINES

TRANSFER PRESSURE

LOW-PRESSURE, HIGH-VOLUME FLOW

TO ENGINE FUEL SYSTEM

FROM ENGINE FUEL SYSTEM

FUEL INSIDE TANK

LEGEND

PAR

REAR SPAR

OVERBOARD VENT LINE

AR MAIN SP

ARD S FORW

VENT VALVE

LEFT SURGE TANK

FUEL PROBE

LP

Figure 5-1.   Fuel Tank System

TEMPERATURE SENSOR

FIREWALL SHUTOFF VALVE

TRANSFER SHUTOFF VALVE

TS

LP LOWPRESSURE SWITCH

FIREWALL SHUTOFF VALVE

PRIMARY EJECTOR PUMP

FLAPPER VALVE

(FROM) (TO) R ENGINE FUEL SYSTEM

SCAVENGE EJECTOR PUMPS

(TO) (FROM) L ENGINE FUEL SYSTEM

REAR SPAR

ELECTRIC BOOST PUMP

AR MAIN SP

FUEL DRAIN

PAR

ARD S FORW

FLOW-RESTRICTOR ORIFICE

FUEL FILLER CAP

VENT SCOOP

SURGE VENT LINE

LEFT MAIN FUEL TANK

SURGE VENT LINE

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Figure 5-2.   NACA Scoop Fuel Vent Figure 5-4.   Sump Drain Valves

Engine Feed Bay An engine feed bay is an integral part of each wing tank. It is the lowest point in the fuel system, which is the location for fuel pick up (in each tank) for the fuel distribution system. Each feed bay holds approximately 8 gallons and has four vent openings (to ensure the bay maintains full capacity under all flight conditions). These openings are restricted by flapper valves that allow the fuel to freely flow into the engine feed bay but restrict the fuel flow out of the engine feed bay. The purpose of the engine feed bay is to keep the boost pump and primary ejector pump submerged even under low fuel conditions. Figure 5-3.   Fuel Tank Filler

5 FUEL SYSTEM

The standpipe is an extension from the filler cap opening into the tank and is used as the full indicator when refueling.

FUEL DISTRIBUTION

Sump Drain Valves

Electric Boost Pumps

Sump drain valves are at the low point in each wing where water can collect (Figure 5-4). In each wing, there is a sump drain in the following locations: • Outboard of the landing gear (behind the main spar) • In the engine feed bay • Between the feed bay and the main spar • Between the forward and main spars (forward of the feed bay) When draining sumps, do not turn any tool in the drain. The drain may lock open resulting in fuel loss.

One 28 VDC boost pump is in each engine feed bay. Either boost pump can be controlled automatically or manually. Circuit protection for the boost pumps is in the aft J-box.

5-4

COMPONENTS

Electric boost pumps are used for: • Engine starting • Low fuel supply pressure • Fuel transfer

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Primary Ejector Pumps

Firewall Shutoff Valves

The primary ejector pump is submerged in fuel in each engine feed bay (Figure 5-5). The primary ejector pump has no moving parts. The pump utilizes a small jet of high-pressure “motive flow” fuel (from the respective high-pressure engine driven fuel pumps). The motive flow fuel passes through a venturi which creates low-pressure in the ejector pump. This low-pressure allows the ejector pump to pull a large low-pressure flow of fuel from the feed bay and pumps it back to the engine. Some of the resulting flow also provides motive flow to the scavenge ejector pumps.

Firewall shutoff valves for each engine are in the respective aft wing fairing (between the wing and fuselage). In the event of a fire, the valve shuts off fuel flow to the respective engine on pilot command. The valve can be commanded closed by the FADEC in the event the normal shutdown valve fails. Circuit protection for the shutoff valves are provided by the L FEED BUS #2 and R FEED BUS #2 through the respective L and R FIREWALL CUTOFF circuit breakers on the aft J-box.

Scavenge Ejector Pumps Two scavenge ejector pumps constantly transfer fuel from the forward and outboard areas of each tank to its engine feed bay. This ensures the engine feed bay is full and the primary ejector and electric boost pump are submerged in fuel until the wing tank is nearly empty. Each scavenge ejector pump receives low-pressure motive flow from the same side primary ejector pump or (when operating) the same side electric boost pump.

Fuel Transfer Valve The fuel transfer valve is in the right engine feed bay. The valve is a direct-acting solenoid that requires normal DC power, to open and with a loss of DC power the valve fails in the closed position. The fuel transfer valve is not powered with a loss of both generators and the BATT switch placed in the EMER position. Circuit protection for the transfer valve is provided by the L and R FUEL CONTROL circuit breaker in the aft J-box. EJECTOR PUMP

Low Fuel Pressure Switches Pressure switches are in the engine fuel supply lines adjacent to each engine. The switch monitors fuel pressure between the engine feed bay and the high-pressure engine-driven fuel pump. The switches can automatically activate the electric boost pumps if the boost pump switches are in the NORM position. The switches deactivate the electric boost pump at 6.4 psig (maximum) and reactivate it when the engine fuel supply pressure drops below 4.65 psig. When fuel pressure drops below this limit, the FUEL PRES LO L-R CAS message appears.

Fuel Flow Transmitter A fuel flow transmitter is on each engine fuel supply line. The transmitter sends a 0–5 volt analog signal to the G1000 system, which translates the signal to pounds/kilograms per hour. HIGH PRESSURE FUEL MOTIVE FLOW (FROM ENGINE DRIVEN FUEL PUMPS)

TO ENGINE DRIVEN FUEL PUMPS

FUEL INLET

5 FUEL SYSTEM

HIGH VOLUME FUEL

Refer to Chapter 8—“Fire Protection” for more information on the firewall shutoff valves and their operation.

FUEL IN TANK

Figure 5-5.   Primary Ejector Pump

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CONTROLS AND INDICATIONS

Fuel Flow Indication

Fuel Quantity Indication

Fuel flow is displayed digitally below the total fuel display. The digits are green and are displayed in pounds per hour (PPH) or kilograms per hour (KGH). Invalid data displays as a red X.

Fuel quantity is displayed on the MFD “AUX” page and in the left EICAS column on the MFD (Figure 5-6). Quantity can be displayed in either pounds or kilograms. Fuel tank levels are displayed with a white pointer on a white scale on the fuel display and by green digits just below the scale. In reversionary mode, only the digits are displayed (Figure 5-7). Total aircraft fuel is the sum of the fuel quantities displayed for each tank. This value is displayed below the individual tank quantities in green digits. Total fuel quantity is displayed in the same units as the fuel tank levels. Invalid data is displayed by a red X or white dashes.

Fuel Temperature Indication The fuel temperature probe is inside the left engine feed bay. The fuel temperature appears at the bottom of the fuel display window as green digits in Celsius (see Figure 5-6). If invalid data is received, a red X is displayed. The temperature displayed is invalid if below –70°C (–94°F) or above 99°C (210°F).

FUEL BOOST Switches The FUEL BOOST switches are on the lower instrument tilt panel (Figure 5-8). Each switch (L and R) has three positions; ON, OFF, and NORM. The switches manually control the respective boost pumps in the ON and OFF positions. In the NORM position, boost pump operation is automatically controlled. The FUEL BOOST L-R or FUEL BOOST L-R CAS message indicates the boost pump is on and operating.

FUEL TRANSFER Selector Knob

Figure 5-6.   Fuel Display on MFD (EICAS Normal Mode)

5 FUEL SYSTEM

Figure 5-7.   Alternate Fuel Display (EICAS Reversionary Mode

5-6

A FUEL TRANSFER selector knob on the lower left instrument tilt panel controls the fuel transfer valve (Figure 5-8). With the FUEL BOOST switches in the NORM position, the FUEL TRANSFER selector knob also commands the boost pumps ON or OFF. The FUEL TRANSFER selector knob has three positions; L TANK, OFF, and R TANK. When the selector is placed in the L TANK or R TANK position, it energizes the fuel transfer valve open and (with the boost pump switches in NORM) energizes the fuel boost pump on the supply side. This allows fuel to be picked up by the boost pump on the supply side engine feed bay and transfer it to the opposite engine feed bay. The transferred fuel comes out through the inoperative boost pump in the receiving engine feed bay. Fuel is transferred at approximately 10 ppm (4.5 kg per minute). Rate varies with engine(s) fuel flow.

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OPERATION Fuel Servicing Fuel servicing includes procedures necessary for fueling and procedures used to check the fuel for contamination or condensation. The fuel is serviced through the flush type cap on the outboard section of either wing (Figure 5-10).

Figure 5-8.   Fuel Controls

ENGINE FIRE Light Figure 5-10.   Fuel Tank Servicing

Refueling Refuel in areas that permit the free movement of fire equipment. Follow approved ground procedures for the aircraft and the fueling equipment. There is one approved grounding point under the outboard end of each wing (Figure 5-11).

5 FUEL SYSTEM

In the event of an engine fire, the fire detector in the engine compartment illuminates the red L or R ENGINE FIRE lights (Figure 5-9). Pushing the light (which has an integral pushbutton switch) closes the corresponding firewall shutoff valve, which shuts off the fuel flow to the engine and illuminates the F/W SHUTOFF L–R CAS message. Refer to Chapter 8—“Fire Protection” for more information.

Figure 5-11.   Grounding Point Figure 5-9.   ENGINE FIRE and BOTTLE ARMED Lights

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Refuel to the bottom of the standpipe to achieve maximum use for flight planning. If the fuel tank is filled above the bottom of the standpipe, there may not be room for expansion which can result in fuel spillage through the fuel vent. Approved fuels and additives for operation of the aircraft are listed in the “Limitations” section of the AFM. Use of avgas is not approved.

Defueling Defueling must be performed as a maintenance function.

Fuel Distribution During normal operation of the fuel system, the L and R FUEL BOOST pump switches are in the NORM position. In this position, each boost pump operates automatically: • D uring engine start— FUEL BOOST L–R CAS message appears. • D uring fuel transfer operation— FUEL TRANSFER and FUEL BOOST L–R CAS messages appear. • W hen low fuel pressure is sensed in the engine fuel supply line—The FUEL PRES LO L–R CAS message appears for a moment, followed quickly by the FUEL BOOST L–R CAS message. As the boost pump increases the fuel pressure, the FUEL PRES LO L–R CAS message extinguishes. If the throttle is in the CUT OFF position the boost pumps do not energize automatically in a low fuel pressure condition, even though the FUEL BOOST switch is in the NORM position. When the switch is OFF, the boost pump does not operate. In the ON position, the pump operates continuously. 5 FUEL SYSTEM

With the L and R FUEL BOOST pump switches in the NORM position, pressing an ENGINE START button energizes the corresponding fuel boost pump. This moves fuel from the wing tank engine feed bay on that side through the firewall shutoff valve to the engine driven fuel pump on the respective engine.

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When the engine start terminates, the boost pump is deenergized and the FUEL BOOST L–R CAS message disappears from the CAS window. During normal operation, each engine is supplied with fuel from the primary ejector pump in the engine feed bay of each tank. The electric boost pump (when energized automatically or by pilot command) may augment the operation of the ejector pump.

Fuel Transfer Operation Using the fuel transfer system, fuel is transferred from one engine feed bay to the opposite engine feed bay. The arrow on the FUEL TRANSFER selector knob points to the wing tank where transfer fuel is directed. Rotating the FUEL TRANSFER selector knob from the OFF position to the R TANK position: • E nergizes the left tank electric boost pump if the FUEL BOOST switch is in the NORM position and displays the FUEL BOOST L CAS message. • Energizes the fuel transfer valve open. The FUEL TRANSFER CAS message is displayed. The left tank boost pump pressure supplies fuel from the left wing tank engine feed bay through the transfer valve and into the right wing tank engine feed bay. Check that the FUEL BOOST L–R CAS message displays indicate the correct boost pump is energized. If both boost pumps are energized, fuel transfer does not occur. To de-energize the pump in the non-selected tank, cycle its L or R FUEL BOOST switch to OFF, then ON, then NORM, and leave in NORM position. To verify fuel transfer, monitor the fuel quantity white tape pointer or the digital indicators (see Figure 5-12). Fuel normally transfers to the selected tank at approximately 10 ppm (600 pph). Maximum normal fuel imbalance is 200 pounds. Maximum emergency fuel imbalance is 600 pounds.

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To terminate fuel transfer and return the system to normal operation, rotate the FUEL TRANSFER selector knob to OFF. The electric boost pump deenergizes (if the FUEL BOOST switch is in the NORM position), the FUEL TRANSFER CAS message disappears, and the fuel transfer valve spring-loads closed.If electrical power fails during fuel transfer operation, the fuel transfer solenoid valve returns to the closed position, preventing fuel transfer.

Table 5-2.   CAS MESSAGES FUEL BOOST L-R DESCRIPTION

INHIBITS

INHIBITS

For specific limitations, refer to the FAA-approved AFM.

Engine fuel filter impending bypass switch is closed. EMER

FUEL LVL LO L-R DESCRIPTION

EMERGENCY/ ABNORMAL

EMER

FUEL FLTR BP L-R DESCRIPTION

LIMITATIONS

Indicates left and/or right low fuel pressure is detected and the boost pumps automatically turn on.

INHIBITS

This message indicates the fuel level is too low. EMER

FUEL PRES LO L-R

For specific information on emergency/abnormal procedures, refer to the appropriate checklist or FAA-approved AFM.

DESCRIPTION INHIBITS

Fuel pressure is under 4.65 psig (decreasing) or 6.4 psig (increasing) EMER, ESDI

FUEL BOOST L-R DESCRIPTION

INHIBITS

Appears when the pilot commands the fuel boost pump on (by selected the FUEL TRANSFER knob to L TANK or R TANK, or by sleected a FUEL BOOST L-R switch to ON, or during engine start. EMER

FUEL LO INOP L-R DESCRIPTION

INHIBITS

Indicates the amber FUEL LVL LO L-R CAS message is not operational and cannot provide reliable indication of fuel level. LOPI, TOPI

DESCRIPTION INHIBITS

Revision 1.1

FOR TRAINING PURPOSES ONLY

Indicates the fuel transfer valve is open. EMER

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5 FUEL SYSTEM

FUEL TRANSFER

CITATION MUSTANG PILOT TRAINING MANUAL

QUESTIONS 1. With the FUEL FLTR BP L message in the CAS window displayed, the best answer would be to: A. Land as soon as practical B. Transfer fuel to the effected engine C. Primary Ejector pump—AUX D. Auxiliary Ejector pump—STANDBY 2. The FUEL BOOST L CAS message indicates that: A. The left fuel boost pump is operating because the fuel boost switch was placed to the ON position B. The primary ejector pump is operating C. The engine-driven high-pressure pump is not operating or has failed D. The left fuel boost pump is automatically operating because of low fuel pressure 3. If the F/W SHUTOFF R message in the CAS window posts when the throttle is moved to CUTOFF, this indicates that: A. The primary ejector pump is inoperative B. A malfunction of the normal shutdown system has occurred C. The electric boost pump is inoperative D. The transfer valve is closed 4. The FUEL TRANSFER message in the CAS window posts: A. During fuel transfer operations B. When the primary flapper valves are open C. When the primary ejector flow control valve is open D. When the FUEL LVL LO message is displayed in the CAS window

5 FUEL SYSTEM

5. With the FUEL LO INOP L message in the CAS window posted: A. The left boost pump has failed B. The fuel quantity signal conditioner is unable to determine the fuel low level C. The fuel quantity is 170 pounds or less D. The primary check valve has failed

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6. The correct statement is: A. The FUEL BOOST pump switches must be ON for engine start B. With both FUEL BOOST pump switches on, fuel can be transferred C. The fuel boost pump automatically energizes anytime the FUEL BOOST switches are in NORM and the START button is depressed, FUEL TRANSFER is selected, or low-pressure is sensed in the engine supply line D. All of the above are correct 7. After engine start, the fuel boost pump is deenergized by: A. The FUEL BOOST pump switch B. Start circuit termination C. Discontinuing fuel transfer D. A time-delay relay 8. To verify that fuel transfer occurs, it is necessary to: A. Monitor the fuel quantity indicators for appropriate quantity changes B. Observe that only the white FUEL TRANSFER annunciator illuminates C. Ensure that both white FUEL BOOST ON annunciators illuminate D. Ensure that the FUEL BOOST pump switch for the tank being fed illuminates 9. With the FUEL BOOST switches in the NORM position and fuel transfer is commanded by placing the FUEL TRANSFER selector knob to the LEFT position, what are the indications: A. Only a FUEL TRANSFER CAS message B. FUEL TRANSFER and FUEL BOOST R CAS messages C. FUEL TRANSFER and FUEL BOOST L CAS messages D. FUEL TRANSFER and FUEL BOOST L CAS messages

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10. Operation of the primary ejector pump directly depends upon: A. DC electrical power B. High-pressure fuel from the engine-driven fuel pump (motive flow) C. AC electrical power supplied by the No. 1 or No. 2 inverter D. Flow from the transfer ejector pump



11. If the engine-driven fuel pump fails: A. The primary ejector pump also fails, but the boost pump energizes by low-pressure and sustains the engine B. The engine flames out C. The transfer ejector pump also is inoperative D. Right or left FUEL TRANSFER must be selected to obtain high-pressure motive flow from the opposite engine 12. The emergency asymmetric fuel differential is: A. 600 pounds B. 400 pounds C. 1,000 pounds D. 200 pounds

5 FUEL SYSTEM

13. Fuel is transferred at a rate of approximately: A. 50 pounds/minute B. 40 pounds/hour C. 10 pounds/minute D. 10 gallons/hour

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5 FUEL SYSTEM

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6   AUXILIARY POWER SYSTEM

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The information normally contained in this chapter is not applicable to this aircraft.

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CHAPTER 7 POWERPLANT CONTENTS GENERAL................................................................................................................................7-1 DESCRIPTION....................................................................................................................... 7-2 Turbofan Engine Basics................................................................................................... 7-3 COMPONENTS..................................................................................................................... 7-4 Engine Systems and Accessories..................................................................................... 7-4 Nacelles And Covers....................................................................................................... 7-5 Engine Systems............................................................................................................... 7-5 Oil System....................................................................................................................... 7-7 CONTROLS AND INDICATIONS...................................................................................... 7-11 FADEC.......................................................................................................................... 7-11 FADEC Reset Switch.................................................................................................... 7-11 THROTTLES................................................................................................................ 7-12 L AND R IGNITION Switches..................................................................................... 7-12 ENGINE START Switches........................................................................................... 7-13 ENGINE SYNC Switch................................................................................................ 7-13 Engine Indicating and Crew Alerting System............................................................... 7-13 N1% Window................................................................................................................. 7-15 N1% RPM...................................................................................................................... 7-15 ITT and Ignition Window.............................................................................................. 7-16 Oil Pressure (PSI) and Temperature (°C) Indications................................................... 7-16

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7 POWERPLANT

INTRODUCTION.................................................................................................................. 7-1

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OPERATION........................................................................................................................ 7-17 Preflight......................................................................................................................... 7-17 START........................................................................................................................... 7-18 Ground Operation.......................................................................................................... 7-18 Flight Operations........................................................................................................... 7-18 7 POWERPLANT

LIMITATIONS...................................................................................................................... 7-19 EMERGENCY/ABNORMAL.............................................................................................. 7-19 QUESTIONS........................................................................................................................ 7-20

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ILLUSTRATIONS 7-1.

Mustang Engine Mounting...................................................................................... 7-2

7-2.

Engine Stations........................................................................................................ 7-2

7-3.

Engine Schematic/Cutaway..................................................................................... 7-3

7-4.

Turbofan Engine Basics........................................................................................... 7-4

7-5.

Turbofan Engine Basics........................................................................................... 7-5

7-6.

Oil Servicing Panel.................................................................................................. 7-5

7-7.

Ignition System (Left Engine Battery Start............................................................. 7-6

7-8.

Engine Fuel System................................................................................................. 7-8

7-9.

Oil System Schematic.............................................................................................. 7-9

7-10.

Oil System............................................................................................................. 7-10

7-11.

Oil Filler Port and Sight Glass Gauge................................................................... 7-10

7-12.

Oil Filter Bypass Indicator..................................................................................... 7-10

7-13.

FADEC/Avionics Interface.................................................................................... 7-11

7-14.

FADEC Switch...................................................................................................... 7-12

7-15.

Throttle Quadrant.................................................................................................. 7-12

7-16.

ENGINE START and IGNITION Switches.......................................................... 7-13

7-17.

ENGINE SYNC Switch......................................................................................... 7-13

7-18.

EICAS Display on MFD........................................................................................ 7-13

7-19.

EICAS Reversionary Display on PFD................................................................... 7-14

7-20.

EICAS Analog Markings....................................................................................... 7-14

7-21.

EICAS Display - Invalid Data............................................................................... 7-14

7-22. N1% Window......................................................................................................... 7-15 7-23.

ITT and Ignition Window...................................................................................... 7-16

7-24. N2% Window......................................................................................................... 7-16 7-25.

Oil Pressure and Temperature Window (Normal).................................................. 7-17

7-26.

Oil Pressure and Temperature Window (Reversionary)......................................... 7-17

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7 POWERPLANT

Figure Title Page

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7 POWERPLANT

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TABLES Table Title Page TARGET N1 BUG................................................................................................. 7-15

7-2.

CAS MESSAGES.................................................................................................. 7-19 7 POWERPLANT

7-1.

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7 POWERPLANT

CHAPTER 7 POWERPLANT

INTRODUCTION This chapter describes the Citation Mustang powerplants, including the engines and their subsystems. The Mustang is powered by two turbofan engines. Each powerplant includes ignition, oil, and fuel systems. This chapter also describes powerplant controls and indicating systems.

GENERAL The Mustang is powered by two Pratt & Whitney PW615F turbofan engines (Figure 7-1). Each Mustang powerplant installation includes a fuel metering unit (FMU), an accessory gear box (to drive accessories with engine power), and ports to provide bleed air for the environmental control system (ECS) and ice-protection systems. A remotely located dual-channel full-authority digital engine control (FADEC) monitors and controls each engine. The two FADECS are in the tail cone on the aft pressure bulkhead. FADECs adjust engine settings in response to pilot throttle settings,

ambient air conditions, and engine conditions to provide optimum engine performance. A dualcoil, permanent-magnet alternator (integral to the FMU) powers each engine FADEC when normal DC power is not available. Each powerplant includes ignition, fuel, and oil systems. Engine indications are integrated into the G1000 electronic cockpit displays. This chapter includes information on normal engine operations (including starting, ground operation, and flight), powerplant limitations, and emergency/abnormal procedures.

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Fire detection and extinguishing systems for the powerplant are described in Chapter 8—“Fire Protection.”

DESCRIPTION 7 POWERPLANT

Two Pratt & Whitney PW615F turbofan engines are in nacelles mounted on pylons on each side of the tail cone. Each engine is flat-rated at 1,460 pounds of maximum continuous thrust (sea level static, standard day). Engine station numbers are assigned at particular points to locate various components and functions, usually relating to air temperature and pressure (Figure 7-2). The PW615F is a twin-spool, counter-rotating turbofan engine (the N1 spool and N2 spool rotate in opposite directions). It has a single-stage, low-pressure axial turbine that directly drives a single-stage, high-efficiency fan. A single-stage, high-pressure axial turbine drives a single-stage, mixed-flow compressor and a single-stage centrifugal compressor (Figure 7-3). Figure 7-1.   Mustang Engine Mounting

1 2 STATIC DYNAMIC

3 COMPRESSED

4 TURBINE INLET

5 TURBINE OUTLET

6 EXHAUST

LEGEND INDUCTION AIR AXIAL COMPRESSOR CENTRIFUGAL COMPRESSION AIR COMBUSTION CHAMBER TURBINE AIR EXHAUST AIR

Figure 7-2.   Engine Stations

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Figure 7-3.   Engine Schematic/Cutaway

TURBOFAN ENGINE BASICS

Combustion

Turbofan engines provide thrust from two sources: the fan and the high-speed engine exhaust. The fan provides thrust like a propeller, pulling air into the fan and pushing it aft. The mass of exhaust exiting aft from the engine at high speed and pressure creates an equal reaction, pushing the engine (and the airplane) in the opposite direction (forward).

The compressed air enters the combustion chamber where it is mixed with fuel and ignited. At engine start, electric ignitors create sparks that ignite the mixture. After each engine start, the flame in the combustion chamber continues burning as long as fuel and air are supplied. The burning fuel/air mixture creates hot, high-pressure exhaust, which expands rapidly and moves aft through the engine.

The core of the engine operates a continuous sequence of air intake, air compression, fuel/air mixture combustion, and exhaust. The exhaust turns turbines that provide torque to the fan and continuous air compression for the engine core. The compressed air is mixed with fuel and ignited, resulting in rapid expansion. The exhaust then exits the engine at high speed to provide the additional thrust (Figure 7-4).

Intake and Compression The spinning fan pulls ambient air into the engine inlet and sends some of it through the fan bypass duct for direct thrust. The fan also pushes air into the compressor section, where the axial-flow and centrifugal-flow compressors compress the incoming air to a very high pressure and temperature.

Revision 1.0

Exhaust As the hot, high-pressure exhaust moves aft through the engine, it turns the high-pressure turbine. The high-pressure turbine is connected to the compressors through a short, hollow shaft. The high-pressure rotors (turbine, shaft, and compressors) are referred to as the “high-pressure (HP) spool.” Its rate of rotation is referred to as “N2 rpm” or simply “N2.” N2 rotation keeps the airflow entering the engine and maintains the intake/compression/ combustion/exhaust cycle. A thermocouple harness at engine Station 6 measures exhaust stream temperature. This information is processed by the FADEC and converted to an equivalent interstage turbine temperature (ITT) for use by the pilot.

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LOW-PRESSURE COMPRESSOR MIXED-FLOW ROTOR

FAN BYPASS DUCT

HIGH-PRESSURE COMPRESSOR (CENTRIFUGAL-FLOW, SINGLE-STAGE)

HIGH-PRESSURE TURBINE

7 POWERPLANT FAN

COMBUSTION CHAMBER

LOW-PRESSURE TURBINE

LEGEND AMBIENT-AIR AND FAN-BYPASS AIR

BURNING FUEL

LOW-PRESSURE COMPRESSED AIR

HOT, SLOW, HIGH-PRESSURE EXHAUST

HIGH-PRESSURE COMPRESSED AIR

HOT, FAST, LOW-PRESSURE EXHAUST

Figure 7-4.   Turbofan Engine Basics

After exiting the HP turbine, the exhaust (now at lower pressure and temperature, but at higher speed) continues through the low pressure (LP) turbine, turning it. The LP turbine turns a long, narrow, inner shaft (which passes through the hollow HP spool) to directly drive the fan. The LP rotors (LP turbine, inner shaft, and fan) are referred to as the “LP spool.” Its rate of rotation is referred to as “N1 rpm” or simply “N1.” The exhaust loses some heat as it turns the turbines, and then mixes with the fan bypass air before exiting aft through the engine exhaust nozzle. As the engine moves the exhaust aft and out, it produces jet propulsion thrust. Jet propulsion thrust and fan bypass thrust combine to produce total engine thrust.

COMPONENTS ENGINE SYSTEMS AND ACCESSORIES On the bottom of the engine is an accessory gearbox (AGB) with an integral oil reservoir, pump, and mechanical power connections for enginedriven accessories. The AGB is driven by the HP spool (N2) through a gear-driven shaft. The AGB drives the engine fuel pump and its associated alternator. The AGB also connects the starter-generator to the engine. Two ports on the outside of the front bypass duct allow for “bleed off” of HP P3 air. This air is used in conjunction with the environmental and iceprotection systems. To prevent engine surge, a bleed valve actuator (BVA) controls pressure in the compressor section of the engine. Compressor surge is managed by bleeding off pressure as required during the differ-

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ent phases of operation to the fan bypass duct. This process is controlled by the FADEC throughout the engine operating envelope.

The engine nacelles (cowlings) are aluminum and consist of the inlet, upper, and lower nacelle doors, and the aft nacelle assembly. The leading edge of the inlet is heated with engine bleed air for antiicing purposes (see Figure 7-1).

VISUAL INSPECTION HOLE

CAUTION When engine anti-ice is operated, and for some time after, the nacelle leading edge and starter-generator cooling inlet may be extremely hot and cause burns to skin. Avoid direct contact. The upper and lower nacelle doors are attached using quarter-turn fasteners, which allow for quick access to the engines for maintenance or inspection (Figure 7-5).

Figure 7-5.   Turbofan Engine Basics

On the lower outboard side of each of the lower nacelles is an oil door, which provides the crew with easy access to the oil level sight glass (Figure 7-6). A spring-loaded closed door is provided on each lower nacelle (outboard on the right nacelle and inboard on the left nacelle), allowing for a visual inspection of the oil filter bypass indicator.

Figure 7-6.   Oil Servicing Panel

Engine contamination is possible from many sources and may cause engine damage. These sources include: • Hail • Condensation and freezing • Salt water spray • Blowing sand • Dirt, dust, or volcanic ash • Birds • Insects • Leaves • Other debris To prevent contamination of the engine on the ground when the engines are off, engine covers are provided for the inlet and exhaust ports of each engine.

ENGINE SYSTEMS Ignition System Each engine has dual ignitors, which produce sparks to ignite fuel in the engine combustion chambers. They are powered by exciter boxes and controlled by the FADEC. Ignitors are normally only operated during starting (Figure 7-7).

FOR TRAINING PURPOSES ONLY

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7 POWERPLANT

NACELLES AND COVERS

CITATION MUSTANG PILOT TRAINING MANUAL

7 POWERPLANT EXCITER BOX

L IGNITION NO. 1 SSR

L IGNITION NO. 2 SSR

7.5

LH IGN SSR NO. 1 (AFT J-BOX)

EXCITER BOX

L ENG L ENG FADEC FADEC A B

R IGNITION NO. 1 SSR 7.5

7.5

LH IGN SSR NO. 2 (AFT J-BOX)

RH IGN SSR NO. 1 (AFT J-BOX)

R IGNITION NO. 2 SSR 7.5

R ENG R ENG FADEC FADEC A B

LEGEND NORMAL DC POWER HIGH-ENERGY IGNITION

Figure 7-7.   Ignition System (Left Engine Battery Start

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RH IGN SSR NO. 2 (AFT J-BOX)

CITATION MUSTANG PILOT TRAINING MANUAL

The FADECs also command the ignitors on and off during an autorelight situation. Pilot control of the ignitors consists of two options: normal and on. In normal mode, the FADECs command the ignitors on and off as required. For the autorelight feature, the FADEC monitors fuel flow and N2 speed. If an uncommanded drop in N2 rpm lasts for more than 0.25 seconds, the FADEC activates the ignitors. Once a positive engine acceleration and adequate rise in ITT are detected, the ignitors are turned off by the FADEC. Along with fuel flow, the ignitors are commanded off by the FADEC when the throttles are put into the CUT OFF position.

Engine Fuel System The engine fuel system consists of the fuel system components between the firewall shutoff valves and the engine. An FMU, under the direction of the FADEC, regulates the fuel flow to the engines (Figure 7-8). The FMU also provides high pressure motive fuel flow to the fuel tank ejector pump system (refer to Chapter 5—“Fuel System”) and the fuel pressure is relayed to the FADEC which controls the bleed valve actuator (BVA).

Bleed Valve Actuator (BVA) The BVA allows for surge free operation of the engine. The BVA is pneumatically operated and is controlled electrically by the FADEC. The valve relieves excess compressed air that the centrifugal compressor cannot use at various power settings.

Fuel/Oil Heat Exchanger A fuel/oil heat exchanger (FOHE) is also part of the FMU assembly. The heat exchanger transfers heat from the hot engine oil to the cooler incoming fuel. This cools engine oil to improve lubrication and warms the fuel to prevent ice formation from water in the fuel system.

Revision 1.0

Fuel Filter and Bypass The engine fuel system includes a fuel filter and a bypass valve, which allows fuel to continue to the engine in the event of a clogged filter. Before the bypass valve opens, a pressure sensor sends a signal to the cockpit, alerting the pilot to an impending bypass situation. This may indicate fuel contamination.

Emergency Fuel Shutoff Valve The FMU incorporates an emergency fuel shutoff valve that is automatically actuated closed in the event of aft N1 shaft movement. This feature prevents N1 overspeed in the event of shaft separation by mechanically closing the emergency fuel shutoff valve.

Permanent Magnet Alternator The FMU also has a dual-coil permanent magnet alternator (PMA) that is integral to the FMU and is driven by the fuel pump drive shaft. It has a single rotor, with dual coils for dual output of electrical power. Under normal operating conditions, power is provided either from DC power by the aircraft electrical system or the PMA (whichever source is providing the greatest voltage). If normal DC power is not available, the PMA provides AC electrical power to the FADEC during all phases of operation.

OIL SYSTEM The oil system provides cooling and lubrication of the engine bearings and the accessory section (Figure 7-9).

Approved Oils Check the current list of engine oils in the “Limitations” section of the Airplane Flight Manual (AFM). Mixing approved oils is permissible if they are from the same brand but is not recommended except in emergency situations. Refer to the AFM for specific procedures.

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7 POWERPLANT

Engine start is initiated when the pilot presses and releases the respective engine start button, then immediately thereafter advances the same side throttle into idle. Once the correct N2 is reached, the FADEC commands the light-off fuel flow and both ignitors on. When an adequate rise in ITT is detected, the ignitors are automatically powered off by the FADEC.

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TO FUEL TANK

FROM FUEL TANK

LEGEND

FUEL BYPASS

FUEL FILTER

FUEL/OIL HEAT EXCHANGER

FADEC

TO SECONDARY MANIFOLD AND NOZZLES

FUEL FLOW TRANSMITTER

TORQUE MOTOR

A B

INTEGRATED FUEL METERING AND FUEL PUMP UNIT

HIGH-PRESSURE ENGINE PUMP

LOW-PRESSURE ENGINE PUMP

LP

LOW-PRESSURE SWITCH

Figure 7-8.   Engine Fuel System

TO PRIMARY MANIFOLD AND NOZZLES

EMERGENCY SHUTOFF VALVE (ESOV)

FMU

OIL IN

OIL OUT

TWO-STAGE FUEL PUMP

FIREWALL SHUTOFF VALVE

L FUEL TANK (FROM) (TO)

7 POWERPLANT R FUEL TANK (TO) (FROM)

THROTTLE

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FUEL

EXHAUST AIR

SCAVENGE OIL

ENGINE-DRIVEN PUMP PRESSURE

LEGEND MAIN OIL PRESSURE

RESTRICTOR

IMPENDING BYPASS POP-UP INDICATOR

PAV/CSV BYPASS

MAIN OIL PUMP

OIL TANK STRAINER

STRAINER

7 POWERPLANT

NO. 5 SCAVENGE PUMP

NO. 4 BRG

AIR/OIL SEPARATOR

Figure 7-9.   Oil System Schematic

FUEL/OIL HEAT EXCHANGER

MAIN OIL FILTER

CHIP COLLECTOR

MAIN OIL TEMPERATURE

STRAINER

NO. 3 BRG

ACCESSORY GEAR BOX

NO. 2 BRG

ACCESSORY GEAR BOX SCAVENGE PUMP

NO. 1 BRG

STRAINER

#5 BRG

TO ENGINE EXHAUST

STRAINER

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NOTE Maximum oil consumption is 1 U.S. quart per 13.5-hour period.

Components Oil Tank 7 POWERPLANT

The oil reservoir is an integral part of the AGB. Total capacity is 5.12 quarts (4.85 liters) (Figure 7-10).

Figure 7-11.   Oil Filler Port and Sight Glass Gauge

Oil Cooling The oil cooler is an oil-to-fuel heat exchanger. It uses output fuel from the low-pressure side of the engine-driven fuel pump to cool engine oil. Fuel is heated in the process so that ice does not form in the fuel (see Figure 7-9). AGB WITH OIL TANK

OIL FILTER BYPASS INDICATOR

Figure 7-10.   Oil System

The engines include a sight glass with MAX and MIN marks, and a sight-glass access door to make it more convenient to check the sight gauge oil level (see Figure 7-6). It has a filler port for servicing (Figure 7-11). The oil volume between MAX and MIN is approximately 0.4 quarts. Do not fill above the MAX mark. After servicing the engine, ensure the engine oil cap is correctly installed and the doors secured. The engine is equipped with a check valve feature to ensure that oil loss is prevented if the cap is not installed or is improperly installed.

Oil Filter The oil filter is a disposable cartridge that removes solid contaminants. It has bypass capability; however, there is no cockpit indication that the oil filter is bypassed. If the filter is approaching bypass, a poppet valve opens, pushing a mechanical indicator out from the valve to indicate that the filter is approaching bypass (Figure 7-12). This oil filter bypass indicator is checked during preflight and postflight inspection.

Oil Pump An engine-driven oil pump on the forward side of the AGB pressures oil throughout the engine to provide for lubrication and cooling. Strategically located engine-driven scavenge pumps collect oil from the extremities and serves to return oil to the tank. Figure 7-12.   Oil Filter Bypass Indicator

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On each engine, a 28-VDC starter-generator is attached to the AGB. To start the engines, the starter uses power from a ground power unit, the aircraft battery, or the opposite side generator. After the ENGINE START switches are pressed and as the engine exceeds approximately 40% N2, the starter-generator transitions to operate as a generator. For details on starter-generators, refer to Chapter 2—“Electrical System.”

CONTROLS AND INDICATIONS FADEC Each engine is controlled and monitored by its own dual-channel FADEC (Figure 7-13). The FADECs are in the tail cone on the aft pressure bulkhead, outside the engine-rotor noncontainment zone. The FADECs are the interface between the engines and pilot throttle control. Additionally, the FADECs are the main source of engine data for cockpit indications. Each FADEC receives signals directly from the engine and communicates through two channels. The Garmin interface adapter (GIA) 1 receives the output from FADEC channel A and GIA 2 from FADEC channel B. The FADEC controls the engine power settings using inputs from the engine sensors, aircraft sensors, and pilot-selected throttle position. The FADECs analyze pilot demands, environmental conditions, and engine operating limits. GIA 1

GIA 2

It then uses these parameters to schedule fuel flow to the engines (through the FMU) as necessary to provide the thrust level selected by the pilot with the throttles (see Figure 7-8).The FADECs monitor rotor speed and ITT, and can schedule fuel to prevent engine damage. During any engine ground start or air restart, the pilot is responsible for monitoring ITT. Depending on pilot settings of ignition switches and engine synchronization, FADEC may also control ignition and engine synchronization. The FADEC channel in control is alternated during each successive engine start. As the engine reaches idle speed on every ground start, the channel in control is switched in order to ensure both FADEC channels are capable of engine control. This also allows the FADEC to check for faults that can only be detected when the FADEC channel is in control. The FADEC does not switch the channel in control during in-flight start attempts. Normal DC power is provided to each FADEC and is available for engine starting and all engine operation. After engine start, if normal DC power becomes unavailable, the engine-driven PMA provides AC electrical power to the FADEC. The aircraft electrical system does not supply electrical power to the FADEC when using the emergency bus. Air data computer (ADC) data is provided to the FADEC in order to allow the FADEC to determine when a stabilized flight condition is established so a signal can be set telling the avionics package to record engine trend monitoring data. The only direct communication between the left and right FADEC is for engine synchronization and fault detection.

LH FADEC

ENGINE 1

B

A

429

429

A

429

429

FADEC RESET SWITCH RH FADEC

B

ENGINE 2

Figure 7-13.   FADEC/Avionics Interface

Revision 1.0

The FADEC RESET switch is on the bottom of the pilot tilt panel below the control yoke. It allows FADEC faults to be reset. After FADEC reset, if the fault is still present, the ENG CTRL SYS CAS message remains displayed. To reset the left engine FADEC, push the switch momentarily to the RESET L position. Push the opposite direction to reset the right engine FADEC (Figure 7-14).

FOR TRAINING PURPOSES ONLY

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Starter-Generators

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7 POWERPLANT

Figure 7-14.   FADEC Switch

THROTTLES One throttle for each engine is in the cockpit. The two throttles are on the throttle quadrant, and are labeled “L THROTTLE” and “R THROTTLE” outboard of their respective tracks (Figure 7-15). Each throttle controls a dual-coil position sensor, which sends pilot commands to the FADEC. Each throttle has detents at five thrust-level positions: • TO (takeoff power)—Commands maximum takeoff power and is intended for brief use at takeoff only. • CLB (climb power)—Commands maximum climb power and is mainly intended for use during climb to cruising altitude. • CRU (maximum cruise power)—Commands maximum cruise power and is mainly intended for use during normal cruise. • IDLE (normal engine idle)—Commands minimum safe continuous power and is mostly used for descent, landing, and stationary ground operations. Varies depending on aircraft on the ground,in flight, or with engine anti-ice turned on. • CUT OFF (engine cutoff)—Commands engine shutdown (fuel cut-off and ignitors off). Refer to the AFM for specific, current guidance on the use of these settings. The pilot can position the throttles at any detent, or at any position between the IDLE and CRU detents. When the throttle is not in a detent, FADEC estimates the intended thrust level based on throttle position and adjusts the engine accordingly. 7-12

Figure 7-15.   Throttle Quadrant

A barrier (gate) between the IDLE and CUT OFF detents prevents accidental engine cutoff and protects against accidental throttle advance out of CUT OFF. To move a throttle above or below the gate, use one finger to pull up the spring-loaded slide latch (triggers) under the throttle handle and hold the slide latch up while using the rest of the hand to move the throttle over the gate. When the throttle is over the gate, release the slide latch, and verify the throttle is full aft (on the IDLE detent). On the outboard side of each throttle handle, a large slide switch controls the speedbrakes, and a small GO AROUND pushbutton switch disconnects the autopilot and sets the flight director for a go-around. Refer to Chapter 15—“Flight Controls” for details on speedbrakes, and refer to Chapter 16—“Avionics” for details on the GO AROUND switch.

L AND R IGNITION SWITCHES The L and R IGNITION switches are on the lower instrument panel, left of the pilot control wheel. Each switch has two positions: ON and NORM (Figure 7-16). In the NORM position, ignition is controlled by the respective engine FADEC, which automatically energizes ignitors as necessary. FADECs energize ignitors during engine start, or if the FADEC detects flameout and activates autorelight. In the ON position, the ignitors operate continuously.

FOR TRAINING PURPOSES ONLY

Figure 7-16.   ENGINE START and IGNITION Switches

ENGINE START SWITCHES ENGINE START switches are grouped on the tilt panel, left of the pilot control wheel. Each is a lighted pushbutton switch. These include the L and R ENGINE START switches and the DISENG switch. Each pushbutton switch is a momentarycontact switch (Figure 7-16): • L and R ENGINE START switches—When the engines are not running, the L and R ENGINE START switches control the corresponding engine starters. Pressing either switch energizes the corresponding engine starter. Refer to Chapter 2—“Electrical Systems” for details on these switches and engine-start operations.

The switch enables or disables the engine synchronization capability of the FADECs. It has two positions: NORM and OFF. In the OFF position, engine synchronization is disabled. In the NORM position, FADECs automatically control engine synchronization in flight when all of the following conditions are true: • ENGINE SYNC switch is in the NORM position • Landing gear are retracted • Each throttle is out of the TO detent and above the IDLE detent • Throttle levers are within 5° of each other • N1 references are within 5% of each other

ENGINE INDICATING AND CREW ALERTING SYSTEM The engine indication and crew alerting system (EICAS) contains all indications for the powerplant and its systems. These include continuous engine indications and crew alerts as necessary. In normal EICAS display mode, these indications are in two columns on the left side of the G1000 multifunction display (MFD) (Figure 7-18).

• D ISENG switch—The DISENG switch (starter-disengage switch) opens the start relay. This may be required if a starter continues to operate too long, or when the engine has reached too high a speed without the starter automatically disengaging.

ENGINE SYNC SWITCH The ENGINE SYNC switch is on the throttle quadrant, to the right of the CUT OFF position of the right engine throttle (Figure 7-17). Figure 7-18.   EICAS Display on MFD

Figure 7-17.   ENGINE SYNC Switch

Reversionary mode is selected by pressing the red DISPLAY BACKUP button at the bottom of either audio control panel. In this mode, most of these indications are presented in a single-column EICAS display. The reversionary mode EICAS display normally is on the left side of the pilot and/or copilot PFD. In reversionary mode, some

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EICAS displays are simplified or eliminated, and crew alerts appear in a box on the right side of the affected display (Figure 7-19). 109.10 ITWI vtf 013° 113.80 ICT LOC HDG

NAV1 113.80 NAV2 116.80 N1%

CRU

CHITO faf AP YD VS

7.1NM BRG 034° 900FPM ASEL

CRU

89.5

3000

100 90

140

80

E

20

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10

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800 700

400 950

FLAPS

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UP

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013°

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BARO MIN 2900FT

E

TERM

BRG1

CDI

-900 2 4

29.92IN

ADF/DME

POINTER (SAFE REGION)

MAJOR GRADUATION SLIDER (SAFE REGION)

OIL PRESS LO L CABIN ALT T2 HTR FAIL R W/S O’HEAT L W/S A/I FAIL L FUEL PRES LO R ENG A/I COLD L CABIN DOOR AFT DOOR P/S HTR L F/W SHUTOFF R SURFACE DE-ICE SPD BRK EXTEND FUEL TRANSFER

POINTER (IN EXCEEDANCE) NORMAL REGION

SLIDER (IN EXCEEDANCE)

EXCEEDANCE REGION

FLIGHT PLAN

Approach –

CHITO faf

16.5NM ICT NAV2 BRG2

EXCEEDANCE REGION

CAS

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S

CABIN PRESS ALT FT 1400 0.8 DIFF PSI

DME NAV1 109.10 –.–– NM 6.8NM CHITO GPS WIND DME

6

3

CRS

15

25 DC VOLTS 20 105 DC AMPS 205 BATT VOLTS 27

M .399 5 +6°C 3 +15°C

RAT ISA

043°

BOTH ON GPS2

2

2900

054°

N

134 66

90

R RF 1 HDG

33

300 1040

70.1

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3000

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N2% OIL PSI °C FUEL LBS PPH 13 °C

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835 68.3

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HDG MISCOMP IAS NO COMP ALT NO COMP

3300 1 20 33 100 00

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118.200 COM1 125.150 COM2

2900 -900

3200

11 8 7

900

24

ITT°C

46.2

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20

126.700 121.900

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DTK

ILS 01L 013° 013° 013°

RW01L map 3000FT ICT mahp XPDR1

1200 ALT

R LCL

IDENT STD BARO BACK

DIS

6.6NM 4.8NM 3.3NM 18:11:37 MSG

Figure 7-19.   EICAS Reversionary Display on PFD

MINOR GRADUATION

Figure 7-20.   EICAS Analog Markings

If the EICAS does not receive valid data for an indication, it replaces the indication with a red “X” (Figure 7-21).

The FADECs pass information to and from the GIAs. The GIAs then send updated engine performance and fault information to the EICAS display. The FADEC provides engine data to the EICAS for: • N1 % rpm • N2 % rpm • ITT Colors of scales, pointers, and digits indicate the current condition of the affected system (Figure 7-20): • Red indicates a warning that a limitation has been exceeded. • Amber indicates a caution that a system is near its limitation, and operating in a timelimited region. • Green indicates normal operation. • White indicates vertical analog “tapes” when in normal operating range, and also for labels on indications. • Cyan indicates pilot-defined settings, or recommended target values as determined by FADEC. On some EICAS indications (ITT, oil pressure, oil temperature), digits only appear when relevant or when abnormal or emergency conditions exist.

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Figure 7-21.   EICAS Display - Invalid Data

NOTE A red “X” on an EICAS indication does not mean that the indicated value is zero or is exceeding normal levels. It only indicates that the EICAS cannot determine the correct value to display, and that the EICAS indication is inoperative. Red lines on some scales, indicating maximum allowable limits, may not appear. This does not mean there is no limit for that item. It means the EICAS cannot determine what the appropriate red line value is.

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EICAS powerplant indications include: • N1 % rpm and thrust mode* • SYNC indications • ITT and ignition indications* • N2 % rpm* • Oil pressure (psi) and temperature (°C) • Crew alerting system (CAS) messages * If normal DC power fails, these items are powered from the permanent magnet alternator (PMA) and remain visible on the EICAS.

N1% WINDOW For each engine, the N1% window of the EICAS (Figure 7-22) indicates: • N1% rpm • N1% target bug • Thrust mode • Engine SYNC

THRUST MODE INDICATOR

N1% RPM The N1% scale indicates the rotation speed of the N1 spool and is calibrated in percent of maximum N1 rpm (as determined by FADEC). It is the primary indication of engine thrust. When in acceptable range, the analog tapes are white and the digits are green. When outside acceptable range, both tapes and digits are red. When N1 is below 20% of maximum, the digits are displayed and the tape display does not indicate below 20% N1. A red line indicating maximum rpm limit (as determined by FADEC) is at 100% rpm on each scale.

N1% Target Bug N1% target rpm, as calculated by FADEC for the selected thrust mode, is indicated by cyan digits in a box centered at the top of the N1 scales, and by a cyan marker (“bug”) on the outboard side of each N1 scale (Table 7-1). Table 7-1.   TARGET N1 BUG GEAR STATUS

THROTTLE POSITION

BUG

DOWN

ANY

TO PWR

UP

CRU DETENT

CRU PWR

UP

CLB DETENT

CLB PWR

UP

TO DETENT

TO PWR

UP

BETWEEN DETENTS

NEXT HIGHER DETENT PWR

N1% TARGET (DIGITAL)

Thrust Mode

N1% RED LINE

Thrust-mode indications appear in cyan at the top of each N1 scale. These indications correspond to the throttle settings currently selected by the pilot. If the pilot selects a setting between detents, the thrust mode indications do not appear.

N1% TARGET (ANALOG) N1% (ANALOG) N1% SYNC INDICATOR N1% (DIGITAL)

Figure 7-22.   N1% Window

The N1% window is powered by normal or emergency DC power. It is always visible and operating when any DC power is active in the aircraft. All N1% window indications remain valid when the aircraft is on emergency DC power.

In normal display mode, these indications are in the upper-left corner of the MFD. In reversionary mode, they are presented at the top of the reversionary EICAS display on the pilot and/or copilot PFD.

Engine SYNC At the bottom-center of the N1% display, the label “SYNC” appears in green letters when the engines are synchronized in flight by the FADEC, when ENGINE SYNC NORM is selected by the pilot.

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In addition to powerplant status indications, the EICAS provides information on most other aircraft systems. For details on those indications, refer to Chapter 4—“Master Warning System.”

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N2% Window

The ITT and ignition window provides current status of ITT and engine ignition systems (Figure 7-23). It is powered by normal or emergency DC power, and is always visible and operating when any DC power is active in the aircraft. All ITT and ignition window indications remain valid when the aircraft is on emergency DC power.

The N2% window indicates the rotation speed of the N2 spool in percent of maximum N2 rpm (as determined by the FADEC) (Figure 7-24). It is a key indication of engine condition. The window appears immediately below the ITT and ignition window, whether in the EICAS normal display or reversionary display. The digit colors are: • White during engine start • Green in acceptable range • Red when outside acceptable range

7 POWERPLANT

ITT AND IGNITION WINDOW

ITT RED LINE ITT AMBER BAND

The N2% window is powered by normal or emergency DC power. It is always visible and operating when any DC power is active in the aircraft. All N2% window indications remain valid when the aircraft is using emergency DC power.

ITT (ANALOG) ITT (DIGITAL)

Figure 7-23.   ITT and Ignition Window Figure 7-24.   N2% Window

ITT Display The ITT°C window appears on the top of the right column of the normal EICAS display, or below the N1% window on the EICAS reversionary display. For each engine, it indicates current ITT, which provides an indication of interstage turbine temperature. Maximum allowable (red line) ITT and abnormal high ITT ranges are calculated by FADEC and displayed as a red line and a short amber band, respectively, on each scale. The ITT scale is calibrated in degrees Celsius (°C). A white tape and pointer moves along the outboard side of each scale. Digits appear at the bottom of the scale to indicate current ITT during engine start or if ITT values are under 200.

Ignition (IGN) Display The green IGN indications appear at the top left or right of the ITT scale and correspond to the respective engine. The green IGN only indicates that the exciter box is powered.

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OIL PRESSURE (PSI) AND TEMPERATURE (°C) INDICATIONS The OIL window appears immediately below the N1% window on the normal EICAS display (Figure 7-25), or immediately below the N2% window on the reversionary display (Figure 7-26). The OIL window is powered by normal DC power. However, when the aircraft is using emergency DC power, only the oil temperature (°C) indication for the left engine remains valid.

Oil Pressure (PSI) In the normal EICAS display, oil pressure (in psi) for each engine is displayed by pointers on the corresponding sides of twin vertical analog scales (and by pointers and digits when at the end of the scale). The scales have color bands indicating normal (green), abnormal (amber), and unsafe (red) ranges. The analog scale bands for the left and

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Figure 7-25.   Oil Pressure and Temperature Window (Normal)

For specific, current instructions on normal operating procedures, refer to the AFM. Where the following information differs from the AFM, use the AFM information and follow the AFM instructions. The following information is only for training and background information, and may change without notice. These procedures focus only on powerplant items in these stages of aircraft operations. Other systems are also involved, and steps are required for them, but are not noted here. Refer to the AFM or checklist for details.

PREFLIGHT In addition to the other systems that must be checked during preflight, the powerplants require particular attention. Before preflight, ensure that all four engine covers are removed and stowed, and that both throttles are selected to CUT OFF. Figure 7-26.   Oil Pressure and Temperature Window (Reversionary)

right engine are separate because the low oil pressure caution region and the red line limit change as a function of N2 speed. The pointers are the same color as the band to which they are pointing. The digits display at the bottom of the oil pressure scales in the same color as the pointer, but only when the pointer is outside the green range. In EICAS reversionary mode, only the color-coded digits appear, but they appear at all pressures.

Oil Temperature (°C) In normal EICAS display, oil temperature (degrees Celsius) for each engine is displayed by pointers on the corresponding sides of a single vertical analog scale. The scale has color bands indicating normal (green), abnormal (amber), and unsafe (red) ranges. The pointers are the same color as the band to which they are pointing. Digits display at the bottom of the oil temperature scale in the same color as the pointer, but only when the pointer is outside the green range. In EICAS reversionary mode, only the color-coded digits appear, but they appear at all temperatures.

Revision 1.0

Inspect the engine intakes and exhaust (including the fan bypass duct) for any indication of ice or foreign object contamination. Ensure the following are all clear: • Engine air inlet • Generator cooling air inlet • Engine anti-ice exhaust • Pylon precooler inlet • Generator cooling air exhaust • Engine fluid drains • Pylon precooler exhaust Check the engine rotors (and the engine T2 probe in the engine inlet) for bent blades, nicks, and blockage of fan stators (stationary blades). Check the oil filter bypass indicator by viewing the indicator button through the access panel on the lower right side of each engine nacelle (see Figure 7-7). If the button is extended (popped), maintenance is required before flight. Open the oil door to check the oil level in the oil level sight glass (see Figure 7-11). It should be between MIN and MAX. The normal time to check engine oil is 10 minutes after engine shutdown.

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When finished checking the oil system, ensure that the access doors are secure.

START

7 POWERPLANT

Before starting the engines, complete preflight inspection and ensure that no inappropriate CAS messages appear. Ensure that no blowing debris is likely to be ingested by the engine. Verify that no aircraft are closer than 50 feet immediately behind the engine and that no people are within the hazard area in front or behind the engine (see Figure 1-4 in Chapter 1—“Aircraft General”).

NOTE Consider wind velocity prior to attempting engine start in order to preclude exceeding wind-related limitations. Reposition the aircraft if required (see section II of the FAA-approved AFM). Prior to starting the first engine, review the STARTING ENGINES checklist to prepare for steps that will take place during the sequence. Verify that adequate voltage is available for the start and then press the START button, verifying that the appropriate START group lights illuminate. Lift the throttle trigger, then place the lever to the idle position. FADEC introduces fuel and energizes the ignition, which should result in combustion. Scan to check that all components of the start occur and monitor ITT as it begins to rise. Ensure that the starting limitations are not exceeded as the engine accelerates, and always stand ready to terminate the start, if required, by guarding the throttle. Check that the ITT rises immediately. If ITT rapidly approaches 830°C or shows no rise within 10 seconds, abort the start. Do not exceed 830°C for more than 5 seconds; and never exceed the maximum limit of 862°C. If engine maintenance has been performed, air in fuel lines may cause a hot start. Accomplish proper purging procedures prior to attempting a start. Be prepared to abort the start. With the throttle at idle, on the ground, FADEC automatically varies fuel flow as required to maintain N2 at 48.6%. Note that the N2 display digits

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change from white to green, indicating that the FADEC start sequence is complete. Also, verify that all EICAS indications are normal and proceed to start the second engine.

GROUND OPERATION When operating on the ground, maintain throttles at IDLE, except as necessary for engine and system checks or for taxiing.

CAUTION When operating on the ground, be aware of the hazardous effect of jet exhaust blast on people and other aircraft in the area. Avoid ground maneuvers and/or power settings that may result in damage or injury to others in the area (see Figure 1-4 in Chapter 1—“Aircraft General”). When beginning to taxi, verify both brakes are operating and nosewheel steering is effective.

FLIGHT OPERATIONS Takeoff At takeoff, while holding brakes, select throttles to the TO detent. Verify the FADEC thrust mode EICAS indicator (top of the N1% window) displays a cyan TO for each engine. Verify all EICAS indications are normal and N1% rpm is at the cyan command bug for each engine. Release the brakes and maintain full takeoff power until reaching safe altitude.

NOTE Takeoff thrust is limited to 5 minutes except during emergency situations (i.e., one engine inoperative). Refer to the AFM.

After Takeoff—Climb During climb, select throttles to the CLB detent. Verify that the FADEC thrust mode indicator displays a cyan CLB for each engine, which indicates the FADECs are automatically setting maximum climb thrust on each engine.

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During cruise, select throttles to CRU detent, or as desired. If using the CRU detent, FADEC automatically sets maximum cruise thrust; verify that the FADEC thrust mode indicator displays a cyan CRU for each engine.

NOTE The throttles should be reduced to the CRU detent or below within 10 minutes after reaching an intermediate or final cruise altitude. The use of CLB during normal operations beyond 10 minutes after reaching cruise altitude will significantly decrease engine life and increase operator costs.

EMERGENCY/ ABNORMAL For specific information on emergency/abnormal procedures, refer to the appropriate checklist or FAA-approved AFM. Table 7-2.   CAS MESSAGES ENGINE FAIL L-R DESCRIPTION INHIBITS

INHIBITS

During descent, approach, and landing, reduce the throttles as necessary to manage descent/approach profile and navigate as required.

Shutdown Prior to shutdown, allow the ITT to stabilize at a minimum value for 2 minutes. When ready to shut down, lift the side latch (triggers) and pull each throttle into CUT OFF, individually. Monitor the EICAS panel during shutdown to verify that operation of each engine has terminated and that the ITT has decreased accordingly. Check the oil level 10 minutes after shutdown. Ensure that the cowl door is secured. When the engine, inlet, and exhaust nozzle are cool, install the four engine covers.

An engine has failed NONE

OIL PRESS LO L-R DESCRIPTION

Descent, Approach, And Landing

7 POWERPLANT

Cruise

This message occurs whenever oil pressure digits turn red based on a low-level exceedance. EMER, ESDI, ESI

ENG CTRL SYS L-R DESCRIPTION

INHIBITS

FADEC has a fault that requires maintenance prior to the next dispatch of the aircraft. NONE

F/W SHUTOFF L-R DESCRIPTION INHIBITS

Fuel firewall shutoff valve is fully closed. EMER

FUEL FLTR BP L-R DESCRIPTION INHIBITS

Engine fuel filter impending bypass switch is closed. EMER

FUEL PRES LO L-R

LIMITATIONS

DESCRIPTION

For specific limitations, refer to the FAA-approved AFM.

Revision 1.1

INHIBITS

FOR TRAINING PURPOSES ONLY

Fuel pressure is under 4.65 psig (decreasing) or 6.4 psig (increasing) EMER

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QUESTIONS 1. The primary thrust indicator for the Pratt & Whitney PW615F-A is: A. Fuel flow B. N1 C. ITT D. N2 7 POWERPLANT

2. If one ignitor fails during engine start: A. The engine starts normally on remaining ignitor B. It results in a “hot” start C. Combustion does not occur D. The exciter box acts as a backup and the engine starts 3. Ignition during normal engine start is activated by FADEC when: A. Turning the IGNITION switches on at 8%–12% N1 B. Moving the throttle to IDLE and the correct N2 is reached C. Depressing the start button D. Nothing; ignition is not needed during normal engine start 4. Ignition and boost pump operation during engine start are normally terminated: A. By turning the IGNITION switches off B. Automatically at the termination of the start sequence C. By turning the boost pump switch off D. By opening the ignition circuit breakers on the right CB panel 5. The maximum tailwind component for engine start is: A. 10 knots B. 30 knots C. 40 knots D. 45 knots

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6. The OIL PRESS LO L–R CAS message on the MFD illuminates when: A. Oil pressure is below minimum acceptable limits and engine failure may occur B. Oil pressure is less than 25 psi C. Oil filter clogs and bypasses oil D. The fuel-oil cooler becomes clogged 7. If the N1 fan shaft shifts aft: A. The engine automatically shuts down B. The vibration detector causes illumination of the MASTER WARNING lights C. The synchronizer shuts the engine down D. Nothing occurs 8. The following engine instruments are available in the event of a total loss of DC electrical power: A. N1 rpm and ITT B. N1 rpm, N2 rpm, and ITT C. N1 rpm (standby digital LCD) D. None of the above 9. The minimum oil temperature for engine start is? A. 0°C B. –40°C C. –20°C D. –30°C 10. If the ITT is rapidly approaching 830°C, terminate the engine start by moving the: A. ENGINE SYNC switch to OFF B. IGNITION switch to ON C. Throttle to CUT OFF D. FADEC to RESET L–R 11. For any flight that exceeds 5 hours, the oil level must be at least: A. MAX B. MIN C. One quart low D. To the top of the oil cap

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12. The maximum crosswind component for engine start is: A. 10 knots B. 15 knots C. 20 knots D. 25 knots 13. If normal DC power is not available, the FADEC receives power from the: A. Aircraft battery B. Standby battery C. Generators D. Permanent magnet alternator

17. If the F/W SHUTOFF R CAS message posts when positioning the throttle to CUTOFF, this indicates: A. A malfunction of the normal shutdown system B. The fuel firewall shutoff valve is performing the auto self-test C. The FADEC does this on every shutdown D. The system bypass valve is performing the auto self-test

14. ITTs should be allowed to stabilize for how long before engine shutdown? A. 1 minute B. 2 minutes C. 3 minutes D. 4 minutes 15. It is recommended that the throttles should normally be reduced to the CRU or below position within how many minutes after reaching cruise altitude? A. 5 B. 10 C. 15 D. 20 16. When the FADEC start sequence is completed, the N2: A. Display digits change from WHITE to GREEN B. Slider scale changes from WHITE to CYAN C. Low rpm alarm silences D. Start light extinguishes

Revision 1.0

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INTENTIONALLY LEFT BLANK

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CHAPTER 8 FIRE PROTECTION CONTENTS INTRODUCTION.................................................................................................................. 8-1 GENERAL................................................................................................................................8-1 DESCRIPTION....................................................................................................................... 8-2 Engine Fire Detection System......................................................................................... 8-2 Engine Fire-Extinguishing System.................................................................................. 8-2 Portable Fire Extinguisher............................................................................................... 8-2

Fire Detection Loop......................................................................................................... 8-2 Pressure Sensor............................................................................................................... 8-2 Engine Fire Bottle Assembly........................................................................................... 8-2 Fuel Shutoff Valve and Generator Disconnect................................................................. 8-4 Portable Fire Extinguisher............................................................................................... 8-4 CONTROLS AND INDICATIONS........................................................................................ 8-4 Engine Fire Lights........................................................................................................... 8-4 MASTER WARNING Lights.......................................................................................... 8-4 White BOTTLE ARMED Lights.................................................................................... 8-4 OPERATION.......................................................................................................................... 8-5 Preflight........................................................................................................................... 8-5 In Flight........................................................................................................................... 8-5 LIMITATIONS........................................................................................................................ 8-6 EMERGENCY/ABNORMAL................................................................................................ 8-6 QUESTIONS.......................................................................................................................... 8-7

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8-i

8 FIRE PROTECTION

COMPONENTS..................................................................................................................... 8-2

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ILLUSTRATIONS Figure Title Page 8-1.

Engine Fire Detection and Extinguishing System................................................... 8-3

8-2.

Engine Fire Bottle.................................................................................................... 8-4

8-3.

Portable Fire Extinguisher....................................................................................... 8-4

8-4.

ENGINE FIRE and BOTTLE ARMED Lights....................................................... 8-5

8-5.

Rotary TEST Knob.................................................................................................. 8-5

TABLES 8-1.

CAS MESSAGES.................................................................................................... 8-6

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CHAPTER 8 FIRE PROTECTION

INTRODUCTION This chapter describes the fire protection system on the Citation Mustang aircraft. The engine firedetection system consists of two separate detection circuits (one for each engine), which provide visual warnings. The engine fire-extinguishing system includes one fire bottle, which is activated from the cockpit. A portable fire extinguisher is in the cabin.

GENERAL The fire-protection system consists of engine fire detection, engine fire extinguishing, a portable fire extinguisher, and aircraft construction that reduces a fire risk. The engine fire-detection system detects fires and overheat conditions in the engine nacelles and alerts the crew. The engine fire-extinguishing system suppresses those fires upon pilot command by supplying fire-extinguishing agent. Fire suppression in the cabin area is accomplished using a portable fire extinguisher.

For additional protection, the engine nacelle fire zone is separated from the pylon and the rest of the aircraft by a stainless steel firewall. At the firewall penetrations and in the nacelle, fuel is contained in stainless steel fittings, stainless steel tubes, and fire-resistant hoses. The rotary TEST knob on the instrument panel is used to test the fire warning system.

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DESCRIPTION

COMPONENTS

ENGINE FIRE DETECTION SYSTEM

FIRE DETECTION LOOP

The engine fire-detection system consists of: • Fire-detection loop • Pressure sensor • Red L and R ENGINE FIRE lights Excessive heat by fire or other heat sources expands an inert gas inside the fire-detection loop. The expansion of gas closes a pressure switch that sends a signal to illuminate the left or right ENGINE FIRE light.

8 FIRE PROTECTION

The fire-detection system requires DC power. FIRE DETECT circuit breakers for the independent sides are on the left and right CB panels within the ENGINE SYSTEMS grouping.

ENGINE FIRE-EXTINGUISHING SYSTEM The single bottle engine fire-extinguishing system enables the flightcrew to suppress a fire in the left or right engine compartment. This action is limited to one use. The engine fire-extinguishing system consists of: • Engine fire bottle assembly • Distribution tubes • Nozzles in each engine nacelle • BOTTLE ARMED lights • Fuel shutoff valve and generator disconnect These components shut off the generator and fuel supply as well as discharge extinguishing agent, which is pressurized with nitrogen and discharged by electrically activated cartridges to the engine nacelles.

PORTABLE FIRE EXTINGUISHER A portable fire extinguisher provides fire protection inside the aircraft.

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Each fire detection loop detects a fire or overheat condition in the respective engine compartment (Figure 8-1). The tube routes along both sides of the engine. An increase in temperature on any part of the tube increases the pressure of the gas (helium) inside the tube. If the helium filled tube develops a leak, the gas escapes decreasing the normal pressure in the tube. This closes a test pressure switch indicated by the failure of the affected side ENGINE FIRE light when checked by the rotary test.

PRESSURE SENSOR A pressure sensor is at the end of the fire-detection loop. When the loop is heated by fire or a bleedair leak, the gas in the tube expands, activating the pressure sensor. This produces an electrical signal that provides a warning to the flightcrew. The signal is in the form of: • L or R ENGINE FIRE lights • MASTER WARNING lights

ENGINE FIRE BOTTLE ASSEMBLY The engine fire bottle (with two squibs) is in the tail compartment. It can be used to extinguish a fire in either engine nacelle (Figure 8-2). The fire bottle contains enough extinguishing agent to protect against one engine fire (0.85 pound of Halon 1301 extinguishing agent). The fire-extinguisher bottle contains two individual firing cartridges (squibs). The cartridges are connected to distribution tubes that are routed to the left and right engine compartments. The bottle has a safety relief valve that thermally relieves (discharges) its contents into the tail cone if the internal bottle temperature rises above 210°F.

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CONTROL BOXES

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SQUIB GAUGE

FIRE DETECTION LOOP (HELIUM-FILLED TUBE)

Figure 8-1.   Engine Fire Detection and Extinguishing System

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CONTROLS AND INDICATIONS ENGINE FIRE LIGHTS Red L and R ENGINE FIRE lights are on the upper part of the center instrument panel (Figure 8-4). They respond to signals from the respective engine fire sensors. Each light is covered by a spring-loaded, transparent plastic guard and has an integral pushbutton switch. Figure 8-2.   Engine Fire Bottle

FUEL SHUTOFF VALVE AND GENERATOR DISCONNECT

8 FIRE PROTECTION

The firewall shutoff valve closes and electrical flow from the generator is stopped when an illuminated ENGINE FIRE light is pushed. The F/W SHUTOFF L-R CAS message in the crew alerting system (CAS) window indicates that the fuel shutoff valve is fully closed. When the generator is disconnected, the respective GEN OFF message appears in the CAS window.

PORTABLE FIRE EXTINGUISHER One portable handheld fire extinguisher is in a drawer in the cabinet behind the pilot (Figure 8-3). It is accessible from either the pilot, copilot, or passenger positions. The Halon 1301 type extinguishing agent discharges as a vapor with no residue or decrease in vision to personnel. The discharge distance is approximately 9–15 feet with a discharge time of 10 seconds.

If the red L and/or R ENGINE FIRE light illuminates steady, it indicates a fire or overheat condition in the corresponding engine.

MASTER WARNING LIGHTS The MASTER WARNING lights are on the instrument panel above each primary flight display (PFD). The MASTER WARNING lights illuminate flashing when the L or R ENGINE FIRE lights illuminate. The pilot acknowledges by pressing one of the MASTER WARNING lights. Pressing will extinguish both lights.

WHITE BOTTLE ARMED LIGHTS A white BOTTLE ARMED light is below each red ENGINE FIRE light on the upper-center panel (Figure 8-4). Each BOTTLE ARMED light has an integral pushbutton switch. These lights indicate when the bottle is armed for the respective engine and prepared to release extinguishing agent. After the extinguishing agent is released, the light extinguishes, indicating to the crew the extinguisher bottle is empty and is no longer available for use.

Figure 8-3.   Portable Fire Extinguisher

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Figure 8-4.   ENGINE FIRE and BOTTLE ARMED Lights

OPERATION PREFLIGHT Rotary TEST Knob Test the engine fire-detection system before each flight by using the rotary TEST knob (Figure 8-5) during the preflight inspection. This test verifies connections to the fire bottles and warning system. Illumination of both ENGINE FIRE lights is an indication the warning system is working properly.

NOTE A successful test of the fire-detection system using the rotary TEST knob, or illumination of either BOTTLE ARMED light, does not confirm that the fire bottle is serviced and full. This can only be confirmed by a visual check of the bottle gauge and comparing the reading to a placard that correlates the acceptable pressure/temperature ranges.

During preflight, check that the portable fire extinguisher is serviced and secure. Verify that the pressure gauge on the extinguisher indicates in the green arc and that the extinguisher is secure in its drawer behind the pilot seat.

Engine Fire Bottle Inspection An inspection door is in the aft compartment to view the fire bottle gauge. A placard is on the back of the door. Check that the gauge pressure matches the acceptable ranges based on outside air temperature (OAT). Refer to the Normal Procedures Checklist.

IN FLIGHT Refer to approved Airplane Flight Manual (AFM) checklist. Pushing the L ENGINE FIRE or R ENGINE FIRE light: • P rovides power to the fuel shutoff valve, which cuts off the fuel supply to the affected engine. The F/W SHUTOFF L-R CAS message appears in the CAS window. • D isconnects the starter-generator on the affected engine. The GEN OFF L-R CAS message appears in the CAS window.

Figure 8-5.   Rotary TEST Knob

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• A rms the engine fire bottle squib (explosive cartridge) that routes extinguishing agent to the selected engine. However, the bottle contents do not yet discharge into the engine. (The corresponding white BOTTLE ARMED light illuminates.)

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Portable Fire Extinguisher

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Once the ENGINE FIRE light is pressed the corresponding BOTTLE ARMED light illuminates indicating that the fire bottle is armed and ready. When the BOTTLE ARMED light is pressed the BOTTLE ARMED light extinguishes and it fires the squib releasing the extinguishing agent into the engine cowling.

CAUTION

LIMITATIONS For specific limitations, refer to the FAA-approved AFM.

EMERGENCY/ ABNORMAL

The white BOTTLE ARMED light does not illuminate (and cannot operate) until after the corresponding red L or R ENGINE FIRE light has been pressed.

For specific information on emergency/abnormal procedures, refer to the appropriate checklist or FAA-approved AFM.

NOTE

Table 8-1.   CAS MESSAGES

8 FIRE PROTECTION

If a crewmember presses the other BOTTLE ARMED light, which is not illuminated, the fire bottle does not discharge, and no extinguishing occurs. When the pilot pushes the illuminated white BOTTLE ARMED light, the light extinguishes.

Portable Fire Extinguisher

GEN OFF L-R DESCRIPTION

INHIBITS

If smoke or fire is present, immediately don oxygen masks and smoke goggles, and set oxygen to EMERGENCY. Ensure that passengers have supplemental oxygen.

LOPI, TOPI

F/W SHUTOFF L-R DESCRIPTION

CAUTION

When either of the ENGINE FIRE lights are pushed, the respective amber GEN OFF message appears in the CAS window.

INHIBITS

When either of the ENGINE FIRE lights are pushed, the respective amber F/W SHUTOFF message displays in the CAS window. This indicates that the corresponding fuel shutoff valve is closed. LOPI, TOPI

To operate the portable fire extinguisher, open the top cabinet drawer and remove extinguisher, hold the extinguisher upright, and aim the extinguisher at the base of fire. Using the attached ring, pull the pin from the extinguisher. Squeeze the handles of the extinguisher together to release the extinguishing agent. Spray the extinguishing agent using a side-to-side motion while aiming at the base of the fire. Anytime the extinguisher is used, even partially, maintenance is required before further dispatch.

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QUESTIONS 1. If the L ENGINE FIRE light extinguishes after moving the left throttle to idle, this indicates a possible: A. Bleed-air leak B. Hydraulic leak C. Fuel leak D. Nitrogen leak

6. If the right side squib fails, fire protection is: A. Still available for both engines from the left side B. Not available to the right engine C. Available through the rotary TEST knob D. Not available to either engine

8 FIRE PROTECTION

2. What indications are visible when the rotary TEST knob is placed in the FIRE WARN position? A. Red L ENGINE FIRE, R ENGINE FIRE, and MASTER WARNING lights illuminate B. BOTTLE ARMED PUSH lights illuminate C. Fire bottle pressure is displayed on the multifunction display D. Confirms the fire bottle is serviced and full 3. During the preflight inspection, sufficient engine fire bottle pressure can be determined by: A. Successful completion of the engine fire position on the rotary test. B. Visual inspection of the bottle pressure gauge against the bottle pressure placard. C. The absence of the ENG FIR BOTTL SERV CAS message. D. Checking to insure that the engine fire bottle gauge is within the green arc. 4. The fire bottle provides how many chances to extinguish a fire? A. 1 B. 2 C. 3 D. 4 5. Pressing the R ENGINE FIRE light: A. Arms the fire bottle B. Disables the right generator C. Stops right fuel flow past the firewall D. All of the above

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CHAPTER 9 PNEUMATICS CONTENTS INTRODUCTION.................................................................................................................. 9-1 GENERAL................................................................................................................................9-2 DESCRIPTION....................................................................................................................... 9-2 Bleed-Air Distribution..................................................................................................... 9-2 Compressed Nitrogen Bottles.......................................................................................... 9-4 LIMITATIONS........................................................................................................................ 9-4 EMERGENCY/ABNORMAL................................................................................................ 9-4 QUESTIONS.......................................................................................................................... 9-5

ILLUSTRATIONS Figure Title Page Mustang Pneumatic Systems................................................................................... 9-3

9 PNEUMATICS

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INTRODUCTION This chapter describes the pneumatic systems on the Citation Mustang aircraft. The pneumatic systems route air or nitrogen from various sources to aircraft systems that use pneumatics for heating, cooling, pressurization, landing gear, and brakes. Because each of the Mustang pneumatic systems is dedicated to a specific purpose, this chapter provides a brief overview of each system, then refers the reader to the appropriate chapter elsewhere in this manual.

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GENERAL The Mustang pneumatic systems are each discrete systems, dedicated to a specific task, and isolated from all other pneumatic systems. The Mustang pneumatic systems (Figure 9-1) include: • Bleed air from engine compressors (outboard bleed-air port on each engine) for pneumatic ice-protection systems • Bleed air from engine compressors (inboard bleed-air port on each engine) for temperature-controlled pressure vessel air supply • Compressed nitrogen from a storage bottle for emergency landing gear extension (blowdown bottle) • Compressed nitrogen from a storage bottle for emergency brakes (emergency braking bottle) Each of these systems is independent of the others and can function when any other pneumatic system fails. Single-engine operation can normally maintain all required pneumatic system functions. However, loss of DC power can cause complete or partial failure of multiple systems. Compressed nitrogen pneumatic systems are not dependent upon engine operation or DC power.

9 PNEUMATICS

Safety devices in each pneumatic system prevent excessive pressure. Each system has its own controls. All systems are controlled directly or indirectly by pilot command. Indications for the compressed nitrogen systems are in the nose baggage compartment. Indications for all other pneumatic systems are displayed in the engine indicating and crew alerting system (EICAS) in the cockpit displays.

DESCRIPTION BLEED-AIR DISTRIBUTION High-temperature engine bleed air is extracted from the high-pressure compressor section of each engine and routed through two separate ports. The outboard port on each engine supplies bleed air for ice protection. The inboard port supplies bleed air for the temperature-controlled pressure vessel air supply. 9-2

From separate engine bleed-air ports, the bleed air enters ducts to the ice protection system and pressure vessel air supply. Check valves prevent flow (in any of the ducts) from reversing and entering an engine, including any crossover flow from the opposite engine.

Ice Protection An outboard bleed air port from each engine supplies bleed air for ice protection. It supplies hot engine bleed air: • D irectly to the respective engine anti-ice system • T hrough a service air regulator to the aircraft pneumatic deice boot system

Engine Anti-Ice System Bleed air for each engine anti-ice system is routed to the leading edge of the engine inlet through a valve. The hot bleed air warms the leading edge as it passes through, then exits overboard through an opening in the bottom of the engine nacelle. The engine anti-ice system is explained in Chapter 10—“Ice and Rain Protection.”

Surface Deice (And Service Air) System Bleed air for the surface deice system is routed to the service air regulator for operation of pneumatic deice boots. Refer to Chapter 10—“Ice and Rain Protection.”

Pressure Vessel Air Supply An inboard bleed air port from each engine supplies air for temperature control and pressurization. The bleed air is routed through a heat exchanger in the respective engine pylon. The heat exchanger dissipates heat from the bleed air to the metal ducts of the heat exchanger. Cooler outside ram air from the pylon ram-air inlets passes over the heat exchanger ducts and carries the heat away. The temperature of the pressure vessel air supply is regulated by a temperature control valve, which varies the amount of pylon ram air flowing over the heat exchangers.

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LEGEND HOT BLEED AIR SERVICE AIR TO SURFACE DEICE COCKPIT/CABIN AIR SUPPLY EMERGENCY GEAR BLOWDOWN

9 PNEUMATICS

EMERGENCY BRAKES

Figure 9-1.   Mustang Pneumatic Systems

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The cooled, pressurized bleed air enters the pressure vessel through a set of valves. Pressure regulating shutoff valves (PRSOVs) ensure that a constant bleed-air pressure is maintained regardless of engine power settings. Flow control valves adjust flow to compensate for single-engine operation. The inboard bleed air from both engines supply temperature-controlled bleed air directly to the pressure vessel. Pressure vessel air supply is discussed in more detail in Chapter 11—“Air Conditioning.” The pressure vessel air supply exhausts overboard through nominal leakage in the cabin and through controlled venting by outflow valves in the aft cabin pressure bulkhead. The outflow valves are controlled by the pressurization system to maintain adequate cabin pressure at all altitudes. The pressurization system is explained in detail in Chapter 12—“Pressurization.”

This system is explained in Chapter 14—“Landing Gear and Brakes.”

LIMITATIONS For specific limitations, refer to the FAA-approved AFM.

EMERGENCY/ ABNORMAL For specific information on emergency/abnormal procedures, refer to the appropriate checklist or FAA-approved AFM.

COMPRESSED NITROGEN BOTTLES Separate bottles of pressurized nitrogen supply emergency pneumatic power for emergency landing gear extension and emergency braking.

Landing Gear Emergency Extension (Blow-Down) 9 PNEUMATICS

An independent emergency pneumatic system uses pressurized nitrogen in a bottle for emergency landing gear extension (blow-down). The high-pressure nitrogen bottle is attached to the right forward bulkhead inside the nose baggage compartment. Emergency gear extension is pilot-activated with the AUXILIARY GEAR CONTROL handle. This system is explained in Chapter 14—“Landing Gear and Brakes.”

Emergency Brakes An independent emergency pneumatic system uses pressurized nitrogen in a bottle for emergency braking. The high-pressure nitrogen bottle is attached to the right forward bulkhead inside the nose baggage compartment. Emergency braking is pilotactivated with the EMERGENCY BRAKE handle.

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QUESTIONS 1. The systems that use pneumatics for operation are: A. Instrument air, emergency brakes, and the entrance door B. Surface deice, engine anti-ice, pressurization, emergency gear extension, emergency brake operation C. Entrance door seal, air-cycle machine, and thrust reversers D. Windshield anti-ice, entrance door seal, and air-cycle machine 2. The left and right PRSOVs: A. Switch settings of flow control valves if one powerplant fails B. Provide 23-psi service air to the main cabin door seal C. Regulate bleed-air inflow regardless of power settings D. Duct bleed air to windshield heat

9 PNEUMATICS

3. Nitrogen bottles for emergency brake and emergency gear extension are on the: A. Inside of the aft storage compartment B. Right forward bulkhead C. Left engine nacelle D. Center spar inside the gear compartment

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CHAPTER 10 ICE AND RAIN PROTECTION CONTENTS INTRODUCTION................................................................................................................ 10-1 GENERAL..............................................................................................................................10-1 ENGINE INLETS................................................................................................................. 10-3 Description.................................................................................................................... 10-3 Components................................................................................................................... 10-3 Controls and Indications................................................................................................ 10-3 Operation....................................................................................................................... 10-5 SURFACE DEICE (WING AND STABILIZERS)............................................................... 10-5 Description.................................................................................................................... 10-5 Components................................................................................................................... 10-6 Controls and Indications................................................................................................ 10-6 Operation....................................................................................................................... 10-7 WINDSHIELD ICE AND RAIN PROTECTION................................................................ 10-7 Description and Components........................................................................................ 10-8 Controls and Indications............................................................................................. 10-10 Operation.................................................................................................................... 10-10 SENSOR ANTI-ICE SYSTEMS....................................................................................... 10-10 Description and Components..................................................................................... 10-10 Controls and Indications............................................................................................. 10-10 LIMITATIONS................................................................................................................... 10-11 EMERGENCY/ABNORMAL........................................................................................... 10-11 QUESTIONS..................................................................................................................... 10-13

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ILLUSTRATIONS Figure Title Page 10-1.

Citation Mustang Ice-Protection Systems............................................................. 10-2

10-2.

Ice-Protection Switches......................................................................................... 10-3

10-3.

Bleed Air/Pneumatic Ice-Protection Systems........................................................ 10-4

10-4.

Wing Stabilizer Automatic Deice Cycle................................................................ 10-7

10-5.

Windshield Anti-ice and Defog Zones................................................................... 10-8

10-6.

Windshield Anti-ice Power Distribution................................................................ 10-9

10-7.

Windshield Anti-Ice Switches............................................................................ 10-10

TABLES Table Title Page CAS MESSAGES............................................................................................... 10-12

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CHAPTER 10 ICE AND RAIN PROTECTION

INTRODUCTION This chapter describes the ice and rain protection systems for the Citation Mustang. Anti-icing is provided for the engine inlets, instrument external sensors, and windshields. Deicing is provided for the wings as well as the horizontal and vertical stabilizers. Rain protection is also provided for the windshield.

Flight into known icing is the intentional flight into icing conditions that are known to exist by either visual observation or pilot weather report information. Icing conditions exist any time the indicated ram air temperature (RAT) is +10°C or below, and visible moisture in any form is present. Engine anti-ice should be selected ON anytime the indicated ram RAT is +10°C or below, and visible moisture in any form is present. WING/

STAB DEICE should be selected as soon as ice is observed to accrue anywhere on the airplane. If ice remains on the airplane during approach and landing, maximum flap extension is limited to the TO/APR position. Ice accumulations significantly alter the shape of airfoils and increase the weight of the airplane. Flight with ice accumulated on the airplane will increase stall speeds and alter the speeds for opti-

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mum performance. Flight at high angle-of-attack (low airspeed) can result in ice building on the underside of the wings and the horizontal stabilizer aft of areas protected by deice boots. Minimum sustained airspeed for flight in icing conditions (except approach and landing) is 160 KIAS. Prolonged flight with the flaps and/or landing gear extended is not permitted except as required for approach and landing. Use of Flaps LAND (30°) is prohibited when any ice is observed adhering to the outside of the airplane. Trace or light amounts of icing on the horizontal stabilizer can significantly alter airfoil characteristics which will affect stability and control of the airplane.

it of ice that usually forms on the leading edges of wings, tail surfaces, pylons, engine inlets, and antennas, etc. Flight crews are to make sure that the airplane is free from ice prior to dispatch.

NOTE

Each engine has two bleed-air ports: one inboard and one outboard. The inboard port provides bleed air for interior air conditioning and pressurization. The outboard port provides bleed air for engine inlet anti-ice and for inflating the deice boots. During single-engine operation, check valves in the supply lines from each engine prevent bleed air from one engine back-flowing to the opposite engine.

With residual ice on the airplane, stall characteristics are degraded and stall speeds are increased. Freezing rain and clear ice will be deposited in layers over the entire surface of the airplane and can “run back” over control surfaces before freezing. Rime ice is an opaque, granular, and rough depos-

The Mustang uses conventional methods of ice protection. The engine inlets are anti-iced using engine bleed air. The wing as well as the horizontal and vertical stabilizer leading edges are protected using pneumatic deice boots. Electrical power protects the windshield, pitot probes, static ports, stall warning vane, and engine T2 probes. A passive rain repellent coating on the windshield provides clear vision in precipitation conditions (Figure 10-1).

HORIZONTAL AND VERTICAL STABILIZER DEICE SYSTEM (PNEUMATIC BOOTS)

ENGINE ANTI-ICE SYSTEM (BLEED AIR) T2 PROBES (ELECTRIC HEAT)

WING DEICE SYSTEM (PNEUMATIC BOOTS) WINDSHIELD ANTI-ICE SYSTEM (ELECTRIC HEAT) PITOT-STATIC SENSORS (ELECTRIC HEAT) STALL WARNING TRANSDUCER (ELECTRIC HEAT)

10 ICE AND RAIN PROTECTION

PITOT-STATIC SENSORS (ELECTRIC HEAT)

WING DEICE SYSTEM (PNEUMATIC BOOTS)

Figure 10-1.   Citation Mustang Ice-Protection Systems

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The ice protection system is controlled by greencapped switches on the pilot tilt panel, which is to the left of the LANDING GEAR handle (Figure 10-2). The status of ice protection systems is displayed by messages on the crew alerting system (CAS).

COMPONENTS Engine Anti-Ice Shutoff Valves The hot bleed air flows into each engine inlet through the engine anti-ice pressure-regulating shutoff valve (PRSOV) (Figure 10-3). When the valve energizes closed, the inlet is not anti-iced. The engine inlet PRSOVs are operated by the ENGINE ANTI-ICE switches and are electrically actuated. In the absence of electrical power, the valves are pushed open by bleed-air pressure. When electrical power is applied, a solenoid powers the PRSOV closed. The PRSOVs control the air pressure downstream of the valve. These valves regulate the airflow of the engine anti-ice system.

Engine Inlet Anti-Ice Undertemperature Switches

Figure 10-2.   Ice-Protection Switches

ENGINE INLETS

The engine inlet undertemperature switches are inside the nacelle leading edge. These switches provide information through the monitoring circuits in the left and right ice protection system printed circuit boards (PCBs) to drive the ENG A/I COLD L-R and ENG A/I COLD L-R CAS messages.

Engine Inlet Anti-Ice Assembly

DESCRIPTION Each engine inlet and the inlet of the generatorcooling scoop is heated by regulated engine bleed air. Temperature of the bleed air is directly related to throttle position. Spent bleed air exits via a vent in the bottom of the inlet. This vent is inspected during preflight.

NOTE There is no crossfeed between engines for inlet heating. If an engine fails, its inlet is no longer heated.

The engine inlet leading edge is hollow. Inside the leading edge, a circular piccolo tube is immediately behind the forward surface of the inlet. The hot bleed air enters the piccolo tube, sprays out of holes in the tube to circulate through the inlet leading edge, then exhausts overboard through a vent in the bottom of each inlet assembly.

CONTROLS AND INDICATIONS Controls for the engine anti-ice system are on the pilot tilt panel with the other ICE PROTECTION controls (see Figure 10-2).

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DC power is provided through an L or R A/I circuit breaker on its respective side ELE #2 bus. If DC power fails, the engine anti-ice valves automatically open, allowing hot air into the nacelle leading edges.

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REGULATED SERVICE AIR

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OOT VERTICAL STAB DEICE BOOT

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HOT HP ENGINE BLEED AIR

LEGEND

SERVICE AIR CHECK VALVES

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Figure 10-3.   Bleed Air/Pneumatic Ice-Protection Systems

DEICE BOOT PRESSURE SWITCHES

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ENGINE INLET UNDERTEMP SWITCH

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The L and R ENGINE ANTI-ICE switches control the flow of hot bleed air to the engine inlet leading edges. Each ENGINE ANTI-ICE switch has two positions: L (or R) and OFF. With the switches in the L or R (up) position, bleed air is directed to the engine inlet for the respective engine. This position also energizes the T2 probe and with the landing gear retracted sets the flight idle N2 speed up to 70%. With the switch in the OFF position bleed air is blocked at the engine antiice PRSOV, the T2 probe is deenergized and the flight idle speed is set back to a minimum of 56.8%. Placing either ENGINE ANTI-ICE switch in the ON position activates the undertemperature warning system for both engines.

Throttles The temperature of the air supplied to the engine inlets is varied only by engine power settings.

OPERATION When in icing conditions or when anticipating icing conditions, set the ENGINE ANTI-ICE switches to the L and R (up) positions. This deenergizes the shutoff valves, allowing hot bleed air to flow through and heat the engine. When not in (or anticipating) icing conditions, set the ENGINE ANTI-ICE switches to the OFF positions. This energizes the shutoff valves closed, stopping the flow of bleed air to the engine inlets. This also increases engine efficiency and available power.

SURFACE DEICE (WING AND STABILIZERS) DESCRIPTION The Mustang uses pressure-regulated engine bleed air (via the service air system) to operate conventional pneumatic deicing boots. The full-span boots protect the wing, vertical, and horizontal stabilizer leading edges.

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Service Air System The pressurized air for inflating the pneumatic boots is supplied by the service air system. The service air is regulated to 20 psig by a service air regulator (see Figure 10-3). The service air system is always active during operation of the aircraft. Bleed air from both engines is routed to the service air regulator through a check valve on each engine outboard bleed-air supply duct. The check valves keep bleed air from backflowing into either engine. The service air system provides regulated bleed air to the deice boot system. If one engine fails, the operating engine can supply enough bleed air to operate the wing, vertical, and horizontal stabilizer deice system.

Surface De-Ice System The Mustang has a surface deice system on the wing, vertical, and horizontal stabilizers. This system uses regulated bleed air to inflate pneumatic boots to remove the ice. The boots, when inflated, normally crack and separate the ice from the leading edge of the protected surface, allowing aerodynamic forces to remove the ice. During normal operation, adequate pressure supplied to the boots is annunciated by the SURFACE DE-ICE message. The WING DE-ICE FAIL or TAIL DE-ICE FAIL message appears in the CAS window if boot pressure is inadequate or boot inflation cycle is not normal. The de-ice boot pressure switches are immediately downstream of each ejector flow control valve (EFCV). The regulated service air is supplied to the wing and tail EFCVs, which supply either pressure or vacuum to the deice boots. There is one EFCV for each of the following four boot sets (see Figure 10-1): • Wing upper de-ice boots (left and right) • Wing lower de-ice boots (left and right) • Left horizontal stabilizer deice boot • Right horizontal de-ice boot and vertical stabilizer deice boots The EFCVs are electrically powered closed to inflate the boots and spring-loaded open.

FOR TRAINING PURPOSES ONLY

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10 ICE AND RAIN PROTECTION

L and R ENGINE ANTI-ICE Switches

CITATION MUSTANG PILOT TRAINING MANUAL

The surface de-ice system operates in one of three modes: • Manual • Automatic • Inactive

COMPONENTS

The switches are set to close at 16 psig and open when the pressure decreases to below 10 psig. They provide information controlling the CAS messages for this system.

CONTROLS AND INDICATIONS WING/STAB Switch

Service Air Regulator The service air regulator reduces the pressure of the engine bleed air to 20 psig from a variable of 25–200 psi. The pressure relief setting of the valve is 27 psig. If the service air regulator regulates too low, the failure is detected by the deice system monitors. If the service air regulator regulates too high, the pressure supplied to the boots is limited to 27 psig by the relief port of the valve.

The control switch for the surface deice system is on the pilot side of the instrument panel with the other ICE PROTECTION controls (see Figure 10-2). The WING/STAB switch has three positions: OFF, AUTO and MANUAL. In the OFF position, no power is supplied to the EFCVs and service air flows through the valves to create a vacuum that holds the boots deflated. The AUTO position activates the deice control and monitor boards which run the 2-minute boot cycle.

Surface De-Ice Boots The wing, horizontal stabilizer, and vertical tail are de-iced by pneumatic boots controlled by the WING/STAB switch. The wing boots are separated into two independent pneumatic chambers: one for the upper surface and one for the lower surface. Each stabilizer boot has one pneumatic chamber. All boots have spanwise tube configurations.

The MANUAL position is spring-loaded and is active only while held in that position. When the switch is in the MANUAL position, power is supplied to the EFCVs to apply pressure to the boots. It also supplies a signal to the deice monitor board to check for adequate pressure supplied to the boots.

Surface De-Ice Control Valves

System Monitoring And Indications

The wing and stabilizer EFCVs are electrically controlled switching valves. When the deice system is turned off or not being inflated, vacuum is applied to the boots by the EFCVs. This is done by passing the supplied service air over a venturi in the valve and is then vented overboard. The other end of the venturi is connected to the boot and the flow created in the venturi creates the vacuum at the boot. When an inflation is triggered, a solenoid closes the EFCV vent and the service air then routes through the venturi to inflate the boot. 10 ICE AND RAIN PROTECTION

Surface De-Ice Pressure Switches

De-ice boot inflation is monitored by a series of pressure switches. One switch is provided for each de-ice boot chamber. The right ice protection system PCB monitors the pressure switches to verify the deice boots inflate when commanded by the left ice protection system PCB. If the deice boots fail to inflate, a discrete output is provided to the CAS to announce the failure.

OPERATION The wing and stabilizer de-ice system operates under electrical control when set with the WING/ STAB switch in MANUAL or AUTO.

The boot pressure switches are immediately between the EFCV and their boots. There are four switches for the wing (two per side) and two switches for the vertical and horizontal stabilizers. 10-6

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CITATION MUSTANG PILOT TRAINING MANUAL

Manual Mode

As in the manual mode, if all of the appropriate boot pressure switches receive adequate pressure, the SURFACE DE-ICE CAS message appears. If one or more of the pressure switches do not receive adequate pressure or if the boots do not inflate in the proper timed sequence, the appropriate amber WING DE-ICE FAIL and/or TAIL DE-ICE FAIL CAS message appears.

During operation in manual mode, the pilot holds the WING/STAB switch in the MANUAL position to inflate all of the deice boots. As long as the pilot holds the switch, the de-ice boots remain inflated. If all of the boot pressure switches receive adequate pressure, the SURFACE DE-ICE CAS message appears.

Inactive

If one or more of the pressure switches does not receive adequate pressure, the appropriate amber WING DE-ICE FAIL and/or TAIL DE-ICE FAIL CAS message appears.

Whenever the boots are not being inflated, vacuum is applied to the boots to hold them down. This is true in automatic mode when not in the inflation period of the deice cycle. This is also true whenever the system is set to OFF.

Following release of the MANUAL switch, the spring-loaded switch immediately returns to the AUTO position, but the automatic cycle is delayed for 2 minutes before restarting the inflation sequence. This prevents the lower wing/left horizontal stabilizer deice boots from immediately inflating after the pilot releases the switch from the MANUAL position.

WINDSHIELD ICE AND RAIN PROTECTION The Mustang glass windshields include electric anti-icing/defogging and have a rain repellent applied. Individual sections of each windshield have different levels and sources of protection.

Automatic Mode During automatic mode, the crew selects the WING/STAB switch to AUTO. This activates a timer, which causes the boots to inflate and deflate in a sequence that repeats every 2 minutes (Figure 10-4). The sequence repeats continuously until the OFF or MANUAL switch positions are selected. When selected OFF, the deice system completes the 2-minute inflation cycle to prevent any asymmetric ice accumulation. Figure 10-4 illustrates the automatic boot inflation cycle.

DESCRIPTION AND COMPONENTS The Mustang uses 28-VDC electric power to provide windshield anti-ice and defog capability (Figure 10-5).

LEGEND PRESSURE APPLIED TO DEICE BOOTS VACUUM APPLIED TO DEICE BOOTS

TIME (SEC) 0

6

12

120

10 ICE AND RAIN PROTECTION

LOWER WING BOOTS LEFT HORIZONTAL STABILIZER UPPER WING BOOTS RIGHT HORIZONTAL STABILIZER VERTICAL STABILIZER

Figure 10-4.   Wing Stabilizer Automatic Deice Cycle

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Each windshield incorporates wire filament heaters in three separate zones: • Inner zone—Provides both anti-ice and defog capability. • Middle zone—Provides both anti-ice and defog capability. • Defog zone—Provides defog capability only.

NOTE The upper region of each zone provides reduced heat. However, it is powered as part of the corresponding anti-ice or defog zones.

feed bus through an extension bus, then through the corresponding L or R WSHLD TEMP circuit breaker (on the same side CB panel). The left and right controller powers the outer and middle portions of the windshield on its respective side and also powers the inner panel on the opposite side windshield. The aircraft has two separate windshield anti-ice controllers. The left generator operates the left windshield controller and the right generator operates the right windshield controller. Each controller can recognize a loss of generator DC power to the other.

DEFOG ZONE DEFOG ZONE ANTI-ICE ZONE TRANSITION REGION

ANTI-ICE ZONES

UNPROTECTED REGION

Figure 10-5.   Windshield Anti-ice and Defog Zones

Each anti-ice zone has two temperature sensors (Figure 10-6). The primary sensor controls the temperature and monitors for overheats. If the primary sensor fails, the corresponding secondary sensor provides those functions.

10 ICE AND RAIN PROTECTION

Electrical power is distributed to each windshield by an anti-ice controller. Each controller receives DC power from two sources: • Heating power comes to each controller directly from its respective feed bus through a 100 amp current limiter. • Control and monitoring functions of each controller are powered from the respective

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Loss of L or R Generator Power Configuration AG aircraft 510-0001 through 0040—With the loss of L or R generator power, the operating windshield controller only provides power to one panel on the pilots side windshield. If the right generator fails then the right windshield controller is load shed and the left windshield controllers is still operating and provides power to the outboard anti-ice panel on the pilot side only. Conversely if the left generator fails then the left windshield controller is load shed and the right windshield controller is still operating and provides power to the inboard anti-ice panel on the pilot side only. Configuration AF aircraft 510-0041 and subsequent—With the loss of L or R generator power, the operating windshield controller will provide power to one panel on the pilots side and one panel on the copilots side windshield. If the right generator fails then the right windshield controller is load shed and the left windshield controller is still operating. In this case the operating left controller provides power to the outboard anti-ice panel on the pilot side and to the inboard panel on the copilot side. Conversely if the left generator fails then the left windshield controller is load shed and the right windshield controller is still operating. In this case the operating right controller provides power to the inboard anti-ice panel on the pilot side and to the outboard panel on the copilot side.

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UNPROTECTED AREA

DEFOG

DEFOG

DEFOG

DEFOG

MID

INNER

INNER

MID

A/I

A/I

A/I

A/I

DEFOG

2

2

3

DEFOG

3

1

1

RTD (TEMP SENSOR) UNPROTECTED AREA

UNPROTECTED AREA

1

2

3

2

LH CONTROLLER

100 A MP

LH FEED BUS

1

3

RH CONTROLLER

200 A MP

100 A MP

200 A MP

CROSSFEED BUS

RH FEED BUS

LEGEND LH CONTROLLER RH CONTROLLER

If both generators are operating and only one of the windshield anti-ice controllers fail, then only the portions of the windshield powered by the failed controller are lost.

tion. The coating requires periodic inspection and refurbishment. The windshield should only be cleaned with a soft cloth and water to preclude damaging the coating.

Rain Repellent A passive rain repellent coating on the windshield external surface provides windshield rain protec-

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10 ICE AND RAIN PROTECTION

Figure 10-6.   Windshield Anti-ice Power Distribution

CITATION MUSTANG PILOT TRAINING MANUAL

CONTROLS AND INDICATIONS WINDSHIELD ANTI-ICE Switches The windshield heat is controlled by two toggle switches in the cockpit (one for each windshield panel) (Figure 10-7). The L and R WINDSHIELD ANTI-ICE switches have two positions: • A NTI-ICE (up) applies power to both the defog and anti-ice zones. • OFF removes power from the system.

SENSOR ANTI-ICE SYSTEMS DESCRIPTION AND COMPONENTS Electric heat is provided to anti-ice the following sensors: • Pitot probes • Static ports • Stall warning vane • T2 probes The heating element for each sensor is monitored by a current sensor to detect failures. Failure of any heating element is indicated on the CAS display.

CONTROLS AND INDICATIONS Sensor Anti-Ice Switch Figure 10-7.   Windshield Anti-Ice Switches

Windshield Ice Detection Lights Two red LED ice detection lights are located at the base of the windshield on the pilot and copilot glareshield. These two lights are powered any time the BATT switch is placed in the ON position. If the windshield is clean, the lights shine through and are undetected. If ice forms on the windshield, the lights reflect back giving the pilot a visible indication at night.

OPERATION When WINDSHIELD ANTI-ICE is selected, the windshield controllers provide a slow increase in temperature to avoid thermal shock to the windshield panels.

The pitot probes, static ports, and stall warning vane heaters are all controlled by a single switch on the ICE PROTECTION panel immediately left of the landing gear handle (see Figure 10-2). This switch has three positions: • RESET STALL WARN—A momentarycontact position. Resets the stall warning to the normal stall airspeed. (Use RESET STALL WARN only when wings are verified free of ice). • PITOT STATIC—Applies power to the sensors (both pitot probes, all four static ports, and the stall warning vane). • O FF—Removes all power from those sensors.

ENGINE ANTI-ICE Switches

10 ICE AND RAIN PROTECTION

The T2 probes are electrically heated when their respective ENGINE ANTI-ICE switches are in the L or R position and the engine is running.

Each windshield controller monitors the windshield temperature sensors in the zones it controls. Using this information, it provides discrete outputs to the CAS for annunciation of controller failures or windshield overheats.

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CITATION MUSTANG PILOT TRAINING MANUAL

OPERATION

LIMITATIONS

The pitot probe, static port, and stall warning vane heaters are powered by selection of the PITOT STATIC position on the sensor anti-ice switch.

For specific limitations, refer to the FAA-approved AFM.

In flight, the sensor anti-ice switch should be in the PITOT STATIC position, which heats the external sensors. On ground, except when ready for takeoff, the switch should normally be OFF to prevent overheating of the sensors and their heating elements.

EMERGENCY/ ABNORMAL For specific information on emergency/abnormal procedures, refer to the appropriate checklist or FAA-approved AFM.

During preflight, the switch may be set to PITOT STATIC for 30 seconds to verify the sensors are heating properly.

CAUTION Limit ground operation of pitot-static heat to 2 minutes to preclude damage to the pitot-static and stall warning heaters.

Stall-Warning System Mode When surface deice is enabled at any time (the WING/STAB switch is selected to MANUAL or AUTO), the stall-warning system changes its mode to a higher airspeed, and does not reset when surface deice is switched OFF. The system remains at this ice-contamination airspeed mode setting until the end of the flight or until RESET STALL WARN is selected. Selecting RESET STALL WARN on the PITOT STATIC switch overrides the automatic ice-contamination setting, and returns stall-warning mode to the normal stalling airspeed, if surface deice is selected OFF.

FOR TRAINING PURPOSES ONLY

10 ICE AND RAIN PROTECTION

The STALL WARN HI CAS message indicates that the stall-warning system is operating on the ice-contamination airspeed mode. Refer to the “Landing Performance” ANTI-ICE-ON landing performance charts in the Airplane Flight Manual (AFM).

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Table 10-1.   CAS MESSAGES ENG A/I COLD L-R DESCRIPTION

INHIBITS

WING DE-ICE FAIL

Indicates the engine inlet temperature is below safe level for satisfactory ice protection.

DESCRIPTION INHIBITS

This message indicates the wing de-ice system is not operating normally. EMER

EMER, ESDI ENG A/I COLD L-R

P/S HTR L-R

DESCRIPTION

DESCRIPTION This message indicates there is no current detected to the pitot static heater and the airplane is in the air or throttles are at or above the cruise detent. INHIBITS

EMER, LOPI INHIBITS

Indicates the engine inlet temperature is below safe level for satisfactory ice protection. This message will post white for up to two minutes after engine anti-ice is turned one while the inlet warms up to the normal operating temperature. EMER, ESDI

STALL WARN HTR DESCRIPTION

INHIBITS

This message indicates no power is being delivered to the stall warning vane heater.

P/S HTR L-R DESCRIPTION

EMER, ESDI INHIBITS

This message indicates there is no current detected to the pitot static heater and the airplane is on the ground. EMER, LOPI

T2 HTR FAIL L-R DESCRIPTION INHIBITS

This message indicates a T2 probe heater fail.

SURFACE DE-ICE DESCRIPTION

EMER, LOPI, TOPI

TAIL DE-ICE FAIL DESCRIPTION INHIBITS

This message indicates the tail de-ice system is not operating normally. EMER

INHIBITS

This message indicates a loss of power to the windshield heater and the WINDSHIELD A/I ON has been ON for 5 seconds.

10 ICE AND RAIN PROTECTION

DESCRIPTION

INHIBITS

INHIBITS

This message indicates the windshield anti-ice power is ON for more than 5 seconds and the windshield temperature is too high.

DESCRIPTION

INHIBITS

EMER, ESI

This message indicates a loss of power to the windshield heater and the WINDSHIELD A/I ON has been ON for less than 5 seconds. EMER

W/S O’HEAT L-R DESCRIPTION

EMER

INHIBITS

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The stall warning system is operating on the ice-contaminated schedule.

W/S A/I FAIL L-R

EMER

W/S O’HEAT L-R

EMER

STALL WARN HI DESCRIPTION

W/S A/I FAIL L-R DESCRIPTION

INHIBITS

The De-ice boots are inflating/deflating as designed. In MANUAL mode, this message displays only if all pressure switches indicate deice boot inflation.

FOR TRAINING PURPOSES ONLY

This message indicates the windshield anti-ice power is ON for less than 5 seconds and the windshield temperature is too high. EMER

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CITATION MUSTANG PILOT TRAINING MANUAL

QUESTIONS

2. The ENG A/I COLD R CAS message may indicate: A. The system is operating normally after initial switch actuation B. The system has failed C. The ejector flow control valve has failed D. The ejector transfer pump has failed 3. With the WING DE-ICE FAIL CAS message displayed: A. The wing boots have failed to inflate adequately B. The ejector transfer pump has failed C. The motive flow pump has failed D. The left weight-on-wheels switch has failed 4. With the W/S A/I FAIL L CAS message displayed, the checklist states to: A. L WSHLD TEMP circuit breaker—Check B. Rotate the rotary TEST knob to ANNU C. Push the TMR/REF soft key on the PFD D. PRESS CTRL switch—STANDBY 5. If the W/S O’HEAT L CAS message remains displayed continuously: A. The temperature decreased using the CABIN TEMP selector B. The temperature decreased using the COCKPIT TEMP C. Power must be removed from the windshield using the L WINDSHIELD ANTIICE switch. D. Power removed using the rotary TEST knob

Revision 1.1

6. With a P/S HTR R message displayed, an erroneous indication may also appear on the: A. Airspeed indicator B. Magnetic compass C. EICAS D. Standby gyro 7. Do not operate the surface deice boots with an indicated RAT below: A. –40°C B. –30°C C. –20°C D. –10°C 8. Failure of the DC electrical system results in: A. Complete failure of the windshield antiicing system B. Continuous flow of hot bleed air, with windshield temperature control possible only through regulation of the volume of bleed air permitted to the windshield C. Continued windshield anti-icing with complete control of the bleed-air temperature D. Continuous isopropyl alcohol flow to the windshield to replace the normal bleed-air anti-icing 9. When using the WING/STAB switch in the MANUAL position: A. They may be cycled at any temperature B. All of the deice boots inflate simultaneously for as long as the switch is held C. Illumination of the SURFACE DE-ICE CAS message always indicates a system malfunction D. MANUAL mode serves as a backup way to inflate the door seal 10. The pitot tubes and stall warning vane have a ground limitation of: A. 1 minute B. 2 minutes C. 3 minutes D. 4 minutes

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10 ICE AND RAIN PROTECTION

1. The ENG A/I COLD L CAS message indicates: A. The system is operating normally after initial switch actuation B. The system has failed C. The ejector flow control valve has failed D. The ejector transfer pump has failed



CITATION MUSTANG PILOT TRAINING MANUAL

11. The ENGINE ANTI-ICE switches must be selected ON when in visible moisture and a RAT of _______ or colder. A. 30°C B. 20°C C. 10°C D. 0°C 12. If the left generator fails: A. The inner zone on the pilot’s windshield and the middle zone on the co-pilot’s windshield are heated B. The engine A/I system becomes inoperative C. The right side of the pitot-static system becomes inoperative D. The stall warning vane heat becomes inoperative 13. If the right generator fails: A. The left side of the pitot-static system becomes inoperative B. The engine anti-ice system becomes inoperative C. The center panel on the pilot’s and the inner panel on the co-pilot’s windshields are heated D. The stall warning vane heat becomes inoperative 14. With the STALL WARN HI CAS message displayed, this indicates: A. The stall warning vane has an overtemperature situation B. The low airspeed awareness tap is operating on a contamination schedule C. The pitot-static switch has failed D. The stall warning system has failed

10 ICE AND RAIN PROTECTION

15. Except for the ground preflight check, the maximum RAT for operation of engine antiice with the throttles above idle is: A. 40°C B. 30°C C. 20°C D. 10°C

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CHAPTER 11 AIR CONDITIONING CONTENTS INTRODUCTION................................................................................................................ 11-1 GENERAL..............................................................................................................................11-1 TEMPERATURE-CONTROLLED BLEED-AIR INFLOW................................................ 11-2 Description.................................................................................................................... 11-2 Components................................................................................................................... 11-4 Controls and Indications................................................................................................ 11-5 Operation....................................................................................................................... 11-7 VAPOR-CYCLE AIR CONDITIONING............................................................................. 11-8 Description.................................................................................................................... 11-8 Components................................................................................................................... 11-8 Controls and Indications............................................................................................. 11-10 Operation.................................................................................................................... 11-10 FRESH AIR AND FANS................................................................................................... 11-10 Description................................................................................................................. 11-10 Controls and Indications............................................................................................. 11-11 Operation.................................................................................................................... 11-11 LIMITATIONS................................................................................................................... 11-11 EMERGENCY/ABNORMAL........................................................................................... 11-11 QUESTIONS..................................................................................................................... 11-12

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ILLUSTRATIONS Figure Title Page 11-1.

Cabin Bleed-Air Schematic................................................................................... 11-3

11-2.

Pylon Ram-Air Inlet.............................................................................................. 11-4

11-3.

Exhaust Ports......................................................................................................... 11-4

11-4.

Environmental Control Panel................................................................................. 11-6

11-5.

AIR SOURCE SELECT Knob.............................................................................. 11-6

11-6.

COCKPIT TEMP and CABIN TEMP Knobs........................................................ 11-7

11-7.

Vapor-Cycle Schematic......................................................................................... 11-9

11-8.

Evaporator Drains (Aft)...................................................................................... 11-10

11-9.

AIR COND Switch............................................................................................. 11-10

11-10. COCKPIT FAN and CABIN FAN Knobs.......................................................... 11-10

TABLES Table Title Page 11-1.

CAS MESSAGES............................................................................................... 11-11

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CHAPTER 11 AIR CONDITIONING

INTRODUCTION This chapter describes the Citation Mustang air-conditioning systems. Information is provided on temperature-controlled pressurized air, vapor-cycle air conditioning, and fresh air supply. Additionally, air distribution and temperature controls are discussed.

GENERAL The Mustang uses modified engine bleed air to heat, cool, and pressurize the cockpit and cabin (pressure vessel). Bleed-air inflow also defogs cabin and cockpit windows. The hot engine bleed air is cooled by heat exchangers, regulated by valves, and enters the pressure vessel through separate ducts on the left and right sides. Outflow valves regulate the outflow of this conditioned air supply to control the air pressure (and resulting pressure altitude) of the pressure vessel (refer to Chapter 12—“Pressurization”).

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The pressure vessel supply is normally provided by bleed air from both engines; however, either engine can supply an adequate bleed air supply to the pressure vessel. The system can operate with complete or partial loss of DC power; however, temperatures will be within 30°C above ambient. A conventional vapor-cycle air-conditioning system provides further cooling and defogging, especially on the ground or at low altitudes on hot days. It moves refrigerant fluid through heat exchangers to extract heat from the pressure vessel, then

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routes the fluid to another heat exchanger to vent the heat overboard. If either generator fails in flight, the vapor-cycle system is loadshed (powered off to conserve DC power). A separate fresh-air vent system allows direct cockpit and cabin ventilation when unpressurized (at low altitudes or on the ground). It requires DC power. These three air-conditioning subsystems (bleed air, vapor cycle, and fresh-air supply) are regulated by air-distribution and temperature-control systems. Engine indicating and crew alerting system (EICAS) indications advise the crew of system status.

TEMPERATURECONTROLLED BLEED-AIR INFLOW In flight, pressure vessel temperature is provided by very hot bleed air from the engines. This hot bleed air is cooled (conditioned) by heat exchangers before entering the pressure vessel. This conditioned air can heat the pressure vessel as well as cool it.

DESCRIPTION Bleed-Air Supply The left engine supplies air to the cockpit (Figure 11-1). The right engine supplies air for the cabin area. Each system is separate and independent, so that a failure of either system does not prevent the other from operating. All bleed-air inflow to both zones flows aft through the pressure vessel and exits through the outflow valves in the aft pressure bulkhead (refer to Chapter 12—“Pressurization”).

Bleed-Air Temperature Control Before entering the pressure vessel, hot, highpressure bleed air from each engine passes through an air-to-air heat exchanger in the engine pylon. Outside (ambient) ram-air enters the pylon ram air 11-2

inlet, passes through the heat exchanger, flows over the bleed-air ducts, and exits overboard through a temperature control valve, carrying away most of the heat from the engine bleed air. (The bleed air and ambient ram air do not mix.) A temperature control valve limits the ambient ram-air flow through each ram-air duct (and through the heat exchanger) to limit how much the heat exchanger can cool the engine bleed air. This indirectly controls the temperature of the bleed air entering the pressure vessel. For each system (cockpit/left engine and cabin/ right engine) the temperature control valve is positioned by a thermal actuator, controlled by a corresponding (cockpit or cabin) environmental printed circuit board (PCB) (Figure 11-1). The environmental PCB responds to crew commands and temperature sensors in the aircraft.

Bleed-Air Flow to Cabin Air Supply On each side, the cooled high-pressure bleed air continues into the cockpit or cabin air supply duct through a series of valves. Mechanical check valves prevent reverse flow of the bleed air into the engines. Pressure regulating shutoff valves (PRSOV) on the left and right engines, allows the bleed air supply to the pressure vessel from either one or both engines to be shut off. Under normal conditions the PRSOVs are both open. With the loss of DC power the PRSOVs fail open, electrical power is required to power them shut. Flow control valves (FCV) on the left and right engines regulate the bleed air supply to the pressure vessel from that engine to either 4 or 8 ppm (pounds per minute). Under normal conditions with both the left and right PRSOVs open, the left and right FCVs are set at 4 ppm. With 4 ppm supplied from both engines this gives a combined supply of 8 ppm to the pressure vessel. If either the left or right PRSOV is commanded shut by the pilot or if an engine fails, the opposite FCV is set to 8 ppm keeping the total bleed supply to the pressure vessel constant.

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COCKPIT FOOTWARMER VENTS

FRESH AIR INLET

T

FOR TRAINING PURPOSES ONLY CABIN FOOT WARMER VENTS T

FLOW CONTROL VALVE

AFT PRESSURE BULKHEAD

CABIN TEMPERATURE SENSOR

DUCT OVERHEAT SENSOR

Figure 11-1.   Cabin Bleed-Air Schematic

CABIN SHOULDER COCKPIT WARMERS TEMPERATURE SENSOR

COCKPIT SHOULDER WARMER

FWD PRESSURE BULKHEAD

PYLON RAM AIR

T

T T 4 8

T

T

4 8

HOT BLEED AIR FROM ENGINE

COCKPIT ENVIRONMENTAL PCB

11 AIR CONDITIONING

EXHAUST OVERBOARD

HOT BLEED AIR

RAM AIR

THERMAL ACTUATOR

EXHAUST OVERBOARD

CABIN ENVIRONMENTAL PCB

REGULATED AIR

LEGEND

PRESSUREREGULATING SHUTOFF VALVE

DUCT TEMP SENSOR

TAIL CONE TEMP SENSOR

HEAT EXCHANGER

TEMPERATURE CONTROL VALVE

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Bleed-Air Distribution Separate ducts route warm bleed air into the cockpit (from the left engine) and into the cabin (from the right engine). The cockpit (left engine) bleed-air system routes warm bleed air to the foot warmer outlets above the pilot and copilot rudder pedals. It also provides warm bleed air to defog the side windows. The cabin (right engine) bleed-air system routes warm bleed air to the cabin shoulder and foot warmer outlets. Air from both bleed-air systems moves aft through the cabin and exits through the outflow valves (refer to Chapter 12—“Pressurization”).

Figure 11-2.   Pylon Ram-Air Inlet

A check valve permits flow from the right engine (cabin) bleed-air system into the cockpit bleed-air system if the left engine fails. This keeps air flowing through the entire length of the aircraft interior.

COMPONENTS Heat Exchangers and Ram-Air Ducts Engine bleed air is cooled by a heat exchanger in each engine pylon. Engine bleed air enters and exits the precooler through bleed-air ducts. Outside (ambient) ram air enters through the pylon ram-air inlet on the leading edge of the pylon and flows into the ram-air duct. It then passes through the heat exchanger and flows over the bleed-air ducts (the bleed air and ambient ram air do not mix). Heat passes from the bleed air to the ram air through the metal walls of the separate ducts, cooling the bleed-air inside the bleed air ducts. The heated ram air exits overboard through a temperature control valve in the ram-air duct, carrying away most of the heat from the engine bleed air. The ram air exits through the aft pylon, into the engine exhaust stream. Forward of the pylon ram air exhaust port, an eductor projects into the engine exhaust stream, creating a vacuum behind it to pull the ram air from the pylon exhaust port. During preflight, check that the pylon ram air duct inlet and exhaust ports are clear (Figures 11-2 and 11-3).

11-4

Figure 11-3.   Exhaust Ports

Temperature Control Valves Aft of the heat exchanger in each ram air duct, a thermally actuated temperature control valve limits the ambient ram-air flow through the duct to limit how much cooling ram air flows through the heat exchanger. This determines how much the heat exchanger cools the engine bleed air and directly controls the temperature of the bleed air before it enters the pressure vessel.

Thermal Actuator A thermal actuator adjusts the position of the temperature control valve. The actuator responds to the temperature of a gas inside the actuator, which is heated by an electrical heating element. The electrical heating element is powered and controlled by an environmental temperature controller. If DC power is removed from the thermal actuator, it cools and retracts, opening its temperature con-

FOR TRAINING PURPOSES ONLY

trol valve to maximum cooling. If the actuator fails by leaking, a spring retracts it to set the temperature control valve to maximum cooling.

Duct Overheat Temperature Sensors

Cockpit and Cabin Environmental PCBs Each environmental PCB compares the temperature in its respective zone (cockpit or cabin) to the temperature setting selected by the crew for that zone. It then compares this to the temperature of the bleed-air supply in the ducts and powers the thermal actuator that adjusts the temperature control valve to allow more or less cooling air through the heat exchanger. The resulting heat exchanger temperature provides cooler or warmer bleed air to that zone. Each environmental PCB (cockpit/left and cabin/ right) is powered from the respective electrical feed bus. If DC power for either cockpit or cabin environmental PCB fails, the temperature control system for that zone fails to the full-cooling mode (30°F above ambient temperature).

Pressure Regulating Shutoff Valves Each pressure regulating shutoff valve (PRSOV) is a normally open electrically actuated solenoid valve. Selecting either L or R with the AIR SOURCE SELECT knob (Figure 11- 5) energizes the opposite (left or right) PRSOV closed, limiting bleed-air inflow to the cabin from only the selected side. The left PRSOV is powered from the right feed bus No. 2. If either PRSOV loses power, the valve is spring-loaded to the open position.

Flow Control Valves Each FCV has two openings: one allowing 4 ppm of bleed air to pass through the system from the corresponding engine, and the other allowing 8-ppm flow. Normally, both valves are deenergized to 4 ppm, for a total flow to the cabin of 8 ppm. If either engine fails, or the crew manually selects bleed air supply from one engine only, the FCV of the supplying engine opens to the 8-ppm setting.

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If power to either FCV is lost, the valve is springloaded to the 4-ppm position.

The duct overheat temperature sensors are downstream of the duct temperature sensors. The sensor opens or closes in response to the temperature of the bleed air in the duct. When closed, it sends a signal to the crew alerting system (CAS), which displays the AIR DUCT O’HEAT L (or R) message.

Duct Temperature Sensors The duct temperature sensors are immediately downstream of the FCV and upstream of the overheat sensors. These duct temperature sensors signal the temperature of the bleed air to the zone PCB for that side.

Zone Temperature Sensors The cockpit and cabin compartments each have a zone temperature sensor that detects the air temperature in that zone. The cockpit zone temperature sensor is behind the pilot instrument panel, below the cooling fan. The cabin zone temperature sensor is in the aft evaporator inlet (behind the aft cabin seats near the floor of the cabin). The environmental PCB compares this zone temperature with bleed-air temperature reported by the duct temperature sensors and crew settings of the temperature control knobs to determine necessary automatic changes to bleed-air temperature. Do not block the airflow at either of the zone temperature sensors. Obstructions to airflow causes errors in a sensor signal to its environmental temperature controller, resulting in incorrect temperature control for that zone.

CONTROLS AND INDICATIONS Controls specifically for the pressure vessel bleedair systems are on the ENVIRONMENTAL control panel below the copilot primary flight display (PFD) (Figure 11-4).

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Figure 11-4.   Environmental Control Panel

AIR SOURCE SELECT Knob The AIR SOURCE SELECT knob is a rotary electrical switch that selects the source(s) of air entering the pressure vessel. To do this, it controls the PRSOVs and FCVs for the bleed-air ducts. It also controls the fresh-air fan.

NOTE On the ground, when the AIR SOURCE SELECT knob is set to BOTH, if the temperature in the pressure vessel is greater than 18°C (65°F) and throttle position is at or below 85% throttle lever angle (TLA), the environmental PCBs automatically energize the PRSOVs closed. This prevents hot bleed air from entering the pressure vessel.

L (Left Engine)

Figure 11-5.   AIR SOURCE SELECT Knob

OFF The OFF position energizes both PRSOVs closed. This stops all bleed-air inflow and fresh-air inflow to the cockpit and cabin.

BOTH (Both Engines) The BOTH position deenergizes both bleed-air PRSOVs open, allowing temperature-controlled, pressurized bleed air to both cockpit and cabin zones from their respective engines (left for cockpit, right for cabin). This also deenergizes both FCVs to 4 ppm, for a total flow of 8 ppm to the pressure vessel. 11-6

The L position energizes the right PRSOV closed and energizes the left FCV to 8 ppm. Bleed air supply to the pressure vessel is from the left engine only. The left engine supplies bleed air to the cockpit and this bleed air flows aft through the pressure vessel to the outflow valves on the aft pressure bulkhead.

R (Right Engine) The R position energizes the left PRSOV closed and energizes the right FCV to 8 ppm. Bleed air supply to the pressure vessel is from the right engine only. Normally the right engine supplies bleed air to the cabin only. However, with the left engine bleed air supply stopped at the PRSOV the right bleed air is supplied to the cabin as well as the cockpit through a one-way check valve. This allows conditioned bleed air to still be supplied to the cockpit.

FOR TRAINING PURPOSES ONLY

FRESH AIR The FRESH AIR position energizes both PRSOVs closed, stopping all bleed air from entering the pressure vessel and starts a fan in the nose of the aircraft. With no bleed air supply, the pressure vessel depressurizes. Once the aircraft is unpressurized, outside fresh air can enter through a fresh air check valve on the forward pressure bulkhead. This fresh air enters between the copilot rudder pedals.

COCKPIT TEMP And CABIN TEMP Knobs The crew uses the COCKPIT TEMP and CABIN TEMP knobs to set the desired temperature in the cockpit and cabin zones. Rotating the knob counterclockwise to COLD selects the coldest possible temperature and rotating the knob clockwise to HOT selects the hottest possible temperature. The normal range of both knobs is 65–85°F.

TEMP knob setting with the cockpit zone temperature sensor and with the cabin bleed air duct temperature sensor. With this information, the cabin PCB makes the necessary adjustment to the right engine temperature control valve.

Throttles Throttles regulate engine power (and hence bleedair temperature and pressure). Bleed-air temperature and pressure can be increased by increasing throttles. Throttles also determine availability of bleed air to the pressure vessel. On the ground with the pressure vessel temperature above 18°C (65°F) and the AIR SOURCE SELECT knob set to BOTH, each throttle closes its respective PRSOV when retarded below 85% TLA. In flight, if a throttle is brought to cutoff, its corresponding PRSOV is shut off and the opposite FCV is switched to 8 ppm if the AIR SOURCE SELECT knob is in the BOTH position.

OPERATION Normally, set the AIR SOURCE SELECT knob to BOTH to ensure proper pressurization inflow and adequate warm air, especially when in flight. Set the COCKPIT TEMP and CABIN TEMP knobs to the desired temperatures. Figure 11-6.   COCKPIT TEMP and CABIN TEMP Knobs

COCKPIT TEMP Knob The COCKPIT TEMP knob sets the desired bleed air temperature entering the pressure vessel from the left engine. The cockpit PCB compares the COCKPIT TEMP knob setting with the cockpit zone temperature sensor and with the cockpit bleed air duct temperature sensor. With this information, the cockpit PCB makes the necessary adjustment to the left engine temperature control valve.

On the ground as determined by either the left or right squat switch, with the AIR SOURCE SELECT knob set to BOTH, bleed-air heat is available only if cabin temperature is below 18°C (65°F) or throttles are above approximately 85% TLA. If the AIR SOURCE SELECT knob is set to L or R, bleed-air inflow is supplied regardless of temperature, TLA, or squat switch position. In flight, bleed-air inflow is always available.

CABIN TEMP Knob The CABIN TEMP knob sets the desired bleed air temperature entering the pressure vessel from the right engine. The cabin PCB compares the CABIN

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11 AIR CONDITIONING

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VAPOR-CYCLE AIR CONDITIONING

COMPONENTS

DESCRIPTION

The Citation Mustang uses R-134a type refrigerant.

The conventional vapor-cycle system provides cool, dry air for the cockpit and passenger cabin. When the system is selected by the crew, the vaporcycle evaporators in the pressure vessel extract moisture and cool the air that is already in the cockpit and passenger cabin. The vapor-cycle system functions in conjunction with the temperature controlled bleed air.

Compressor

A pressurized refrigerant gas circulates through the system to absorb heat from the pressure vessel and dissipate it into the atmosphere. A DC-powered compressor compresses the refrigerant gas, heating it, then pumps it into a condenser, which transfers heat from the refrigerant to cooler ambient air passing over the condenser coils. The cooled refrigerant condenses into a pressurized liquid. Evaporator units depressurize the liquid refrigerant into a spray, cooling the liquid during the process. The cold spray circulates inside the evaporator coils. A fan in each evaporator unit blows interior air (cockpit or cabin air) across the cold coils, cooling and dehumidifying the air (Figure 11-7). As the interior air transfers its heat through the heat exchanger, the refrigerant in the coils absorbs the heat, warms and vaporizes, then returns it to the compressor, carrying heat away from the aircraft interior. The refrigerant circulates continuously throughout the system, transferring heat from the interior air to the ambient air. The forward evaporator is adjacent to the copilot seat. It provides conditioned air to the crew through two vents outboard of the instrument panel. The aft evaporator is on the aft bulkhead and provides cooling air to the passengers through individual overheat vents.

11-8

Refrigerant

The DC-powered compressor in the tail cone area compresses warm, low-pressure refrigerant vapor from the evaporators into a hot, high-pressure gas, then pumps it through the condenser. On the ground as determined by either the left or right squat switch, the compressor can operate from either the left or right generator or an external power unit (EPU). When airborne, it only operates when BOTH generators are operating. If a generator fails while airborne, the compressor is automatically load-shed (disconnected from DC power to reduce electrical loads). The compressor operation hours are not the same as aircraft hours. A separate hour-meter for the compressor is on the compressor assembly, above the battery in the tail cone. To read the hour-meter, access it through a panel directly opposite the baggage door on the right forward side of the baggage compartment.

Condenser The condenser is in the tail cone and cools the hot, high-pressure refrigerant gas flowing from the compressor prior to entry into the pressure vessel. It utilizes finned coils through which the refrigerant gases flow to derive a cooling benefit from transient, ambient air, thereby acting as a heat exchanger. This ambient, atmospheric air is ducted through the condenser by an inlet on the right side of the aft tail cone. The cooling air is then routed over the condenser coils, allowing for the transfer of heat, and then is ducted overboard through a duct on the upper right side of the tail cone.

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SERVICE CONNECTION

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HEAT EXCHANGER

RECEIVER/ DRYER

P

DRAIN

HEAT EXCHANGER

PT T

DISCHARGE LINE (HIGH PRESSURE)

AIR-CONDITIONING EXHAUST

LOAD SHEDDING FUNCTION

Figure 11-7.   Vapor-Cycle Schematic

AIR-CONDITIONING INLET

CONDENSER ASSEMBLY

100 AMP CURRENT LIMITER

PRESSURE RELIEF VALVE

SERVICE CONNECTION

EVAPORATOR BLOWER FROM AFT CABIN

SUCTION LINE (LOW PRESSURE)

TEMPERATURE SENSOR

TEMPERATURE SENSOR

FROM FWD CABIN

EVAPORATOR BLOWER

COMPRESSOR

HOUR METER

COMPRESSOR ASSEMBLY

H

DRAIN

COPILOT OUTLET

T

AFT EVAPORATOR

HEAT EXCHANGER

CABIN VENTS

WINDSHIELD OUTLETS

PT

CONDENSER BLOWER

EXPANSION VALVE

AFT CABIN OUTLET

PILOT OUTLET

EXPANSION VALVE

LIQUID LINE (HIGH PRESSURE)

FWD EVAPORATOR

11 AIR CONDITIONING

AIR-CONDITIONING INLET

COMPRESSOR DISCHARGE

AIR-CONDITIONING EXHAUST

HIGH-PRESSURE LIQUID

TO OUTLETS

SUCTION

FROM FWD CABIN

LEGEND

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Evaporators There are two evaporators in the vapor-cycle system. Cold refrigerant chills the evaporator coils. Electrically powered blowers push cockpit or cabin air over the cold evaporator coils and then force the cold air into the cockpit or cabin air distribution system. Water vapor from the cockpit or cabin air condenses on the evaporator coils and the liquid water is routed overboard through two heated vents under the fuselage (one near each evaporator) (Figure 11-8). This dehumidifies cabin air.

COCKPIT FAN and CABIN FAN Knobs Switches electrically energize the fans at the forward (cockpit) and aft (cabin) evaporators, to push interior air over the cold evaporator coils and into the interior through ducts: • The LOW, HIGH, and FLOOD positions direct air through individually adjustable cockpit and cabin outlets. • The OFF position deenergizes the fan.

Figure 11-10.   COCKPIT FAN and CABIN FAN Knobs Figure 11-8.   Evaporator Drains (Aft)

CONTROLS AND INDICATIONS Controls specifically for the vapor-cycle air conditioning are on the ENVIRONMENTAL control panel.

AIR COND Switch The AIR COND switch activates the vapor cycle system when selected on, provided either or both fan switches are also selected to an operationally energized position (i.e., LOW, HI, or FLOOD).

OPERATION To operate the vapor-cycle air conditioning system, select the AIR COND switch to AIR COND and select either (or both) fan switches to any position other than OFF.

FRESH AIR AND FANS DESCRIPTION A separate fresh-air system is in the cockpit between the copilot rudder pedals. If FRESH AIR is selected with the AIR SOURCE select knob, the duct (Figure 11-8) routes fresh air to the cockpit and an electric fan blows the air through the duct into the cockpit, between the copilot rudder pedals. When the aircraft is pressurized, a check valve in the fresh-air duct closes and fresh air does not enter the cockpit. This prevents pressurized bleed air inside the pressure vessel from leaking out through the fresh-air duct.

Figure 11-9.   AIR COND Switch

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Additional circulation of air in the pressure vessel is provided by the evaporator fans. They route cockpit and cabin air through the evaporators and then through ducts to cockpit and cabin outlets. The fans can be directly commanded by the crew when the vapor-cycle system is not operating.

Table 11-1.   CAS MESSAGES DUCT O’HEAT L-R DESCRIPTION

CAUTION The FRESH AIR position on the AIR SOURCE SELECT knob shuts off pressurized bleed-air inflow. The pressure vessel will depressurize at nominal leak rate.

INHIBITS

EMER, LOPI

PRESS OFF DESCRIPTION

CONTROLS AND INDICATIONS AIR SOURCE SELECT Knob

This message is displayed when either the cabin or cockpit air supply duct temperature exceeds approximately 300°F (149°C). Crew action is required. This message disappears if the temperature falls below approximately 285°F (141°C).

INHIBITS

This message is displayed when the air selector knob is in the OFF or FRESH AIR position. EMER, LOPI

The AIR SOURCE SELECT knob, when set to FRESH AIR, energizes the fresh-air fan. If the AIR SOURCE SELECT knob is at any other position, the fresh-air fan is deenergized (refer to Temperature-Controlled Bleed-Air Inflow earlier in this chapter).

OPERATION When on the ground or in unpressurized flight, if fresh-air ventilation is desired, set the AIR SOURCE SELECT knob to FRESH AIR. At any time that increased circulation of air in the cabin is desired, set either the COCKPIT FAN or CABIN FAN fan switch (or both switches) to any setting other than OFF.

LIMITATIONS For specific limitations, refer to the FAA-approved AFM.

EMERGENCY/ ABNORMAL For specific information on emergency/abnormal procedures, refer to the appropriate checklist or FAA-approved AFM.

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QUESTIONS 1. If both generators fail and the BATT switch is positioned in the EMER position, what is the failed position for the PRSOVs and FCVs? A. Both PRSOVs fail open with a loss of normal DC power and the FCVs fail to 4 ppm B. Both PRSOVs fail closed with a loss of normal DC power and the FCVs fail to 4 ppm C. Both PRSOVs fail open with a loss of normal DC power and the FCVs fail to 8 ppm D. Both PRSOVs fail closed with a loss of normal DC power and the FCVs fail to 8 ppm 2. Which of the following is required for the vapor cycle air conditioner to operate while in fight? A. Both generators operating, the AIR COND switch placed in the up (AIR COND) position, and either the COCKPIT or CABIN FAN knobs out of the OFF position. B. Both generators operating, the AIR COND switch placed in the up (AIR COND) position, and the COCKPIT and CABIN TEMP knobs set to the COLD position. C. Either the left or right generator operating, the AIR COND switch placed in the up (AIR COND) position, and either the COCKPIT or CABIN FAN knobs out of the OFF position. D. Both generators operating, the AIR COND switch placed in the up (AIR COND) position. There is no need to move the COCKPIT or CABIN FAN knobs out of the OFF position.

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3. Which of the following conditions allows bleed air inflow to the pressure vessel to be stopped? A. Positioning the AIR SOURCE SELECT knob to either the L or R position B. Positioning the AIR SOURCE SELECT knob to FRESH AIR C. On the ground with the AIR SOURCE SELECT knob positioned to L or R and a pressure vessel temperature above 18°C (65°F) D. In flight with the AIR SOURCE SELECT knob positioned in the BOTH position and a pressure vessel temperature above 18°C (65°F) 4. Selecting the FRESH AIR position: A. Activates the vapor cycle air conditioner in flight to cool the bleed air from the engine B. Causes the cabin to depressurize in flight at a nominal rate C. Must not be used during ground operations D. Does not affect normal pressurized flight 5. The right engine supplies bleed air to what part of the pressure vessel? A. Under normal operating conditions the right engine supplies bleed air to the cockpit. However, if the left engine bleed air supply is stopped the right engine can also supply bleed air to the cabin through a one-way check valve. B. Under normal operating conditions the right engine supplies bleed air to the cabin. However, if the left engine bleed air supply is stopped the right engine can also supply bleed air to the cockpit through a one way check valve. C. The right engine only supplies bleed air to the cockpit. D. The right engine only supplies bleed air to the cabin.

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6. On the ground, the vapor cycle air conditioning system can be operated by: A. Either the left or right generator as well as a ground power unit. B. The left generator or a ground power unit C. The right generator or a ground power unit D. Ground power unit only 7. In the event of a total DC power failure, the flow control valves fail to: A. 8 ppm B. 4 ppm C. 10 ppm D. 20 ppm 8. In the event of a total DC power failure, the temperature control valves fail to: A. 8 ppm B. 4 ppm C. Max hot position D. Max cold position 9. If either engine fails, the opposite flow control valve automatically goes to: A. 4 ppm B. 8 ppm C. 10 ppm D. 12 ppm 10. On the ground with the throttles below 85% TLA and a pressureabove 18°C (65°F), bleed air inflow is: A. Stopped by the PRSOVs B. Selected to 8 ppm C. Selected to 4 ppm D. Stopped by the flow control valves 11. The hot bleed air from the engines is cooled by: A. Either the Left Generator or the Right Generator needs to be operating B. Heat exchangers utilizing gas refrigerants C. Centrifugal diffusers that drop the air pressure and thereby cools the bleed air D. Heat exchangers utilizing liquid refrigerants Revision 1.0

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INTENTIONALLY LEFT BLANK

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CHAPTER 12 PRESSURIZATION CONTENTS GENERAL..............................................................................................................................12-1 DESCRIPTION..................................................................................................................... 12-3 COMPONENTS................................................................................................................... 12-3 Outflow Valves............................................................................................................... 12-3 CONTROLS AND INDICATIONS...................................................................................... 12-5 Pressurization Controller............................................................................................... 12-5 Destination Elevation Setting........................................................................................ 12-6 PRESS CONT (STANDBY—NORM) Switch............................................................. 12-6 Manual CABIN DUMP Switch..................................................................................... 12-6 Pressurization Display................................................................................................... 12-6 OPERATION........................................................................................................................ 12-8 Ground/Flight Modes.................................................................................................... 12-8 Setting Destination Elevation..................................................................................... 12-10 High-Altitude Airport Operation (Autoschedule)...................................................... 12-10 Manual CABIN PRESSURE DUMP......................................................................... 12-11 LIMITATIONS................................................................................................................... 12-11 EMERGENCY/ABNORMAL........................................................................................... 12-11 QUESTIONS..................................................................................................................... 12-13

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12 PRESSURIZATION

INTRODUCTION................................................................................................................ 12-1

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ILLUSTRATIONS 12-1.

Pressurization System Schematic.......................................................................... 12-2

12-2.

Outflow Valve Positions......................................................................................... 12-4

12-3.

Pressurization Controls.......................................................................................... 12-5

12-4.

Standard (MFD) Pressurization Display................................................................ 12-7

12-5.

Pressurization Display (Reversionary Mode)........................................................ 12-7

12-6.

Cabin Pressure Display....................................................................................... 12-10

TABLES Table Title Page 12-1.

CAS MESSAGES............................................................................................... 12-12

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Figure Title Page

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INTENTIONALLY LEFT BLANK

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12 PRESSURIZATION

CHAPTER 12 PRESSURIZATION

INTRODUCTION The pressurization system on the Citation Mustang maintains cabin altitude lower than actual aircraft altitude to provide a suitable environment for the crew and passengers. The cabin altitude is maintained by regulating the bleed air escaping overboard from the pressure vessel. The system consists of a pressurization controller, two outflow valves, safety valves, pilot controls, and system monitoring.

GENERAL Two elements provide cabin pressurization. One is a constant source of temperature controlled bleed air to the cabin (refer to Chapter 11—“Air Conditioning”) (Figure 12-1). The other is a method of controlling the outflow of the bleed air from the cabin. This control of bleed air inflow and outflow results in a cabin differential pressure (difference between cabin pressure and outside air pressure). This difference in pressure equates to the cabin

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pressure altitude being lower than the aircraft cruising pressure altitude. The maximum cabin pressure differential is 8.5 + 0.1 psid. Normal cabin pressure differential is 8.3 psid. This permits a normal cruise cabin altitude at the aircraft top altitude (FL 410) of 7,800 ± 200 feet.

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STATIC SOURCE

12 PRESSURIZATION

MAX P LIMITER

CABIN PRESS

GRILLE

CABIN EXHAUST

CABIN PRESS MAX ALT LIMITER CABIN PRESS

THROTTLE SWITCH

10,00 feet and is not in high altitude mode. This message will also be displayed if the aircraft is >15,000 feet regardless of what mode the aircraft is in.

DESCRIPTION

LOPI, TOPI, ON GROUND INHIBITS

This message indicates failure of the ARINC 429 data link from the G1000 system, indicating that the controller may no longer have valid data on outside air pressure, actual aircraft altitude, or selected destination elevation. EMER, LOPI, TOPI

CABIN ALT DESCRIPTION

INHIBITS

In high altitude mode, The amber CABIN ALT message displays if the cabin altitude exceeds 10,000–15,000 feet for more than 30 minutes.

PRESS OFF DESCRIPTION

INHIBITS

LOPI, TOPI, ON GROUND

This message is displayed when the air selector knob is in the OFF or FRESH AIR position. EMER, LOPI

PRESS CTRL DESCRIPTION

INHIBITS

This message indicates failure of the pressurization controller or that the pilot has selected the PRESS CONT switch to STANDBY, disabling the pressurization controller. EMER, LOPI, TOPI

PRESS OFF DESCRIPTION

INHIBITS

In high altitude mode, The amber CABIN ALT message displays if the cabin altitude exceeds 10,000–15,000 feet for more than 30 minutes. LOPI, TOPI, ON GROUND

CABIN ALT DESCRIPTION

INHIBITS

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In high altitude mode, The amber CABIN ALT message displays if the cabin altitude exceeds 10,000–15,000 feet for less than 30 minutes. LOPI, TOPI, ON GROUND

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1. With a destination field elevation below 8,000 feet, the CABIN ALT CAS message displays when the cabin altitude exceeds: A. 10,000 feet B. 8,500 feet C. 14,000 feet D. 12,000 feet

6. The maximum altitude limit valves are only effective as long as: A. The digital auto controller is functioning B. There is bleed inflow into the pressure vessel C. AC power is available D. DC power is available

2. With the PRESS CONT switch in the STANDBY position, the CAS displays a(n): A. PRESS CTRL CAS message B. WOW MISCOMPARE CAS message C. FDR FAIL CAS message D. ENG CTRL R CAS message

7. If the CABIN DUMP switch is pressed while the aircraft is at 30,000 feet, and there is bleed air being supplied to the pressure vessel, the cabin altitude will climb or descend to what altitude? A. 30,000 feet because the outflow valves are fully opened B. Sea level if the dump switch is held in position C. Remains at 8,000-foot cabin altitude because full cabin pressure dumping does not occur at that altitude D. Stays below 15,000 feet because the pneumatic maximum cabin altitude limit valve overrides the CABIN DUMP

3. Placing the PRESS CONT switch in the STANDBY position causes the pressurization controller to: A. Switch to high altitude mode when the destination field is above 8,000 feet B. Deenergizes and unpressurizes the pressure vessel C. Deenergizes, failing the climb and dive solenoid closed D. Deenergizes, failing the climb and dive solenoid open 4. The max deferential pressure is: A. 7.5 + 0.1 B. 8.0 + 0.1 C. 8.7 + 0.1 D. 8.5 + 0.1 5. To pressurize the aircraft and operate normally, the AIR SOURCE SELECT knob must be positioned to: A. L, R, or BOTH B. STANDBY C. DUMP D. FRESH AIR

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8. If total DC power is lost, automatic pressurization without pilot input is controlled by: A. The CABIN DUMP switch B. The pressurization controller C. Maximum delta-P and maximum cabin altitude limit valves D. The COCKPIT TEMP knob 9. If both generators fail and the BATT switch is placed in the EMER position, the pilot control of the pressurization system is with the use of the: A. AIR SOURCE SELECT knob B. CABIN DUMP switch C. Flow control valves D. Pressure regulating shutoff valves (PRSOVs)

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QUESTIONS

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12 PRESSURIZATION

10. The pressurization system operates in high altitude mode if: A. There is a destination field above 10,000 feet set and the aircraft has descended below 24,500 feet B. There is a destination field above 12,000 feet set and the aircraft has descended below 25,000 feet C. There is a destination field above 8,000 feet set and the aircraft has descended below 24,500 feet D. There is a destination field above 5,000 feet set and the aircraft has descended below 24,500 feet

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CHAPTER 13 HYDRAULIC POWER SYSTEM CONTENTS INTRODUCTION................................................................................................................ 13-1 GENERAL..............................................................................................................................13-1 DESCRIPTION..................................................................................................................... 13-3 COMPONENTS................................................................................................................... 13-3 Reservoir....................................................................................................................... 13-3

Hydraulic Accumulator................................................................................................. 13-3 Pressure Switches.......................................................................................................... 13-3 System Relief Valve...................................................................................................... 13-5 Filters............................................................................................................................. 13-5 CONTROLS AND INDICATIONS...................................................................................... 13-5 Hydraulic Reservoir Sight Gauge................................................................................. 13-5 Hydraulic Accumulator Pressure Gauge....................................................................... 13-5 Manual Accumulator Bleed Valve................................................................................. 13-5 OPERATION........................................................................................................................ 13-6 Preflight......................................................................................................................... 13-6 In Flight......................................................................................................................... 13-6 LIMITATIONS...................................................................................................................... 13-6 EMERGENCY/ABNORMAL.............................................................................................. 13-6 QUESTIONS........................................................................................................................ 13-7

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13 HYDRAULIC POWER SYSTEM

Pump.............................................................................................................................. 13-3

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ILLUSTRATIONS Figure Title Page 13-1.

Hydraulic System Schematic................................................................................. 13-2

13-2.

Hydraulic Reservoir............................................................................................... 13-4

13-3.

Hydraulic Reservoir Sight..................................................................................... 13-4

13-4.

Hydraulic Accumulator Sight Gauge..................................................................... 13-5

13-5.

Manual Accumulator Bleed Valve......................................................................... 13-5

TABLES Table Title Page CAS MESSAGES.................................................................................................. 13-6 13 HYDRAULIC POWER SYSTEM

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13 HYDRAULIC POWER SYSTEM

CHAPTER 13 HYDRAULIC POWER SYSTEM

INTRODUCTION This chapter provides information on the hydraulic system in the Citation Mustang aircraft. Hydraulic fluid driven by a pump and regulated by valves provides pressure for landing gear and brakes. Operation of these devices is presented in Chapter 14—“Landing Gear and Brakes.” This chapter describes the portions of the hydraulic system used by both subsystems.

GENERAL The hydraulic system permits the application of substantial force by converting a volume of fluid flow into pressure on a hydraulic piston. Hydraulic fluid lines provide the capability to transmit that force wherever it is required in the aircraft without heavy or complex mechanical linkages. In the Mustang, a reservoir stores hydraulic fluid for the pump and receives return flow from the system (Figure 13-1). One electrically driven pump supplies hydraulic power. Hydraulic fluid is rout-

ed through lines, regulated by system valves, and cleaned by filters. The hydraulic system responds automatically to the activation of controls for the landing gear and brakes. The engine indicating and crew alerting system (EICAS) indicates system status. Also, landing gear and brakes each have a pneumatic backup system to provide pressure to their respective systems in the event of hydraulic system failure (refer to Chapter 14—“Landing Gear and Brakes”).

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13-1

13-2

RESERVOIR DRAIN

FOR TRAINING PURPOSES ONLY LH BUS

FILTER BYPASS VALVE (100 PSID)

RESERVOIR DRAIN VALVE

PUMP

OVERBOARD VENT LINE

BRAKE SYSTEM TO GEAR RETRACT

TO GEAR EXTEND

CABIN AIR

RETURN

CONTROL MANIFOLD

SYSTEM PRESSURE

NITROGEN

LEGEND

ACCUMULATOR

Figure 13-1.   Hydraulic System Schematic

FILTER

FILTER

FILTER BYPASS VALVE (100 PSID)

MANUAL ACCUMULATOR BLEED VALVE

RESTRICTOR VALVE

RESERVOIR SIGHT GAUGES

RESERVOIR

13 HYDRAULIC POWER SYSTEM

CABIN AIR

LOW PRESSURE SWITCH (750–950 PSIG) (ACTIVATES CAS MESSAGE)

GEAR RETRACT PRESSURE SWITCH (1,275–1,475 PSIG)

NORMAL PRESSURE SWITCH (1,100–1,475 PSIG)

PRESSURE RELIEF VALVE (OPENS AT APPROX 1,750 PSIG)

P

P

P

TO LANDING GEAR BLOWDOWN SYSTEM

EMERGENCY BLOWDOWN BOTTLE

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DESCRIPTION

COMPONENTS

The hydraulic system is classified as a “closed center” system. When no subsystems are in use (landing gear or brakes), there is normally no flow in the system, except to maintain accumulator charge.

The majority of the hydraulic system components are in the lower left nose compartment outside the pressure vessel.

The LANDING GEAR selector and gear-position sensor switches determine which pressure switch regulates pump operation and resultant system pressure. When the combination of LANDING GEAR selector, weight on wheels, and gear position sensor switches indicate the gear is ready for retraction or is in transit, a gear-retraction pressure switch is automatically selected to regulate the pump to maintain approximately 1,300 psig minimum system pressure. At all other times when the hydraulic system is energized, a normal pressure switch is automatically selected to regulate pressure, maintaining 1,050 psig minimum. (For details about the LANDING GEAR selector and gear-position sensor switches, refer to Chapter 14—“Landing Gear and Brakes”). Both pressure switches limit maximum pressure to approximately 1,550 psig. The system relief valve provides additional protection if system pressure exceeds 1,750 psig. A hydraulic reservoir stores hydraulic fluid for the pump and receives return flow from the system. During system operation, an accumulator stores hydraulic fluid under nitrogen pressure to maintain system pressure. Two filters clean the system: one in the pressure flow upstream of the manifold and downstream of the pump, and one in the return flow, upstream of the reservoir. The system uses red MIL-PRF-87257 hydraulic fluid and is designed for operation in ambient temperatures ranging from –65°F to 160°F (–54°C to 71°C).

RESERVOIR The hydraulic reservoir is attached to the left forward pressure bulkhead (Figures 13-2 and 13-3) and stores hydraulic fluid for the pump and receives return flow from the system. When fluid flows, the excess hydraulic fluid returns to the reservoir. Cabin air pressurizes the reservoir to reduce foaming and assure positive flow to the pump. A relief valve opens at approximately 10 psi to prevent overpressurization. There are two sight gauges on the hydraulic reservoir that are visible through the forward baggage compartment liner. The reservoir capacity is approximately 2.2 quarts (2 liters) as indicated by the full mark.

PUMP The hydraulic pump is powered by normal DC power through the HYD PUMP circuit breaker. The pump operates whenever the pressure drops below the lower set point of the normal gear retract pressure switch. Pump operation discontinues once pressure is restored to the upper set point of the regulating pressure switch.

HYDRAULIC ACCUMULATOR The hydraulic accumulator receives fluid from the pump and stores a supply of hydraulic fluid under pressure. During pump inactivity this pressurized fluid maintains pressure against normal internal leakage within the system. It is also used to supplement pump flow during landing gear extension and retraction.

PRESSURE SWITCHES There are a total of three pressure switches in the control manifold. Two of these switches control pressure to the overall system and gear retraction cycle; the third pressure switch provides a warning

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13 HYDRAULIC POWER SYSTEM

The Mustang uses hydraulic power for retraction and extension of the landing gear, and operation of the brakes. The system includes a single electrically driven hydraulic pump, which functions to maintain and supplement accumulator pressure.

CITATION MUSTANG PILOT TRAINING MANUAL

for low system pressure. The switches are identified as NORMAL, GEAR RETRACTION and LOW PRESSURE. The NORMAL pressure switch maintains a constant pressure on the system under no load. The GEAR RETRACTION pressure switch provides increased system pressure for the gear retraction cycle. The LOW PRESSURE switch alerts the crew to an overall low system pressure through the HYD PRESS LO message in the CAS window.

NORMAL Pressure Switch Whenever the GEAR RETRACTION pressure switch is not active, the NORMAL pressure switch regulates pump operation. It closes to energize the pump whenever system pressure falls below approximately 1,050 psig and opens to deenergize the pump when system pressure rises above approximately 1,550 psig. 13 HYDRAULIC POWER SYSTEM

GEAR RETRACTION Pressure Switch Figure 13-2.   Hydraulic Reservoir

When the LANDING GEAR selector is selected UP and the landing gear is not fully retracted (as indicated by uplock sensor switches) or when the LANDING GEAR selector is selected GEAR DOWN and the aircraft is airborne (as indicated by the squat switches), the GEAR RETRACTION pressure switch regulates pump operation. It closes to energize the pump whenever system pressure falls below approximately 1,300 psig and opens to deenergize the pump when system pressure rises above approximately 1,550 psig.

LOW PRESSURE Switch The LOW PRESSURE switch (750 psig–1,000 psig) signals the HYD PRESS LO message in the CAS window to display. This message advises the crew that the hydraulic pressure is low and that the hydraulic pump failed to operate automatically when signaled by the NORMAL or GEAR RETRACTION pressure switches. Figure 13-3.   Hydraulic Reservoir Sight Gauge

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SYSTEM RELIEF VALVE A mechanical relief valve in parallel with the landing gear and brakes maintains the system pressure at a maximum of 1,750 psig.

FILTERS The system incorporates two filters: one for filtering fluid leaving the pump and one for filtering return fluid prior to entering the reservoir. Each filter incorporates a bypass valve that opens at 100 ± 10 psid if the filter element clogs. As a bypassed filter is a standard maintenance interval item, there is no cockpit indication or filter indication of bypassing.

Figure 13-4.   Hydraulic Accumulator Sight Gauge

13 HYDRAULIC POWER SYSTEM

CONTROLS AND INDICATIONS The hydraulic system functions automatically during normal DC power (supplied by generator, battery, or ground power unit). All inflight indications for the hydraulic system appear as crew alerting system (CAS) messages on the EICAS as displayed on the multifunction display (MFD) or (in reversionary mode) on the primary flight displays (PFDs).

HYDRAULIC RESERVOIR SIGHT GAUGE The hydraulic reservoir sight gauge indicates current hydraulic fluid quantity. The gauge is integral to the hydraulic reservoir, and is visible through the aft wall of the nose baggage compartment (see Figure 13-3).

HYDRAULIC ACCUMULATOR PRESSURE GAUGE

Figure 13-5.   Manual Accumulator Bleed Valve

MANUAL ACCUMULATOR BLEED VALVE The manually actuated accumulator bleed valve is in the left forward baggage compartment next to the accumulator pressure gauge. It is accessed through a hinged access panel (Figure 13-5). On the ground, it enables the crew to release pressurized fluid from the accumulator to the reservoir. This is done during preflight inspection if the hydraulic system was previously energized to ensure accurate reservoir fluid level and accumulator precharge readings.

A pressure gauge on the accumulator (Figure 13-4) indicates current pressure in the accumulator. The gauge is on the aft wall of the nose baggage compartment, and indicates 0–2,000 psig.

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OPERATION

Table 13-1.   CAS MESSAGES HYD PRESS LO

PREFLIGHT

DESCRIPTION

Bleed the accumulator prior to checking the accumulator precharge and fluid level in the reservoir. Verify that the accumulator is precharged per the placard and that the hydraulic fluid level is adequate (no air visible in lower sight gauge).

IN FLIGHT Hydraulic System

INHIBITS

The hydraulic system operates automatically to maintain pressure and sends cautionary CAS messages to the crew if there is a fault.

EMER TOPI

HYD PUMP ON DESCRIPTION

13 HYDRAULIC POWER SYSTEM

Hydraulic Subsystems Hydraulically powered subsystems include landing gear and brakes. Application of hydraulic power to the two subsystems is presented in Chapter 14, “Landing Gear and Brakes.”

The low pressure switch in the hydraulic control manifold controls the amber HYD PRESS LO message. As hydraulic system pressure decreases below 750 psig, the HYD PRESS LO message appears, accompanied by MASTER CAUTION lights. As the pump increases system pressure to greater than 1,000 psig, a circuit opens to extinguish the message and the MASTER CAUTION lights.

INHIBITS

The amber HYD PUMP ON message indicates that the hydraulic pump has been operating continuously for over 60 seconds. Refer to the checklist. EMER, LOPI

LIMITATIONS For specific limitations, refer to the FAA-approved AFM.

EMERGENCY/ ABNORMAL For specific information on emergency/abnormal procedures, refer to the appropriate checklist or FAA-approved AFM.

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1. The hydraulic system: A. Uses only red MIL-PRF-87257 fluid B. Has a reservoir pressurized by pylon scoop ram air C. Uses two electrically powered pumps D. Only functions when both pumps are operating

6. The HYD PUMP ON message in the CAS window appears when: A. The pump has been operating for longer than 60 seconds B. The system is operating normally C. The speedbrakes are extended D. The flaps are extending

2. If normal DC power is lost: A. The hydraulic system is powered by the standby battery pack B. The hydraulic system is inoperative C. The hydraulic system is powered by the emergency bus D. Both B and C

7. On preflight inspection, to check accumulator precharge pressure and hydraulic fluid levels: A. Aircraft power must be on B. System must be discharged with the manual accumulator bleed valve C. Aircraft power must be off D. Both B and C

3. The hydraulic system provides pressure to operate the: A. Landing gear, speedbrakes, and flaps B. Landing gear and wheel brakes only C. Antiskid brakes, landing gear, and flaps D. Speedbrakes, landing gear, and wheel brakes

8. The hydraulic accumulator is used: A. For hydraulic fluid storage tank B. As a means to lower the gear in an abnormal situation C. For storing hydraulic fluid under pressure D. As a backup to normal brakes

4. Access to the hydraulic reservoir sight glass is: A. Through the left forward baggage compartment door B. On the copilot instrument panel C. Ahead of the tail cone baggage compartment D. Inside a door behind the right flap 5. On the ground with the HYD PRESS LO message in the CAS window, the MASTER CAUTION: A. Can be reset regardless of the validity of the CAS message B. Can be reset by powering down the aircraft C. Does not reset as long as this indication remains valid D. Can be reset with the hydraulic pump circuit breaker

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13 HYDRAULIC POWER SYSTEM

QUESTIONS

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CHAPTER 14 LANDING GEAR AND BRAKES CONTENTS INTRODUCTION................................................................................................................ 14-1 GENERAL..............................................................................................................................14-1 LANDING GEAR................................................................................................................ 14-2 Description.................................................................................................................... 14-2 Main Gear System......................................................................................................... 14-3 Nose Gear System......................................................................................................... 14-7 Components................................................................................................................... 14-7 Controls and Indications............................................................................................. 14-10 Operation.................................................................................................................... 14-11 NOSEWHEEL STEERING............................................................................................... 14-13 Description and Operation......................................................................................... 14-13 BRAKES............................................................................................................................ 14-14

Antiskid System......................................................................................................... 14-14 Parking Brakes........................................................................................................... 14-14 Emergency Brakes...................................................................................................... 14-16 Components................................................................................................................ 14-16 Controls and Indications............................................................................................. 14-17 Operation.................................................................................................................... 14-17 LIMITATIONS................................................................................................................... 14-19 EMERGENCY/ABNORMAL........................................................................................... 14-19 QUESTIONS..................................................................................................................... 14-21

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Description................................................................................................................. 14-14

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ILLUSTRATIONS Figure Title Page 14-1.

Left Main Landing Gear and Door........................................................................ 14-2

14-2.

Nose Landing Gear and Doors.............................................................................. 14-2

14-3.

Emergency Gear Release Handle (Cover Removed)............................................. 14-3

14-4.

Landing Gear Schematic - Extension.................................................................... 14-4

14-5.

Landing Gear Schematic - Emergency Extension................................................. 14-5

14-6.

Landing Gear Schematic - Retraction................................................................... 14-6

14-7.

Landing Gear Control Panel............................................................................... 14-10

14-8.

Landing Gear Handle Locking Solenoid and Switches...................................... 14-10

14-9.

Landing Gear Position Indications..................................................................... 14-12

14-10. Nosewheel Steering............................................................................................ 14-13 14-11. Stop Bolt Location.............................................................................................. 14-14 14-12. Power Brake and Digital Antiskid System.......................................................... 14-15 14-13. PARKING BRAKE Knob................................................................................... 14-16

14-15. EMERGENCY BRAKE Handle........................................................................ 14-19

TABLES Table Title Page 14-1. CAS MESSAGES............................................................................................... 14-20

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14-14. Emergency Brake System................................................................................... 14-16

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CHAPTER 14 LANDING GEAR AND BRAKES

INTRODUCTION

GENERAL The Citation Mustang has retractable tricycle landing gear that is electrically controlled and hydraulically actuated. Each gear is retracted by its own hydraulic actuator. When retracted, the nose gear and the struts of the main gear are enclosed by mechanically actuated doors connected to the gear struts. The trailing-link main gear wheels remain uncovered in the wheel wells. Gear position and warning are provided by colored indicator lights and an aural warning. In the event of hydraulic gear extension system failure, an independent mechanical uplock release and pneumatic system provide for emergency gear extension.

Nosewheel steering is mechanically actuated through linkage from the rudder pedals. A friction shimmy damper is contained within the nose gear strut. A bungee allows tighter turns with differential power and braking. The aircraft is towed by connections on the nosewheel strut. Power braking (hydraulically actuated) is provided with or without antiskid protection. A “touchdown protection” feature is provided to prevent landing with brakes locked. A spindown feature stops tires from spinning before retracting into the wheel wells. In the event of a hydraulic brake system failure, an independent pneumatic system provides for emergency braking. A parking brake system is available to temporarily lock the brakes on the ground.

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14 LANDING GEAR AND BRAKES

This chapter describes the landing gear, nosewheel steering, and brake system of the Citation Mustang.

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Crew alerting system (CAS) messages report the status of the braking and hydraulic systems and related systems. The rotary test switch tests all indications of the landing gear and brakes.

LANDING GEAR DESCRIPTION The main landing gear struts are trailing-link struts, supporting the wheels with a trunnion and air-oil (oleo) strut, connected by a trailing link (Figure 14-1). The nose landing gear strut is a conventional air-oil (oleo) strut extending from the trunnion (Figure 14-2). Normally, the landing gear is hydraulically actuated, but if the normal gear actuation system fails, the gear can be mechanically and pneumatically released and extended. At airspeeds up to 250 KIAS, the gear can be extended (VLO). The aircraft can be flown with the gear extended at airspeeds up to 250 KIAS (VLE). However, the gear cannot be retracted when the airspeed is above 185 KIAS.

Figure 14-1.   Left Main Landing Gear and Door

It takes 6 seconds to extend the landing gear. At airspeeds between 100 and 160 KIAS, it takes 11–14 seconds to retract the landing gear. At airspeeds between 160 and185 KIAS, retraction takes 18–20 seconds. 14 LANDING GEAR AND BRAKES

Each inboard-retracting main gear uses two hydraulic actuators (one for uplock release and one for gear actuation). Three more hydraulic actuators perform these duties for the forward-retracting nose gear. An electrically positioned gear-control valve directs hydraulic pressure for gear extension or retraction. If hydraulic extension fails, the emergency gear T-handle allows the landing gear to mechanically release and free fall. A bottle of compressed nitrogen, activated by the knob behind the T-handle, provides pressure to ensure gear extension and downlock (Figure 14-3).

Figure 14-2.   Nose Landing Gear and Doors

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Abnormal Main Gear Extension System Two backup gear-extension systems are provided in addition to normal gear extension: • T he T-handle operates a cable system to mechanically release the uplock hooks from the struts. The gear should free-fall into position, aided by the pilot yawing the aircraft.

NOTE

MAIN GEAR SYSTEM Main Gear Extension System Figure 14-4 displays the landing gear actuation during gear extension.

Uplocks The main landing gear struts are mechanically locked in the retracted position by a spring-loaded, hydraulically released uplock hook. In normal operation, to release a strut from its uplock, the gear-control solenoid valve routes fluid to the uplock hook actuator, retracting the piston into the actuator. When the piston retracts completely, it pulls the uplock hook free of the uplock roller, and the gear strut is unlocked to start extension.

Extension When the uplock actuator is fully retracted, fluid passes through the uplock actuator to the gearextend side of the gear actuator. Hydraulic pressure is then applied to the actuators, which extend until the gear is down- and-locked. To speed gear extension and improve free-fall capability of the gear, a regenerative shuttle valve allows fluid to flow from the retract side of the actuators to the extend side.

It may require an acceleration above 150 KIAS to lock the nose landing gear into place. • T he round knob behind the T-handle releases pneumatic pressure (high-pressure nitrogen) from a bottle in the nose compartment to pneumatically operate the uplock actuators and release the uplock hooks, then extend the gear actuators, which extend the gear. Figure 14-5 displays the landing gear actuation during abnormal/emergency extension. Once the uplock hooks are released, the pneumatic pressure is applied to ensure that the gear actuators properly extend and lock the gear. Without using the pneumatic gear-extension system, it is still possible to lock the gear down by yawing the aircraft to force the gear into position. However, mechanical release and downlock should always be followed by the pneumatic extension procedure to ensure complete and proper extension of the gear.

Main Gear Downlock Mechanism Each main gear has a mechanical downlock mechanism (integral to the gear actuator), which locks the main landing gear in the down position. Applying hydraulic pressure is the only way to release the downlocks; therefore, no blocks or external downlock pins are required.

Main Gear Retraction System Figure 14-6 shows landing gear actuation during retraction.

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14 LANDING GEAR AND BRAKES

Figure 14-3.   Emergency Gear Release Handle (Cover Removed)

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LOW-PRESSURE SWITCH

CABIN AIR

NITROGEN

STATIC

RETURN

CONTROL MANIFOLD

RETRACT LINE

PNEUMATIC DUMP VALVE

UPLOCK ACTUATOR ASSEMBLY

NOSEGEAR ACTUATOR

NOSE DOWNLOCK RELEASE ACTUATOR

CABIN AIR

TO BRAKES

MAIN LANDING GEAR ACTUATOR

ACCUMULATOR

FROM BRAKES

HYDRAULIC RESERVOIR

NOSE GEAR UPLOCK ACTUATOR

SHUTTLE VALVE

UPLOCK ACTUATOR ASSEMBLY

ACCUMULATOR BLEED VALVE

GEAR CONTROL SOLENOID VALVE

EXTEND LINE

TO BRAKE METERING VALVE

UPLOCK HOOK ACTUATOR

Figure 14-4.   Landing Gear Schematic - Extension

P

GEAR RETRACT PRESSURE SWITCH

SYSTEM PRESSURE

LEGEND

P

MAIN LANDING GEAR ACTUATOR

REGENERATIVE SHUTTLE VALVE

NORMAL-PRESSURE SWITCH

EMERGENCY BLOWDOWN BOTTLE

14 LANDING GEAR AND BRAKES EMERGENCY GEAR EXTENSION HANDLE

CITATION MUSTANG PILOT TRAINING MANUAL

FOR TRAINING PURPOSES ONLY P

LOW-PRESSURE SWITCH

CABIN AIR

NITROGEN

STATIC

RETURN

CONTROL MANIFOLD

RETRACT LINE

PNEUMATIC DUMP VALVE

UPLOCK ACTUATOR ASSEMBLY

NOSEGEAR ACTUATOR

NOSE DOWNLOCK RELEASE ACTUATOR

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CABIN AIR

TO BRAKES

MAIN LANDING GEAR ACTUATOR

ACCUMULATOR

FROM BRAKES

HYDRAULIC RESERVOIR

NOSE GEAR UPLOCK ACTUATOR

SHUTTLE VALVE

UPLOCK ACTUATOR ASSEMBLY

ACCUMULATOR BLEED VALVE

GEAR CONTROL SOLENOID VALVE

EXTEND LINE

TO BRAKE METERING VALVE

UPLOCK HOOK ACTUATOR

Figure 14-5.   Landing Gear Schematic - Emergency Extension

P

GEAR RETRACT PRESSURE SWITCH

SYSTEM PRESSURE

LEGEND

P

MAIN LANDING GEAR ACTUATOR

REGENERATIVE SHUTTLE VALVE

NORMAL-PRESSURE SWITCH

EMERGENCY BLOWDOWN BOTTLE

EMERGENCY GEAR EXTENSION HANDLE

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LOW-PRESSURE SWITCH

CABIN AIR

NITROGEN

STATIC

RETURN

CONTROL MANIFOLD

RETRACT LINE

PNEUMATIC DUMP VALVE

UPLOCK ACTUATOR ASSEMBLY

NOSEGEAR ACTUATOR

NOSE DOWNLOCK RELEASE ACTUATOR

CABIN AIR

TO BRAKES

MAIN LANDING GEAR ACTUATOR

ACCUMULATOR

FROM BRAKES

HYDRAULIC RESERVOIR

NOSE GEAR UPLOCK ACTUATOR

SHUTTLE VALVE

UPLOCK ACTUATOR ASSEMBLY

ACCUMULATOR BLEED VALVE

GEAR CONTROL SOLENOID VALVE

EXTEND LINE

TO BRAKE METERING VALVE

UPLOCK HOOK ACTUATOR

Figure 14-6.   Landing Gear Schematic - Retraction

P

GEAR RETRACT PRESSURE SWITCH

SYSTEM PRESSURE

LEGEND

P

MAIN LANDING GEAR ACTUATOR

REGENERATIVE SHUTTLE VALVE

NORMAL-PRESSURE SWITCH

EMERGENCY BLOWDOWN BOTTLE

14 LANDING GEAR AND BRAKES EMERGENCY GEAR EXTENSION HANDLE

CITATION MUSTANG PILOT TRAINING MANUAL

CITATION MUSTANG PILOT TRAINING MANUAL

COMPONENTS Main Gear

NOSE GEAR SYSTEM

Each main gear assembly includes: • Trunnion • Trailing link • Oleo strut • Main gear actuators with integral downlocks • Uplock assembly • Main wheel tire and brake assembly • Squat switch

Nose Gear Extension System

Trunnion

Nose gear extension is similar to the main landing gear. In the wheel well, the uplock hook is hydraulically retracted, releasing the uplock roller on the gear. The nosewheel rotates down and aft from the nose wheel well. The nose gear is mechanically locked in the extended position by a spring-actuated downlock. A position switch on the drag brace indicates down-and-locked.

A trunnion is the main support (leg) for each main gear. It connects to the wheel through the oleo strut and the trailing link and is extended or retracted by the main gear actuator. During extension, the trunnion (with the main gear components attached) rotates down-and-outboard on pivots attached to the forward and aft wing spars.

Whenever the nosewheel is extended, nosewheel steering is engaged, regardless of whether the aircraft is in flight or on the ground.

Nose Gear Retraction System During nose gear retraction, a hydraulic actuator releases the nose gear downlock. The nose gear actuator extends, causing the nose gear to retract forward into the nosewheel well. On takeoff, with weight off wheels, the nosewheel steering remains engaged until retraction. During retraction, nosegear steering is disengaged and the nose gear is mechanically centered. In the wheel well, a spring-loaded mechanical uplock hook catches the uplock roller on the gear when it retracts. A position-sensor switch in the uplock indicates up-and-locked.

Nose Gear Door System Nose gear movement actuates two doors to completely enclose the nose gear and wheel at retraction. The doors open during gear extension and remain open after the gear is extended.

Trailing Link The trailing link connects the trunnion to the wheel through a pivot and an oleo strut. It allows the wheel to simultaneously move up and aft when landing or during ground operations.

Oleo Strut The oleo (air-oil) strut is a sealed hydraulic piston and cylinder that uses compressed nitrogen to absorb landing and taxiing shocks. It absorbs shocks between the trailing link (attached to the wheel) and the trunnion (attached to the aircraft). On the ground, the oleo struts support the weight of the aircraft.

Main Gear Actuators Inboard of (and attached to) each main gear trunnion is a fluid-driven actuator, which extends or retracts the main landing gear. The main gear actuators are normally driven hydraulically but can be extended pneumatically for emergency gear extension.

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14 LANDING GEAR AND BRAKES

The main gear retracts when hydraulic pressure is applied to the retract side of the actuators. This first releases the downlocks, then forces the actuators to retract, pulling the gear into the wheel wells. Before the wheels enter the wheel wells, an automatic braking feature (spindown) stops the wheels from spinning to prevent loose wheel tread or debris from striking the interior of the wheel well. When the gear is fully retracted into the uplocks, a switch in each uplock detects that the main gear is up-and-locked.

CITATION MUSTANG PILOT TRAINING MANUAL

Each main landing gear actuator includes its own integral mechanical locking system to lock the actuator in place when it is fully extended, thereby locking the gear down. Hydraulic retraction pressure retracts the locking system and permits gear retraction.

Uplock Assembly In each wheel well, a spring-loaded mechanical uplock hook catches the uplock roller on the gear when it retracts. (During preflight, check that the rollers rotate.) This locks the gear in the up position. A switch in the uplock assembly detects when the trunnion uplock roller is in the lock (gear is up-and-locked). At the start of gear extension, a hydraulic uplock-sequencing valve/actuator unit retracts the uplock hook, releasing the main landing gear, then passes hydraulic fluid to the gear actuator. In case of an emergency gear extension, pressurized nitrogen gas retracts the uplock hook.

Main Wheel, Tire, and Wheel Assembly Each main gear assembly includes a single wheel with tire and a fluid-actuated multiple-disc brake assembly. Each main wheel has three fusible plugs that melt to deflate the tire if excessive temperature is generated by an overheated brake. Inflate with dry nitrogen to 85 ± 5 psi (586 KPa ± 34 KPa) unloaded. Maximum tire ground speed is 160 knots. 14 LANDING GEAR AND BRAKES

Squat Switch There is one squat switch on each left and right main landing gear. The switch indicates when weight is on that wheel. The squat switches are positioned in the landing gear assembly, and detect when wheel position changes up or down, as caused by weight-on-wheels or weight-off-wheels. Each squat switch is connected to several aircraft systems. Some systems are connected to both squat switches. Some systems function differently depending on whether the weight is on one wheel or on both wheels. Some systems are sensitive to whether only a specific wheel (left or right) has weight on it.

14-8

If the squat switches are not in the same position (weight-on-wheels or weight-off-wheels) for more than 2 seconds, the amber CAS message WOW MISCOMPARE warns of the difference. Malfunctions of the squat switches and their associated circuits may cause abnormal functioning of any or all of the aircraft systems that use squat switch information.

Dependent Systems The following systems require squat switch information for normal functioning: • External doors • Engine/FADEC • Pneumatics • Windshield anti-ice • Air conditioning • Pressurization • Landing gear • Brakes (antiskid) • Avionics • Stall warning

L and R SQUAT SWITCH Circuit Breakers Each squat switch (left and right) is powered through the corresponding L or R SQUAT SWITCH circuit breaker in the ENGINE SYSTEMS section of the corresponding CB panel (left or right).

Main Gear Door and Fairing Each main landing gear strut has a gear door mechanically attached to the trunnion assembly of the gear. When the gear is operated, the door moves up and down with the gear itself and is not separate from the gear in its operation. The door will cover the gear strut but the tire is partially exposed.

Nose Gear The nose gear assembly supports the nose section of the aircraft while on the ground and provides steering and a linkage for towing.

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Strut The nose gear strut includes the trunnion (attached to pivots and moved by the gear actuator), the shock strut, and the nosewheel fork. The integral oleo (airoil) shock strut absorbs landing impact and other shocks. The lower end of the shock strut attaches to the nosewheel fork, which holds the nosewheel assembly. A folding torque link holds the shock strut cylinder and nosewheel fork piston together and keeps them aligned with each other.

Shimmy Damper To reduce nose gear shimmy during takeoff, landing, and taxiing, the nose gear has a shimmy damper. It is a friction band around the center of the shock strut cylinder. It rubs against the inside of the trunnion using friction to reduce nosewheel shimmy.

Drag Brace and Downlock In the down (gear-extended) position, the gear is mechanically locked down by an integral locking mechanism in the drag brace. A position switch on the drag brace signals when the nose gear is down-and-locked. During extension, the downlock is spring-actuated to lock mechanically. During retraction, the downlock is released by a hydraulic actuator on the drag brace.

Uplock Assembly A mechanical latching system (uplock hook) similar to the main gear system is attached to the airframe in the nose wheel well. It locks the nose gear in the up (gear-retracted) position. During gear extension, a fluid actuator/valve unit releases the hook and then passes hydraulic fluid (or pressurized nitrogen in an emergency) to the gear actuator.

Revision 1.0

On gear retraction, the uplock latch catches a roller on the rising nosewheel fork to lock the gear in the up position. (During preflight, check that the roller rotates.) A switch in the uplock hook mechanism detects whether or not the gear is up-and-locked.

Gear Actuator The fluid-driven nose gear actuator retracts to extend the nose gear. It also triggers nose gear door operation through linkages.

Single Wheel and Tire Assembly The nosewheel assembly includes a wheel and tire. The nose gear tire has chines to deflect water and slush. The tire must be inflated to 120 ± 5 psi or 827 KPa (± 34 KPa). Maximum tire limit speed is 160 knots.

Gear Control Solenoid Valve The gear control solenoid valve regulates the flow of hydraulic fluid to the gear actuators, uplockrelease actuators, and nose gear downlock-release actuator. It is an electrically driven solenoid valve, actuated by two opposing solenoids that respond to electrical commands from the LANDING GEAR handle. When the LANDING GEAR handle is commanded DOWN, the gear-extend solenoid on the valve moves the valve to the gear-extend position, routing fluid pressure to the uplock-release actuators and then to the gear-extend side of the gear actuators (see Figure 14-4). When the LANDING GEAR handle is commanded UP, the gear-retract solenoid on the valve moves the valve to the gear-retract position, which routes fluid to the nose gear downlock release actuator and the gear-retract side of all three gear actuators (see Figure 14-6). Without DC power, the valve centers, releasing pressure from the gear-extend side of the system. This permits emergency extension of the gear using mechanical and pneumatic actuation (see Figure 14-5). DC power can be disconnected from the gear control solenoid valve by pulling the LDG GEAR CONTROL circuit breaker on the SYSTEMS panel of the left CB panel.

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14 LANDING GEAR AND BRAKES

The nose gear assembly (see Figure 14-2) includes: • Strut • Shimmy damper • Drag brace and downlock • Uplock assembly • Gear actuator • Single wheel and tire assembly

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CONTROLS AND INDICATIONS

PLUNGER

LANDING GEAR Control Handle The LANDING GEAR control handle is on the left side of the center tilt panel (Figure 14-7) and controls the normal landing gear retraction and extension procedure. The handle actuates switches to complete circuits to the extend or retract side of the gear control solenoid valve. The gear handle must be pulled out of a detent prior to movement to either the GEAR UP or GEAR DOWN position. On the ground, to keep the gear handle in the GEAR DOWN position, the left and right main gear squat switch deenergizes a locking-solenoid in the instrument panel to extend a spring-loaded plunger into the gear handle path. This prevents inadvertent movement of the handle to the GEAR UP position (Figure 14-8). Airborne, with the left and right main gear squat switches in the in-flight position, the landing gear handle locking solenoid energizes to retract the plunger. This frees the handle for movement to the GEAR UP position. This safety feature cannot be overridden. If the solenoid fails or electrical power is lost, the gear handle cannot be moved to the GEAR UP position.

RETRACT SWITCH EXTEND SWITCH

LOCKING SOLENOID

FW

D

Figure 14-8.   Landing Gear Handle Locking Solenoid and Switches

When the gear handle is up, if the locking-solenoid plunger deenergizes and extends (due to DC power failure or another cause), the plunger does not prevent moving the gear handle to the down position. Never attempt to pull the gear handle up during taxi. Before energizing the aircraft electrical system, ensure that the gear handle is in the down position to prevent inadvertent gear retraction.

14 LANDING GEAR AND BRAKES

In particular, if the squat switches do not agree (as indicated on the EICAS by the WOW MISCOMPARE CAS message), it is possible to raise the gear handle, possibly resulting in gear retraction on the ground.

Circuit Breakers Circuit breakers provide protection for the components and wiring of the landing gear system. Specific components are: • • LH ELE #1 °° HYD PUMP 2 °° LDG GEAR CTL 2 • • LH ELE EMERG °° LDG GEAR MONITOR 2

Figure 14-7.   Landing Gear Control Panel

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One red and three green position indicators on the landing gear control panel provide gear position indication. In addition, an aural warning sounds if gear is up when the combination of throttle position, flap position, and airspeed indicate landing configuration. The green NOSE–LH–RH lights on the LDG GEAR control panel indicate gear down and locked. As each gear locks down, its respective green light illuminates. The red landing gear UNLOCK light indicates an unsafe gear condition. It illuminates when the gear handle is moved out of GEAR UP detent and remains on until all three gear are down and locked. At retraction, the light illuminates when the gear handle is moved out of the down position and remains visible until all three gear are up and locked. Normal indication with the gear down is three green lights visible. With the gear retracted, all lights extinguish and the red UNLOCK light extinguishes. Figure 14-9 shows indicator displays for various gear positions. Test the landing gear indicator lights and warning horn by positioning the rotary TEST knob to LANDING GEAR.

Aural Warning The warning/caution advisory system provides a landing gear aural warning if one or more gear are not locked down and either of the following situations occurs: • Both throttles are retarded below approximately 85% N2 and airspeed is below 130 KIAS. Pressing the HORN SILENCE– PUSH button on the gear control panel (see Figure 14-7) silences this warning. • Flaps are extended beyond the TAKE OFF AND APPROACH setting. In this situation, the aural warning cannot be silenced with the HORN SILENCE– PUSH button.

Aural warning is DC-powered. Circuit protection for the aural warning system and position lights is on the right CB panel labeled WARN LIGHT within the lights grouping.

Rotary TEST Knob The rotary TEST knob is at the top of the copilot panel. This knob is used to test the landing gear and antiskid warning systems.

OPERATION Preflight During preflight, inspect the pressure gauge for the emergency landing gear extension pneumatic bottle. It is in the nose baggage compartment, on the right side of the aft wall. Determine the current temperature, and then compare the gauge pressure indication to the pressures listed on the placard next to the gauge to determine if the pneumatic pressure is appropriate.

Retraction And Extension Moving the LANDING GEAR handle to the GEAR DOWN position energizes the gear control solenoid valve. The DC power for the landing gear control circuit is from the left bus through the LDG GEAR CONTROL circuit breaker in the SYSTEMS section of the pilot CB panel.

Retraction Placing the LANDING GEAR handle in the GEAR UP position energizes the retract solenoid of the gear control valve. The gear control valve is positioned to direct pressure to: • The nose gear downlock-release actuator to release the nose gear downlock • The gear-retract side of each gear actuator (also releases downlocks inside the main gear actuators) • The extend side of the uplock actuators, which position the uplock hooks to catch the rising gear All downlocks are released and retraction begins (see Figure 14-6).

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Gear Position Indicators

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DOWN AND LOCKED

UP AND LOCKED

14 LANDING GEAR AND BRAKES NOSE GEAR NOT DOWN AND LOCKED

ONE OR MORE GEAR NOT UP AND LOCKED

Figure 14-9.   Landing Gear Position Indications

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As each gear reaches the fully retracted position, a spring-loaded uplock hook catches it and an uplock microswitch actuates. When all three uplock microswitches actuate, the gear control solenoid valve circuit is interrupted and the valve returns to the neutral position. All position indicators on the control panel extinguish.

BUNGEE SPRING

STEERING OUTPUT LEVER

STEERING PIN

Extension Placing the landing gear handle in the GEAR DOWN position energizes the gear control solenoid valve to the extend position. The gear control solenoid valve is positioned to route pressure to the uplock actuators, which releases the gear uplocks.

NOSEWHEEL STEERING DESCRIPTION AND OPERATION

Figure 14-10.   Nosewheel Steering

Rudder pedals mechanically steer the nose gear to 20° either side of center. A spring linkage provides an additional 55° of nosewheel deflection (±75° total) via castering accomplished with application of differential engine power or braking. For towing, ensure that the rudder (gust) lock is disengaged and do not exceed 75° nosewheel deflection. If 75° is exceeded, the steering system or airframe structure will be damaged. If the rudder (gust) lock is engaged, towing beyond 55° may cause structural and/or steering system damage. During preflight, check that the stop bolts are present and intact (Figure 14-11). If they are not, the steering system is damaged. Maintenance is required before flight.

WARNING

Mechanical linkage from the rudder pedals mechanically actuates the nosewheel steering system (Figure 14-10). Whenever the nosewheel is extended, nosewheel steering is engaged regardless of whether the aircraft is in flight or on the ground. During retraction, nosewheel steering is disengaged and the nose gear is mechanically centered.

If damage to nosewheel steering is suspected or crewmembers detect abnormal steering system action, do not attempt to fly the aircraft. If the system is damaged, the crew does not have full steering control of the aircraft on takeoff or landing. If the aircraft flies, even if the gear remains extended after takeoff, the nosewheel may not remain centered, and may not be controllable.

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14 LANDING GEAR AND BRAKES

When the spring-loaded uplocks release, pressure continues to the gear-extend side of the gear actuators. As each gear reaches the fully extended position, a downlock microswitch actuates. When all three downlock switches actuate, the gear control solenoid valve circuit is interrupted and the valve returns to the neutral position. With pressure no longer applied to the gear actuator, three actions occur: 1. The internal locking mechanism within each main gear actuator assumes the downlocked position. 2. The spring-loaded nose gear mechanical downlock latches. 3. Downlock switches illuminate the green NOSE–LH–RH position indicators on the gear control panel.

CENTERING FITTING

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ANTISKID SYSTEM With the ANTI SKID switch in the up (on) position, the antiskid system provides maximum braking efficiency on all runway surfaces. The antiskid control system can only reduce pressure; the applied pressure can never be more than that commanded by the crew. A wheel speed transducer on each wheel electronically transmits wheel-speed signals to the antiskid control box as a variable frequency. If the control box detects sudden deceleration of a wheel (impending skid) it commands the antiskid valve to reduce pressure to that specific wheel/brake. When the slow wheel catches up to the fast wheel and the transducer signal returns to normal, braking pressure is restored to the brakes. Figure 14-11.   Stop Bolt Location

CAUTION Anytime the gear is extended, the nosewheel deflects with rudder pedal movement. During a crosswind landing, center the pedals immediately before nosewheel touchdown.

BRAKES DESCRIPTION 14 LANDING GEAR AND BRAKES

Disc brakes are on the main gear assemblies. The aircraft hydraulic system provides normal power braking with a pneumatic (pressurized nitrogen) system for backup (Figure 14-12). The hydraulic system automatically maintains constant pressure for brake operation. The brakes are normally used as antiskid power brakes but can operate as power brakes without antiskid protection. In the event that brake system hydraulic pressure is lost, emergency braking is available. The crew initiates braking by pressing on the tops of the rudder pedals. The pedals connect by cables to actuate the brake metering valve. If both the pilot and copilot apply brakes simultaneously, the one applying the greater force on the rudder pedals has control. 14-14

The antiskid system includes “touchdown protection,” which prevents landing with the brakes locked. Any time both squat switches indicate that the aircraft is in the air and the gear is extended, the antiskid unit dumps brake pressure (except during gear retraction when braking is applied to spindown the wheels). Upon landing, this dump continues for 3 seconds after weight-on-wheels or until wheel spinup (whichever occurs first), before brake pressure is enabled. During high-speed ground movement, “lockedwheel crossover protection” prevents sudden yawing due to differential braking. If the antiskid controller senses a 70% difference in speed between left and right brakes, it reduces brake pressure to both wheels. At low speeds (approximately 12 knots), this feature is disabled to permit tight turns during taxiing.

PARKING BRAKES Parking brakes are a locked configuration of the brakes. Brakes are locked when the parking brake valve traps hydraulic fluid in the brake lines. The valve (and hence the parking brakes) can only be set by pulling on the PARKING BRAKE knob on the right lower side of the pilot instrument panel (Figure 14-13) while pressing on the brake pedals.

FOR TRAINING PURPOSES ONLY

FOR TRAINING PURPOSES ONLY

NITROGEN EXHAUST

CABIN AIR

NITROGEN

METERED PRESSURE

RETURN

SYSTEM PRESSURE

LEGEND OVERBOARD VENT LINE

TO GEAR RETRACT TO GEAR EXTEND

DIGITAL ANTISKID CONTROL UNIT

FROM GEAR RETRACT PRESSURE LINE (FOR TIRE SPINDOWN)

P P P CONTROL MANIFOLD

MANUAL ACCUMULATOR BLEED VALVE

29-VDC HYDRAULIC PUMP

HYDRAULIC SYSTEM RESERVOIR

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14 LANDING GEAR AND BRAKES

Figure 14-12.   Power Brake and Digital Antiskid System

EMERGENCY BRAKE HANDLE

PARKING BRAKE VALVE

SHUTTLE VALVES

ANTISKID CONTROL VALVE

BRAKE METERING VALVE

EMERGENCY BRAKE NITROGEN BOTTLE

EMERGENCY BRAKE VALVE

PILOT/COPILOT RUDDER PEDALS

CABIN AIR

ACCUMULATOR

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CITATION MUSTANG PILOT TRAINING MANUAL

prevent locking the brakes and blowing the tires. The emergency brakes are effective for six pulls on the handle. No electrical power is required for the emergency brakes.

COMPONENTS Brake Pedals Figure 14-13.   PARKING BRAKE Knob

EMERGENCY BRAKES If the hydraulic brake system fails, a pneumatic brake system is available to actuate the wheel brakes (Figure 14-14). The system uses nitrogen pressure from a pneumatic bottle independent of emergency landing gear extension system. Pulling on the EMERGENCY BRAKE handle (under the center of the pilot left tilt panel) actuates the emergency brake system. Pull out and hold the handle aft to apply and modulate emergency braking pressure. The lever has a stop to

Brakes are normally actuated by the pilot or copilot pressing on the tops of one or more of the rudder pedals. Each pedal is mechanically linked to the brake metering valve and to the corresponding pedal. Both pilot and copilot foot forces are transmitted to the brake metering valve by cables. For each wheel, the pilot applying the greater force to the corresponding pedal determines brake pressure to that wheel, and the position of the corresponding pedal for the other pilot. The pilot and copilot pedals move together.

Brake Metering Valve The brake metering valve regulates left or right brake pressure according to brake pedal inputs as commanded by the crew.

POWER BRAKE SYSTEM

14 LANDING GEAR AND BRAKES

SHUTTLE VALVE

SHUTTLE VALVE

EMERGENCY BRAKE VALVE

LEGEND NITROGEN PRESSURE

OVERBOARD VENT LINE

METERED BRAKE FLUID MECHANICAL

EMERGENCY BRAKE NITROGEN BOTTLE

Figure 14-14.   Emergency Brake System

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Wheel Speed Transducers One transducer is in each main gear axle. A drive clip on the wheel hubcap spins the inner rotor of the transducer. Wheel speed data is provided directly to the antiskid control unit.

separately as commanded by the crew and/or the antiskid control unit. However, a shuttle valve at each brake allows high-pressure nitrogen from the emergency brake valve to bypass hydraulic flow and apply pressure directly to both brakes evenly.

CONTROLS AND INDICATIONS

Antiskid Control Valve To prevent skids, the antiskid control valve regulates the distribution of brake pressure, as required to prevent the skid. The antiskid control valve is electrically controlled by the anti-skid control unit (ASCU).

Antiskid Control Unit A digital ASCU monitors wheel speed to provide wheel skid protection and optimum braking efficiency on all runway surfaces. Based on wheel speed inputs, the control unit reduces brake pressure as required to prevent the skid.

ANTI SKID Switch The ANTI SKID switch (see Figure 14-12) is on the LANDING GEAR control panel and is normally in the up (on) position. In the OFF (down) position, the antiskid system deactivates; brake operation remains the same except that antiskid protection is not available. Before turning the antiskid system off, ensure that brake pressure is released. Before turning the antiskid system on, ensure that the wheels are not rotating.

Circuit Breakers The power brake and antiskid systems receive DC power from the left electrical bus.

The parking brake valve is in the brake lines between the antiskid control valve and the brake assemblies. When a crewmember pulls the PARKING BRAKE knob and depresses the brake pedals, the valve engages check valves to trap brake fluid pressure in the brake lines. To relieve pressure due to fluid expansion when the parking brake is engaged shortly after heavy braking, 1,200-psi thermal relief valves are in the parking brake valve.

Emergency Brake Valve The emergency brake valve is lever operated to provide metered pneumatic pressure from the emergency nitrogen bottle directly to the brake assemblies. The emergency brake valve connects through a cable to the emergency brake lever, which is under the instrument panel near the right knee of the pilot.

The HYD PUMP circuit breaker is in the SYSTEM section of the left CB panel. Disengaging the HYD PUMP circuit breaker electrically disables the hydraulic pump. This action reduces or eliminates hydraulic system pressure. This results in limited or improper functioning of the power brake system or completely eliminates the power brake system. The SKID CONTROL circuit breaker is also in the SYSTEM section of the left CB panel. Disengaging the SKID CONTROL circuit breaker disables the antiskid system and touchdown protection.

Rotary TEST Knob The rotary TEST knob is at the top of the copilot panel. This knob tests the antiskid system.

OPERATION

Brakes And Shuttle Valves

Antiskid Touchdown Protection

Disc brakes are in each main gear assembly. The brakes respond to hydraulic or pneumatic pressure. Normally, fluid from the brake metering valve hydraulically actuates the left and right brakes

During landing, the antiskid system “touchdown protection” feature prevents the aircraft from touching down with locked brakes. Touchdown pro-

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14 LANDING GEAR AND BRAKES

Parking Brake Valve

CITATION MUSTANG PILOT TRAINING MANUAL

tection mode is active anytime all three of the following conditions exist: • The ANTI SKID switch is ON. • Both squat switches indicate that the aircraft is in flight. • The gear is extended. Under these conditions (regardless of pilot or copilot pedal position), the touchdown protection mode releases all brake pressure from the brakes. To ensure adequate wheel spinup on contaminated runways, the touchdown protection mode stays active for 3 seconds after the first wheel touches down (either left or right squat switch indicates weight-on-wheels). Under normal conditions, the wheels spinup almost immediately; therefore, a spinup override feature is incorporated. Anytime wheel speed is above 50 knots (regardless of squat switch position), touchdown protection is overridden and normal antiskid braking is available.

Power Braking (Antiskid ON)

14 LANDING GEAR AND BRAKES

For normal operation of the power brake and antiskid system, all three of the following conditions must exist: • The ANTI SKID switch is up (on). • B oth wheels are rotating at aircraft groundspeed. • Either squat switch (left wheel or right wheel) senses weight on wheel. Maximum braking technique is obtained by: 1. Lowering the nose to the ground 2. Firmly applying and holding the brakes until the desired speed has been reached 3. Extending the speedbrakes while applying the wheel brakes

NOTE Do not pump the brakes. The antiskid system is not operative during emergency braking.

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Depressing the brake pedals moves cables attached to the power brake metering valve, which meters hydraulic pressure to the brake assemblies in direct proportion to pedal force. With the ANTI SKID switch on and a ground speed of at least 12 knots, maximum braking with skid protection is available. Any tendency of a wheel to rapidly decelerate (skid) is detected by the wheel speed transducer, and the antiskid control valve is signaled to momentarily dump pressure from affected brakes. As wheel speed returns to normal, dumping ceases and pressure is once again increased in the brake assemblies. When wheel speed drops below approximately 12 knots, the antiskid function disengages. Braking on each main wheel is controlled by the corresponding pedal; therefore, differential braking is available.

Power Braking (Antiskid OFF) The ANTI SKID switch is normally in the up (on) position. In the OFF position, the antiskid system deactivates and the ANTISKID FAIL CAS message appears. The power brakes, powered by the hydraulic system, still function without the assistance of the antiskid system. With the loss of the antiskid system, touchdown protection and tire spin-down are inoperable.

Parking Brakes To set the parking brakes, apply the brakes in the normal manner, then pull out the PARKING BRAKE knob (see Figure 14-13). This mechanically actuates the parking brake valve and traps fluid in the brakes. To release the parking brakes, depress the brake pedals, then push in the PARKING BRAKE knob. One-way check valves allow setting increased trapped pressure once the brakes are set by simply depressing the brake pedals harder. Do not use parking brakes after using emergency brakes. Hold emergency brakes until the aircraft can be secured. Parking brakes can hold the aircraft for only a limited time. They are not intended to secure an

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unattended aircraft. Parking brakes are a temporary function, to be used only until the aircraft can be secured.

EMERGENCY BRAKE PULL

NOTE If brakes are suspected of being hot, do not set the parking brake.

The antiskid system is checked as part of the rotary test procedure in the Cockpit Preparation checklist. This same test can be conducted in flight if a problem is suspected. With the antiskid switch ON (up), the antiskid control unit is continuously conducting integrity checks of the system. If any faults are detected, the ANTISKID FAIL CAS message appears in the CAS window. To ground test the antiskid system: • Rotate the rotary TEST knob to the ANTI SKID test position. • The ANTISKID FAIL and NO TIRE SPINDOWN CAS messages flash for 6 seconds. • Test is valid if the ANTISKID FAIL and NO TIRE SPINDOWN CAS messages are confirmed extinguished after 6 seconds. • Test is failed if the ANTISKID FAIL and/ or NO TIRE SPINDOWN CAS messages remain illuminated after more than 6 seconds.

Emergency Brakes Pulling the red EMERGENCY BRAKE handle aft actuates the emergency brake valve mechanically (see Figure 14-15). The valve meters nitrogen pressure through shuttle valves on the brake assemblies in direct proportion to the amount of lever movement. Since nitrogen pressure is applied to both brakes simultaneously, differential braking is not possible. Returning the lever to its original position releases pressure from the brakes and vents it overboard, which releases the brakes.

Revision 1.1

Figure 14-15.   EMERGENCY BRAKE Handle

Apply the emergency brakes only enough to obtain the desired rate of deceleration, then hold them until the aircraft stops. Best performance can be obtained using a smooth, steady, continuous pull of the handle to obtain the desired deceleration rate. Multiple pulls and releases of the handle deplete the nitrogen charge. Do not depress the brake pedals while applying emergency airbrakes. Shuttle valve action may deplete nitrogen pressure, reducing available braking power. Repeated applications deplete nitrogen pressure. If the emergency nitrogen bottle is full, six applications are available for emergency braking. Antiskid protection is not available during emergency braking. Do not attempt to taxi after clearing the runway when using the emergency brakes. Maintenance action is required subsequent to emergency braking.

LIMITATIONS For specific limitations, refer to the FAA-approved AFM.

EMERGENCY/ ABNORMAL For specific information on emergency/abnormal procedures, refer to the appropriate checklist or FAA-approved AFM.

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14 LANDING GEAR AND BRAKES

Antiskid Test

CITATION MUSTANG PILOT TRAINING MANUAL

Table 14-1.   CAS MESSAGES ANTISKID FAIL DESCRIPTION

INHIBITS

WOW MISCOMPARE

This message displayes when the ACU cannot perform antiskid functions. This message indicates that an internal test routine has failed or there are electrical problems. Electrical problems may include DC power that is off or below operational levels, or shorts or opens in the antiskid system wiring. This message also appears when the rotary TEST knob is selected to ANTI SKID.

DESCRIPTION

EMER

The miscompare may be caused by: • Different wheel positions • Stuck squat switch • Electrical short or open circuit • Problems with the multi-function PCB • Popped L or R SQUAT SWITCH circuit breaker • Loss of power to a squat switch

HYD PRESS LO DESCRIPTION

INHIBITS

The low pressure switch in the hydraulic control manifold controls the amber HYD PRESS LO message. As hydraulic system pressure decreases below 750 psig, the HYD PRESS LO message appears, accompanied by MASTER CAUTION lights. As the pump increases system pressure to greater than 1,000 psig, a circuit opens to extinguish the message and the MASTER CAUTION lights.

INHIBITS

DESCRIPTION

14 LANDING GEAR AND BRAKES

INHIBITS

The amber HYD PUMP ON message indicates that the hydraulic pump has been operating continuously for over 60 seconds. Refer to the checklist. EMER, LOPI

INHIBITS

14-20

EMER, LOPI, TOPI

NO TIRE SPINDOWN

EMER TOPI

HYD PUMP ON DESCRIPTION

The amber CAS message WOW MISCOMPARE indicates that the squat switch system is indicating different status (miscompare) of the two switches. One squat switch appears to indicate weight-on-wheels while the other appears to indicate weight-off-wheels. The message does not display until the miscompare has continued for 2 seconds. This allows for momentary differences during takeoff and landing.

FOR TRAINING PURPOSES ONLY

Ten seconds after gear retraction begins, if either tire is spinning above 10 knots, the white NO TIRE SPINDOWN message appears, indicating failure of the tire spindown function of the antiskid system. This could cause damage to the wheel wells from loose tire tread and debris. This message also appears when the rotary TEST knob is selected to ANTI SKID. EMER, LOPI

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CITATION MUSTANG PILOT TRAINING MANUAL

QUESTIONS

2. With NO TIRE SPINDOWN CAS message displayed in the CAS window: A. Damage to the wheel well from loose tread and debris could result B. Airspeed should be reduced to 150 KIAS C. Do not select flaps beyond approach D. Do not deploy the speedbrakes 3. The ANTISKID FAIL CAS message displays when: A. Loss of touchdown protection B. Loss of the power brakes C. Loss of the emergency brakes D. Loss of the speedbrakes 4. What landing gear systems are affected if a WOW message appears prior to touchdown? A. The antiskid system may not operate normally B. The parking brake may be inoperative C. The emergency brake system may be inoperative D. The power brakes may be inoperative 5. When using the emergency brakes, do not: A. Pump the emergency brake handle B. Use the parking brake C. Use the speedbrakes D. Both A and B 6. Before movement of the aircraft on the ground: A. Disengage the rudder (gust) lock B. Speedbrakes should be extended C. Flaps should be extended to TO D. Emergency exit should be removed

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7. The landing gear uplocks are: A. Mechanically held engaged by springs B. Hydraulically disengaged C. Pneumatically engaged D. Both A and B 8. On the ground, the LANDING GEAR handle is prevented from movement to the GEAR UP position by: A. Mechanical detents B. A spring-loaded locking solenoid C. Hydraulic pressure D. A manually applied handle locking device 9. Landing gear downlocks are disengaged: A. When hydraulic pressure is applied to the retract side of the main gear actuators and the nose gear downlock release actuator B. By action of the gear squat switches C. By removing the external downlock pins D. By mechanical linkage as the gear begins to retract 10. Each main gear wheel incorporates three fusible plugs that: A. Blow out if the tire is overserviced with air B. Melt, deflating the tire if an overheated brake event occurs C. Is thrown out by centrifugal force if maximum wheel speed is exceeded D. None of the above 11. At extension, if the nose gear does not lock in the down position, the gear panel indications are: A. Red light illuminated, green LH and RH lights illuminated B. Red light extinguished, green LH and RH lights illuminated C. Red light illuminated, all three green lights extinguished D. All four lights extinguished

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14 LANDING GEAR AND BRAKES

1. For maximum effective braking, do not switch off the: A. Power brake system B. Antiskid system C. FADEC D. G1000

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12. When using the emergency brake: A. Differential braking is not available B. Antiskid protection is provided C. The handle should be pumped D. Nosewheel steering is inoperative

14 LANDING GEAR AND BRAKES

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CHAPTER 15 FLIGHT CONTROLS CONTENTS INTRODUCTION................................................................................................................ 15-1 GENERAL............................................................................................................................ 15-1 PRIMARY FLIGHT CONTROLS........................................................................................ 15-2 Description.................................................................................................................... 15-2 Aileron System.............................................................................................................. 15-2 Rudder System.............................................................................................................. 15-3 Elevator System............................................................................................................. 15-3 CONTROL LOCK SYSTEMS............................................................................................. 15-4 Aileron/Elevator Control Lock...................................................................................... 15-5 Rudder Lock.................................................................................................................. 15-5 SECONDARY FLIGHT CONTROLS.................................................................................. 15-6 Trim Systems................................................................................................................. 15-6 Flaps.............................................................................................................................. 15-9 Speedbrakes................................................................................................................ 15-11 LIMITATIONS................................................................................................................... 15-13 EMERGENCY/ABNORMAL........................................................................................... 15-13

15 FLIGHT CONTROLS

QUESTIONS..................................................................................................................... 15-14

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15 FLIGHT CONTROLS

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ILLUSTRATIONS Figure Title Page 15-1.

Flight Control Surfaces.......................................................................................... 15-2

15-2.

Rudder Control System Installation...................................................................... 15-4

15-3.

Aileron/Elevator Control Lock.............................................................................. 15-5

15-4.

Rudder Lock System (Left Side of Tailcone)........................................................ 15-5

15-5.

Aileron and Rudder Trim....................................................................................... 15-6

15-6.

Trim Display.......................................................................................................... 15-7

15-7.

Elevator Trim System............................................................................................ 15-8

15-8.

Flaps - LAND Posiiton....................................................................................... 15-10

15-9.

Flap System Schematic....................................................................................... 15-10

15-10. Flaps Position Display........................................................................................ 15-11 15-11. Flaps Position Display - Reversionary Mode..................................................... 15-11 15-12. Rotary TEST Knob............................................................................................. 15-11 15-13. Speedbrakes (Extended)..................................................................................... 15-12 15-14. Throttle Knob Speedbrake Switch (Left Throttle).............................................. 15-12

TABLES Table Title Page

15 FLIGHT CONTROLS

15-1. CAS MESSAGES............................................................................................... 15-13

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CHAPTER 15 FLIGHT CONTROLS

INTRODUCTION This chapter describes the flight controls of the Cessna Model 510 Citation Mustang. The aircraft has fixed and moveable surfaces that provide stability and control during flight. The primary flight controls are ailerons, rudder, and elevators. Secondary flight controls include trim devices, flaps, and speedbrakes. Control locks are also described.

GENERAL

The primary flight controls (elevators, ailerons, and rudder) directly control aircraft movement around the three axes of flight (pitch, roll, and yaw). They are manually actuated through cables by dual conventional control yokes and dual sets of rudder pedals in the cockpit. They can be immobilized by

control locks when on the ground to prevent damage to the control surfaces and systems from wind gusts striking the aircraft. The secondary flight controls include trim, flaps, and speedbrakes. Trim tabs, electrically or mechanically adjusted through controls on the cockpit pedestal or control yoke, assist flight control on all three axes. Mechanical elevator trim, adjusted through a cockpit pedestal wheel, is also provided.

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15 FLIGHT CONTROLS

The flight control systems consist of the control surfaces, trim control surfaces, trim indicating systems, and the related mechanical and electrical systems that control the airplane during flight.

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Flaps and speedbrakes directly adjust airplane lift and drag. Both controls are electrically actuated. Flaps are operated by a handle on the cockpit pedestal. Speedbrakes are operated by a switch on the throttle. All flight control surfaces are shown in Figure 15-1.

PRIMARY FLIGHT CONTROLS

The primary flight controls can also be controlled by the autopilot and yaw damper (see Chapter 16—“Avionics”). The rudder, both elevators, and the left aileron are each equipped with a trim tab that is electrically actuated from the cockpit. The elevator tabs can also be mechanically positioned by the pitch trim wheel on the control pedestal.

AILERON SYSTEM

DESCRIPTION The primary flight controls (ailerons, rudder, and elevators) are manually operated by either the pilot or the copilot through a conventional control yoke and rudder pedal arrangement. Control inputs are transmitted to the control surfaces through cables, bellcranks, and pushrods. The rudder pedals also operate the nosewheel steering and wheel brakes (see Chapter 14—“Landing Gear and Brakes”). A flexible mechanical interconnect between the rudder and ailerons provides improved lateral stability.

Two ailerons (one on the outboard trailing edge of each wing) provide roll control. Neutral aileron position is 2° up. The ailerons are controlled through cables connected to the cockpit control yokes and the autopilot aileron electric servo. The control yoke rotates 70° in each direction to provide maximum aileron deflection.

Operation When the pilot rotates the control yokes counterclockwise, the right aileron rotates down and the left aileron rotates up, causing the aircraft to roll left. By turning the control yokes clockwise, the opposite is true. ELEVATOR TRIM TAB

TRIM TAB FLAP

RUDDER

STRAKE

SPEEDBRAKE

15 FLIGHT CONTROLS

TRIM TAB AILERON

Figure 15-1.   Flight Control Surfaces

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Either pilot can manually override the servo motor by applying force to the control yoke. For information on the AFCS (including autopilot), refer to Chapter 16—“Avionics.”

Aileron-Rudder Interconnect A flexible mechanical interconnect between rudder and ailerons provides improved lateral stability. Movement of the ailerons results in a comparable movement of the rudder (as sensed through the rudder pedals). If the pilot rolls the aircraft to the left, the interconnect also causes some rudder deflection (and resultant airplane yawing) to the left. Conversely, pressure on the rudder pedals and movement of the rudder results in a coordinated movement of the ailerons and control yoke. In flight, to intentionally slip or skid/yaw the airplane, the pilot can override the interconnect by applying opposite forces to the control yoke and rudder pedals (“cross-controlling”). On the ground, the interconnect may cause some aileron and control yoke movement, as a coordinated response to rudder movements caused by the crew steering with the rudder pedals.

RUDDER SYSTEM The rudder on the trailing edge of the vertical stabilizer provides yaw control. It moves as much as 35° left or right of center. It is controlled through cables connected to the cockpit control pedals and the autopilot yaw servo. The rudder is moved by fore and aft movement of the pedals.

Revision 1.0

The rudder pedals are floor-mounted and nonadjustable. The pedals are connected to the rudder through mechanical linkages and cables. Two separate rudder cable loops, routed differently, provide redundancy to protect against an engine rotor noncontainment (Figure 15-2).

Operation Pressing either pilot rudder pedal (left or right) moves the rudder in that direction, which yaws the airplane. Copilot controls work the same. Pilot and copilot pedals are mechanically linked so the pilot applying the greater force controls yawing, and controls the amount of pedal movement for both pilots. The rudder pedals also control nosewheel steering (refer to Chapter 14—“Landing Gear and Brakes.”). The single autopilot yaw servo is mechanically connected to the rudder. When the autopilot is engaged, the yaw servo provides input to the rudder system in response to the AFCS commands. The yaw damper can be disengaged by: • P ressing the YD button on the AFCS controller • P ressing the AP TRIM DISC switch on either control yoke Additionally, pilots can manually override the yaw servo motor by pushing the rudder pedals. For information on the AFCS (including autopilot), refer to Chapter 16—“Avionics.”

ELEVATOR SYSTEM The elevators are on the trailing edge of the horizontal stabilizer and provide longitudinal (pitch) control of the airplane. The elevators are mechanically controlled through cables by either pilot moving the control yoke forward, aft, or by the autopilot pitch servo. The pitch system is a manual system consisting of conventional mechanical flight control components. A cable run from the pilot and copilot control yokes to a common elevator pulley provides output to the elevator surfaces. The aft elevator pulley is attached to each surface by a pushrod and horn.

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15 FLIGHT CONTROLS

When the autopilot is operating, the autopilot roll servo provides inputs to the aileron control system. A single autopilot roll servo is mechanically connected to the aileron cable system. When the autopilot is engaged, the autopilot servo provides autopilot input to the aileron system in response to the automatic flight control system (AFCS) commands. Disengaging the autopilot can be accomplished by three normal means: • The AP or YD button on the AFCS controller • The AP TRIM DISC switch on either control yoke • By commanding pitch trim

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Figure 15-2.   Rudder Control System Installation

Motion from the aft elevator pulley is transmitted to the elevators by their respective pushrod. In the event of engine rotor non containment, separate elevator trim systems provide sufficient pitch control for elevator control redundancy.

Operation By moving the control column aft (approximately 4 inches maximum deflection), the elevators rotate up, causing the nose of the aircraft to pitch up. By moving the control column forward (approximately 3 inches maximum deflection), the opposite motion occurs.

15 FLIGHT CONTROLS

A single pitch servo is mechanically connected to the elevator cables. When the autopilot is engaged, the pitch servo provides autopilot input to the elevator system in response to the AFCS commands.

15-4

Normally, the autopilot can be disengaged by: • Pressing the AP or YD button on the AFCS controller • Pressing the AP TRIM DISC switch on either control yoke • Commanding electric pitch trim The pitch servo can also be manually overridden by either pilot applying a force to the control yoke. For information on the AFCS (including autopilot), refer to Chapter 16—“Avionics.”

CONTROL LOCK SYSTEMS Control locks, when engaged, restrain the primary flight controls. The control lock system prevents damage to the control surfaces and systems from wind gusts striking the aircraft while it is on the ground. There are two parts to the control lock system: aileron/elevator control lock and rudder lock.

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AILERON/ELEVATOR CONTROL LOCK

When removing the lock:

The yoke position and the visual obstruction from the flag provide unmistakable warning of control lock engagement. To insert the lock: • Rotate the yoke to the center position so the receptacle in the bushing and the receptacle on the control yoke are aligned. • Move the yoke forward until both ends of the pin can be inserted into their respective receptacles. • Insert the U-shaped pin of the flag device into the receptacles. • Check the control wheel is locked in both pitch and roll.

Figure 15-3.   Aileron/Elevator Control Lock

RUDDER LOCK The rudder control lock inserts a pin into the aft rudder pulley, preventing movement of the rudder. The rudder lock must be operated from outside the aircraft (Figure 15-4).To lock the rudder, the pilot rotates a handle on an external lever on the left side of the tail cone 60° counterclockwise (up), which inserts a pin into the aft rudder pulley to lock the rudder torque tube. The lock disengages when the external lever is rotated to point aft (streamlined). The lock can also be disengaged from the cockpit by pulling the control yoke aft from the neutral position.

NOTE With the rudder lock engaged, the nosewheel system allows up to 55° of free castering when the pilot steers with differential power and/or differential braking. However, taxiing or steering with rudder lock engaged is not recommended. To release the rudder lock from the cockpit, pull aft on the control yoke.

Figure 15-4.   Rudder Lock System (Left Side of Tailcone)

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15 FLIGHT CONTROLS

To lock the aileron and elevator control surfaces, a removable flag-insert device fits through a hole in both the control yoke bushing at the control panel and the back of the pilot control yoke (Figure 15-3). A special U-shaped lock pin locks the yoke in a most-forward position, nose down with the wheel at ailerons-neutral. The U-shape of the pin ensures that no single pin device can engage the lock. The device, installed from the top, pins the control yoke to the instrument panel. The flag on the pin covers the pilot primary flight display (PFD) airspeed tape and horizontal situational indicator (HSI). The flag portion of the pin is keyed to the instrument panel receptacle so the control lock cannot be installed without obstructing the view of the pilot.

• G rasp the U-shaped pin between the receptacles (with the right hand) and remove, raising it straight up until clear of both receptacles. • Stow the control lock. • Check the control yoke is free and clear for both roll and pitch.

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Aileron (Roll) Trim

CAUTION Disengage the rudder lock before towing. The rudder lock can be released and re-engaged externally.

SECONDARY FLIGHT CONTROLS The secondary flight controls consist of the trim systems for the primary flight controls, and the lift and drag controls (flaps and speedbrakes).

TRIM SYSTEMS Trim is provided by a tab on the inboard trailing edge of most primary flight controls (both elevators, the left aileron, and the rudder). Trim systems are electrical on all three axes, with additional mechanical trim also available for pitch. Rudder and aileron trim are electrically actuated by trim switches on the lower pedestal. The elevator is operated by a manual trim wheel on the left side of the pedestal next to throttles. In addition, the electric trim switches on either pilot control wheel can control the elevator trim.

The single aileron trim tab is on the trailing edge of the left aileron only. The AILERON TRIM control knob (Figure 15-5) controls the aileron trim tab through an electrical trim actuator in the leading edge of the aileron. The electric actuator uses two independent control rods to move the aileron trim tab. The aileron trim control knob is on the aft face of the center pedestal. To operate the knob, depress it before rotating it. Rotating the trim knob left (counterclockwise) causes the aircraft to roll left and trims the left wing down. Near the bottom of the engine indication and crew alerting system (EICAS) display, a white horizontal analog scale and cyan pointer indicate position of aileron trim (Figure 15-6). If cockpit displays are set to reversionary mode, the trim display does not appear. The aileron trim circuit breaker is on the left CB panel in the FLIGHT CONTROLS grouping.

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15 FLIGHT CONTROLS

LEGEND ELECTRICAL CONTROL

Figure 15-5.   Aileron and Rudder Trim

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Rudder (Yaw) Trim The cockpit rudder trim tab is on the center of the trailing edge of the rudder. It is driven by an electrical trim actuator in the leading edge of the rudder, controlled by the RUDDER TRIM switch. The RUDDER TRIM switches (centering dual sliding rocker switches) are on the lower pedestal (Figure 15-5). Pushing the RUDDER TRIM switches to the left, toward the NOSE–L position, causes the aircraft to yaw toward the left. Pushing it right, toward NOSE–R, causes opposite movement. Near the bottom of the EICAS display, a white horizontal analog scale and cyan pointer indicate position of rudder trim (Figure 15-6). If displays are set to reversionary mode, the trim display does not appear.

The mechanical elevator trim system is a single cable loop system that routes from the command wheel in the cockpit to the tail cone, then through the vertical tail to mechanical actuators in each of the horizontal tail. The mechanical actuators move linkages, which move the trim tabs. Additional cables connect the autopilot elevator trim servo to the system. When the control yoke switches are used or the autopilot is active, the autopilot servo electrically commands the entire mechanical system through its cable linkage. The elevator pitch trim servo is powered from the left avionics bus through the AFCS circuit breaker on the FLIGHT CONTROL pilot CB panel. It operates only if the AVN PWR switch is set to the up position. The AVN PWR switch is on the AVIONICS switch panel, below the pilot PFD.

Manual Trim By rotating the trim wheel forward toward the nose-down position, the trim tabs rotate upward, causing the elevator system to pitch the nose of the aircraft down. By rotating the trim wheel aft, the opposite is true. Elevator trim position is indicated by a mechanical pointer, which rides in a slot of the LED-backlit panel of the throttle quadrant, connected to the elevator trim cable loop.

The yaw trim circuit breaker is on the left CB panel in the FLIGHT CONTROLS grouping.

Elevator (Pitch) Trim The elevator trim tabs are at the trailing edges of both elevator surfaces. Both tabs travel synchronously. Each trim tab is connected to a mechanical actuator by two control rods. The trim tabs are controlled mechanically through cables by a mechanical trim control wheel on the left side of the pedestal beside the throttle controls (Figure 15-7), or electrically by split switches on the outboard grip of the control yokes.

Electric Trim Electric elevator trim is controlled by a split-element centering thumb switch on the outboard side of each control yoke (see Figure 15-7). When the pilot moves both elements of the rocker switch, the electric elevator trim actuator moves the cables to trim the aircraft in the direction selected. Selecting forward trims nose-down. The electric trim can be overridden by the mechanical trim. Additionally, pilot pitch trim inputs override the copilot trim inputs. As the trim switch is moved to the UP or DOWN position, the elevator tabs are repositioned as indicated by the elevator trim indicator. Prior to flight, the electric pitch trim system can be checked for proper operation by moving both elements of the switch in both directions, noting

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15 FLIGHT CONTROLS

Figure 15-6.   Trim Display

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LEGEND

MANUAL TRIM

MECHANICAL ELECTRICAL

ELECTRIC TRIM

Figure 15-7.   Elevator Trim System 15 FLIGHT CONTROLS

whether trim occurs in the appropriate directions. Check for system malfunction by attempting to trim with one element of the switch. If trimming occurs, the system is malfunctioning and must be restored to normal operation prior to flight.

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NOTE If a pilot holds only one element of the trim switch in either the UP or DOWN position for more than 3 seconds, the red PTRM message appears on the upper left of the PFD.

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NOTE The autopilot will not engage if electric trim is not operating properly.

FLAPS Flaps on the wings provide control of lift and drag. On the Citation Mustang, flaps increase both lift and drag.

Components The flap system consists of electrical and mechanical components. The flap panels are on the inboard trailing edge of each wing, one on either side of the aircraft (Figure 15-8). They are hinged for operation in three positions: UP (retracted), TO/APR (takeoff/approach), and LAND (landing). Each flap panel is directly connected to a mechanical actuator on the rear wing spar (Figure 15-9). The two actuators are driven through flexible drive shafts connected to a common electric motor (electrically powered, electronically controlled power drive unit). The power drive unit (PDU) is behind the rear wing spar at the aircraft centerline. A mechanical interconnect system links the two flap panels together at their inboard ends via pushrods, pulleys, and cables (Figure 15-9). This system ensures that even with linkage failure, flap position remains synchronized, preventing asymmetrical flap positions.

Controls and Indications A flap handle is in the cockpit, to the right of the throttle levers on the control pedestal (see Figure 15-8). The flap handle can be moved aft from the UP detent and forward from the LAND detent.

The flap handle can be set in any of the three detent positions: • UP—On retraction, should not be selected until TO/APR flap position is achieved • TO/APR • LAND—On extension, may not be selected until TO/APR flap position is achieved Three switches under the flap handle supply command signals for the control and monitoring circuits. The flap handle has a three-position mechanical detent, which requires that the handle be pushed down before it can be moved forward or aft to a new position. Flap panel movement is directly controlled by a flap controller circuit that controls the flap power drive unit (PDU). The flap controller evaluates command signals from the flap handle and position signals from the left interconnect pulley to determine the appropriate operation of the flap PDU to drive the flap movement. When the pilot moves the flap handle from one position to another, the flap controller senses the disagreement between the flap handle position and the flap panel position, and energizes the PDU to move the flap panels until the signals are brought back into agreement. If the flap controller detects a fault, it immediately stops the PDU and goes into idle mode and the FLAPS FAIL CAS message appears. Flap panel position is monitored by a flap monitor circuit and is graphically depicted in the cockpit on the EICAS display, usually on the multifunction display (MFD) (Figure 15-10). To generate the analog flap position indication, the flap monitor evaluates: • Command signals from the flap handle • Position signals from the right interconnect pulley • A monopole signal from the PDU If the flap monitor loses all flap-position inputs (i.e., TO/APR and LAND positions), it replaces the EICAS flap position display with a red “X.”

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Interrupt runaway or malfunctioning trim by depressing the red AP/TRIM DISC switch on the control yoke (see Figure 15-5) and pulling the AFCS circuit breaker in the FLIGHT CONTROLS section of the left CB panel.

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LEGEND ELECTRICAL CONTROL

Figure 15-8.   Flaps - LAND Posiiton FLAP DRIVE SHAFT FLAP ACTUATOR (MECHANICAL)

FLAP POWER DRIVE UNIT (PDU)

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FLAP CONTROL SWITCHES

MECHANICAL INTERCONNECT CABLE

FLAP ACTUATOR (MECHANICAL)

FLAP MONITOR SWITCHES

Figure 15-9.   Flap System Schematic

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Operation Preflight During preflight, visually check that the flap position indication and the flap handle agree on position.

Rotary Test Select the FLAPS position with the rotary TEST knob (Figure 15-12). The flap position display on the MFD is replaced with a red X and the amber STALL WARN FAIL and FLAPS FAIL CAS messages appear for 3 seconds, then extinguish. Figure 15-10.   Flaps Position Display

In reversionary mode, only text appears on the EICAS to indicate flap position (Figure 15-11). Three positions are identified: • UP (fully retracted) • TO/APR (takeoff/approach) • LAND (landing)

Figure 15-11.   Flaps Position Display Reversionary Mode

In emergency power mode, flap information is not shown on the EICAS, because the flap monitor circuit is not powered. The flap system is DC-powered three ways: • The flap drive motor is powered from the left feed bus No. 2, through a current limiter in the aft J-box, to the flap system printed circuit board (PCB). • The flap control is powered from the left feed bus No. 2, through the FLAP CONTROL circuit breaker in the aft J-box. • The flap monitoring system is powered from the left feed bus No. 1, through the FLAPS circuit breaker in the FLIGHT CONTROL section of the left (pilot) CB panel.

Revision 1.1

Figure 15-12.   Rotary TEST Knob

Normal Operation To reposition the flaps, push in on the flap handle and select the desired position. Allow the flap panels to stabilize in the new position. Confirm that flap indication and handle position agree before selecting the next position. Takeoff/approach flaps are limited to airspeed at or below 185 KIAS. Landing flaps are limited to airspeeds at or below 150 KIAS.

SPEEDBRAKES Speedbrakes on the wings provide control of lift and drag. On the Citation Mustang, flaps increase both lift and drag, while speedbrakes increase drag and slightly reduce lift (acting as wing lift spoilers).

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Components The speedbrakes are on the upper and lower surface of the wing forward of the flaps, pivoting on hinge lines at (and parallel to) the aft spar (see Figure 15-1). The speedbrake system consists of an upper panel and a lower panel on each wing, driven by push rods connected to an electromechanical actuator (one on each wing) (Figure 15-13). The speedbrakes can be set at two different positions: stowed and extended. The panels are commanded to either extend or retract; there are no intermediate positions.

SPEEDBRAKE MOMENTARY SWITCH R E

BR

K

T

X

E

EXTEND

FW

GA

D

T

RETRACT

GA

SP

D

RH THROTTLE KNOB LH THROTTLE KNOB

Figure 15-14.   Throttle Knob Speedbrake Switch (Left Throttle)

When a direction is commanded, the actuator moves to that position and remains there until commanded (extend or retract) to the opposite direction. The pilot does not need to hold the speedbrake switch forward or aft during the entire cycle; only a momentary push or pull is required to initiate a sequence. Figure 15-13.   Speedbrakes (Extended)

Controls And Indications The speedbrakes are controlled by a three-position, momentary thumb switch on the outboard side of each throttle lever knob in the cockpit (Figure 15-14). The pilot uses the switch to select the desired speedbrake position (RET or EXT). Speedbrake position indication is provided by the SPD BRK EXTEND CAS message when the speedbrakes are not in the stowed position.

NOTE A commanded extension can immediately be reversed by the pilot and the speedbrakes will stow. However, when a retraction is commanded, the speedbrakes cannot be reversed (opened) until fully stowed (closed). The speedbrakes control-and-monitoring circuit monitors speedbrake positions and commands retraction when either throttle is set for greater than approximately 85% TLA.

Operation 15 FLIGHT CONTROLS

Move the momentary switch on the throttle control lever knob aft to EXT (extend) and forward to RET (retract). The speedbrakes extend when the speedbrake switch is pushed to EXT and the throttle levers are set for less than approximately 85% TLA.

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The speedbrake system is DC-powered three ways: • The left speedbrake actuator and the speedbrakes control-and-monitoring circuit are powered from the left electrical bus No. 1, through the L SPEED BRAKE circuit breaker in the FLIGHT CONTROL section of the left (pilot) CB panel. • The right speedbrake actuator is powered from the right electrical bus No. 1, through the R SPEED BRAKE circuit breaker in the FLIGHT CONTROL section of the right (copilot) CB panel. • The speedbrake position sensors (proximity switches) are powered from the left feed bus No. 2, through the SPD BRK SW circuit breaker in the aft J-box. In the event of DC power failure, the speedbrakes remain in their current position.

Table 15-1.   CAS MESSAGES FLAPS FAIL DESCRIPTION

INHIBITS

Indicates flap system failure has occurred. Message may or may not coincide with loss of all flap indication, which results in removal of the analog flap signal and a red X on the EICAS flap display EMER

STALL WARN FAIL DESCRIPTION INHIBITS

Indicates a failure has been detected in the stall warning system. LOPI, TOPI, ESI, EMER

SPD BRK EXTEND DESCRIPTION INHIBITS

Indicates the speedbrakes are extended on either side. EMER

LIMITATIONS For specific limitations, refer to the FAA-approved AFM.

EMERGENCY/ ABNORMAL

15 FLIGHT CONTROLS

For specific information on emergency/abnormal procedures, refer to the appropriate checklist or FAA-approved AFM.

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QUESTIONS 1. Before towing the aircraft: A. The rudder gust lock must be unlocked B. All intake plugs should be removed C. The pitot covers should be removed D. Fuel vents should be cleared 2. With a total electrical failure, the only trim available is the: A. Manual aileron B. Manual rudder C. Manual elevator D. All of the above 3. In reversionary mode, the flap position indication is displayed: A. As normal on the EICAS B. As text only on the EICAS C. On the left MFD D. On the right side of the MFD 4. A white SPD BRK EXTEND message in the CAS window indicates: A. A loss of hydraulic power to the speedbrakes B. Speedbrakes are not in the stowed position C. A loss of electrical power to the speedbrakes D. A loss of pneumatic power to the speedbrakes

15 FLIGHT CONTROLS

5. If hydraulic power is lost: A. The flaps are inoperative B. The flaps operate with the backup electrical system, but extend and retract at a reduced rate C. There is no effect on wing flap operation D. A split flap condition could result if the flaps are lowered

7. The speedbrakes fully retract if: A. A complete electrical failure occurs B. A hydraulic failure occurs C. Either throttle is advanced above approximately 85% TLA position with the electrical system operating normally D. Hydraulic quantity drops below 0.2 gallons 8. Speedbrakes must not be extended within: A. 50 feet AGL on landing B. 110 feet AGL on landing C. 40 meters D. 50 meters 9. The wing flaps: A. Can be preselected to only three positions (UP, TAKEOFF, and LANDING) B. Depend on both actuators to function to prevent a split-flap condition C. Can be lowered manually if electrical power is lost, but only if all hydraulic fluid has not been lost D. Can be selected to the GROUND FLAP position on the ground or in flight; the ground flap selection is prohibited in flight 10. If the flap position is unknown (red “X” on the flap indicator), the maximum KIAS is: A. 250 B. 200 C. 180 D. 150 11. The maximum airspeed with the flaps in the TO/APR position is: A. 140 KIAS B. 150 KCAS C. 185 KIAS D. 200 KCAS

6. The ailerons are operated by: A. Hydraulic pressure B. Mechanical inputs from the control wheels C. A fly-by-wire system D. An active control system that totally eliminates adverse yaw 15-14

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CHAPTER 16 AVIONICS CONTENTS INTRODUCTION................................................................................................................ 16-1 GENERAL............................................................................................................................ 16-1 G1000 Integrated Flight Deck Overview...................................................................... 16-3 Standby Flight Instruments Overview........................................................................... 16-3 Air Data Reference Sensors.......................................................................................... 16-3 G1000 ARCHITECTURE.................................................................................................... 16-4 Data Communications................................................................................................... 16-4 Garmin Integrated Avionics Units................................................................................. 16-4 Displays......................................................................................................................... 16-7 Other Units.................................................................................................................... 16-7 AVIONICS POWER SWITCHES........................................................................................ 16-9 Battery Switch............................................................................................................... 16-9 Avionics Power Switch............................................................................................... 16-10 Standby Flight Instruments Switch............................................................................ 16-10 PRIMARY FLIGHT DISPLAY......................................................................................... 16-10 Description................................................................................................................. 16-10 Graphical Flight Instrumentation............................................................................... 16-12 Inset Map.................................................................................................................... 16-19 Auxiliary Information Window.................................................................................. 16-19 Reversionary Mode.................................................................................................... 16-20 NAV/COM Frequencies and Navigation Data Windows............................................ 16-20 PFD Controls.............................................................................................................. 16-20

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Transponder Display and Control............................................................................... 16-25 Supplemental Flight Data........................................................................................... 16-25 MULTIFUNCTION DISPLAY.......................................................................................... 16-27 Description................................................................................................................. 16-27 MFD Controls............................................................................................................ 16-27 Engine Indicating and Crew Alerting System ........................................................... 16-32 Main MFD Display Area............................................................................................ 16-32 Navigation Status Box................................................................................................ 16-35 FLIGHT MANAGEMENT SYSTEM............................................................................... 16-35 Weight Planning......................................................................................................... 16-35 FMS and Flight Plans................................................................................................. 16-35 MFD/FMS Flight Plan Controls and Indications....................................................... 16-36 User-Defined Waypoints............................................................................................. 16-36 IFR Procedures........................................................................................................... 16-36 Vertical Navigation..................................................................................................... 16-36 AUTOMATIC FLIGHT CONTROL SYSTEM................................................................. 16-37 Description................................................................................................................. 16-37 Controls and Indications............................................................................................. 16-40 Operation.................................................................................................................... 16-41 SYNTHETIC VISION SYSTEM (SVS)........................................................................... 16-43 Operation.................................................................................................................... 16-43 AUDIO PANEL................................................................................................................. 16-47 Description................................................................................................................. 16-47 Controls and Indications............................................................................................. 16-47 TRAFFIC INFORMATION SERVICE............................................................................. 16-49 Traffic Advisory System............................................................................................. 16-50

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AIRBORNE WEATHER RADAR.................................................................................... 16-51 Weather Radar Page and Controls.............................................................................. 16-52 Radar Display and Indications................................................................................... 16-52 Antenna Stabilization................................................................................................. 16-53 Antenna Tilt................................................................................................................ 16-53 Ground Mapping........................................................................................................ 16-53 TERRAIN AWARENESS AND WARNING SYSTEM.................................................... 16-53 Hazard Depictions and Alerts..................................................................................... 16-53 XM Weather and GDL 69/69A Data Link................................................................. 16-54 STANDBY FLIGHT INSTRUMENTS............................................................................. 16-55 Standby Attitude Indicator.......................................................................................... 16-57 Standby Airspeed Indicator........................................................................................ 16-57 Standby Altimeter Display......................................................................................... 16-58 Standby Magnetic Compass....................................................................................... 16-58 LIMITATIONS................................................................................................................... 16-59 EMERGENCY/ABNORMAL........................................................................................... 16-59 QUESTIONS..................................................................................................................... 16-60

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ILLUSTRATIONS Figure Title Page 16-1.  G1000 Integrated Flight Deck............................................................................... 16-2 16-2.  G1000 Integrated Avionics Architecture............................................................... 16-5 16-3.  Battery Switch....................................................................................................... 16-9 16-4.  Avionics Power Switch and Standby Instruments Switch.................................. 16-10 16-5.  PFD Graphical Callouts..................................................................................... 16-11 16-6.  Airpseed Indicator.............................................................................................. 16-12 16-7.  Red Pointer......................................................................................................... 16-12 16-8.  Flap Speed.......................................................................................................... 16-13 16-9.  AFCS Reference................................................................................................. 16-13 16-10.  Attitude Indication.............................................................................................. 16-14 16-11.  Pitch Attitude Warnings..................................................................................... 16-14 16-12.  Slip/Skid Indication............................................................................................ 16-14 16-13.  Flight Director Single-Cue Command Bars....................................................... 16-14 16-14.  Flight Director Cross-Pointer Command Bars................................................... 16-15 16-15.  Altimeter............................................................................................................. 16-15 16-16.  Altitude Alerting Display................................................................................... 16-15 16-17.  Barometric MDA Displays................................................................................. 16-16 16-18.  Barometric MDA Altitude Alert Setting............................................................ 16-16 16-19.  Horizontal Situation Indicator (HSI).................................................................. 16-17 16-20.  Navigation Sources............................................................................................ 16-17 16-21.  Glide-Slope Indicator......................................................................................... 16-19 16-22.  Marker Beacon Indications................................................................................ 16-19 16-23.  Vertical Speed Indicator..................................................................................... 16-19 16-24.  G1000 System Messages.................................................................................... 16-20

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16-25.  Primary Flight Display....................................................................................... 16-21 16-26.  Softkey Chart for Garmin Software (Sheet 1 of 2)............................................ 16-23 16-26. Softkey Chart for Garmin Software (Sheet 2 of 2)............................................. 16-24 16-27.  HSI with Bearing and DME Indication.............................................................. 16-26 16-28.  Wind Data Box................................................................................................... 16-26 16-29.  Ram Air Temperature Box.................................................................................. 16-26 16-30.  System Time Box............................................................................................... 16-27 16-31.  Timer Reference Window................................................................................... 16-27 16-32.  Multifunction Display........................................................................................ 16-28 16-33.  MFD Controller.................................................................................................. 16-29 16-34.  MFD with MAP Displays.................................................................................. 16-33 16-35.  VNAV Indications.............................................................................................. 16-37 16-36.  Automatic Autopilot........................................................................................... 16-37 16-37.  AFCS Controller................................................................................................. 16-38 16-38.  AFCS Status Bar................................................................................................ 16-40 16-39.  AFCS Status Box................................................................................................ 16-41 16-40.  Control Yoke Switches....................................................................................... 16-41 16-41.  CWS Display...................................................................................................... 16-41 16-42.  GA Switch.......................................................................................................... 16-41 16-43.  Go-Around Mode............................................................................................... 16-42 16-44.  Emergency Descent Mode................................................................................. 16-42 16-45.  SVS on the PFD................................................................................................. 16-45 16-46.  Terrain Alerts with SVS on the PFD.................................................................. 16-46 16-47.  Audio Panel........................................................................................................ 16-47 16-48.  Traffic and Terrain Display................................................................................. 16-53 16-49.  Terrain Colors..................................................................................................... 16-54

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16-50.  NEXRAD Legend.............................................................................................. 16-56 16-51.  Standby Attitude Indicator................................................................................. 16-57 16-52.  Standby Airspeed Indicator................................................................................ 16-57 16-53.  Standby Altimeter............................................................................................... 16-58 16-54.  Magnetic Compass............................................................................................. 16-58

TABLES Table Title Page 16-1.

G1000 SYSTEM COMPONENTS....................................................................... 16-6

16-2.

V-SPEED TABLE............................................................................................... 16-13

16-3.

Flight Phases and CDI Scaling........................................................................... 16-18

16-4.

VERTICAL DEVIATION DISPLAY................................................................. 16-19

16-5.

PRIMARY FLIGHT DISPLAY CONTROL DESCRIPTIONS......................... 16-22

16-6.

MFD SOFTKEYS.............................................................................................. 16-30

16-7.

MFD CONTROLLER DESCRIPTIONS........................................................... 16-31

16-8.

AFCS CONTROLLER DESCRIPTIONS.......................................................... 16-39

16-9.

AUDIO PANEL DESCRIPTIONS..................................................................... 16-48

16-10. PRECIPITATION INTENSITIES...................................................................... 16-52 16-11. GROUND TARGET RETURN INTENSITY LEVELS.................................... 16-53

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CHAPTER 16 AVIONICS

INTRODUCTION This chapter is an overview of the avionics systems and does not contain complete details of every part of each system. Detailed operational information on the G1000 integrated flight deck system is available in the Garmin Pilot’s Guide as revised for the Cessna Citation Mustang. It is incumbent upon the pilot to adhere to the procedural policies stated within Garmin and Cessna FAA-approved documents, which include warnings, cautions, and notes. Refer to Chapter 1—“Aircraft General” for a list of Mustang publications.

GENERAL The Cessna Citation Mustang utilizes a highly integrated electronics/ instrumentation package (Figure 16-1). The Garmin G1000 integrated flight deck avionics suite is the main element of the system.

Revision 1.0

In addition to normal flight operations with the G1000, standby and manual systems provide backup capabilities for essential flight operations and system control.

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AUDIO PANEL

PILOT PRIMARY FLIGHT DISPLAY

FOR TRAINING PURPOSES ONLY MFD/FMS CONTROLLER

MULTIFUNCTION DISPLAY

STANDBY ALTIMETER

Figure 16-1.   G1000 Integrated Flight Deck

STANDBY AIRSPEED INDICATOR

AUTOMATIC FLIGHT CONTROL SYSTEM MODE CONTROLLER

STANDBY ATTITUDE INDICATOR

AUDIO PANEL

COPILOT PRIMARY FLIGHT DISPLAY

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G1000 INTEGRATED FLIGHT DECK OVERVIEW

STANDBY FLIGHT INSTRUMENTS OVERVIEW

The G1000 provides the pilot with communication, navigation, flight guidance, flight instrumentation, and monitoring of most aircraft systems. Functions are performed by various individual units. Three large displays and four control panels give the crew access to all functions.

The G1000 is supplemented by three 2”-display, stand-alone backup flight instruments on the topcenter area of the instrument panel. These flight instruments are powered by the standby instrument battery pack: • Airspeed indicator • Altimeter • Attitude indicator

Two primary flight displays (PFDs) provide flight instrument indications. The larger multifunction display (MFD) provides a moving map display, incorporating information from navigation instruments, terrain/obstruction databases, weather information, and traffic alerts. The MFD also provides engine and airframe systems monitoring.

AIR DATA REFERENCE SENSORS

Supporting G1000 units provide data to these three displays, including air data, attitude and heading data, communications, navigation, transponders, weather radar, satellite-broadcast weather information, traffic information, and sensors for the engines and airframe systems.

Outside air data is supplied to the Mustang avionics (air data computer and standby instruments) through dual pitot-static systems, dual outside-air probes and a stall-warning vane. Two pitot probes, one on each side of the nose, supply ram-air inputs to the respective side (pilot/left or copilot/right) air data computer (ADC). A separate, two-port static system connects to each ADC. To minimize yaw effects, both static systems have a static port on each side of the fuselage.

The G1000 analyzes aircraft systems status and report crew alerts to the displays for various emergency, abnormal and advisory situations. The G1000 also provides automatic flight control, flight direction and an integrated flight management system. Finally, the processors automatically detect, report, and adjust for abnormal conditions in the G1000.

All four static ports and both pitot tubes are electrically heated whenever the PITOT-STATIC switch (in the ICE PROTECTION switch panel) is on. To ensure continued air data reference if normal DC power fails in icing conditions, the pilot pitot-static system is electrically heated through the emergency bus (refer to Chapter 10—“Ice and Rain Protection”).

NOTE

Two outside air temperature (OAT) probes are on the left fairing below the cabin door (one for each ADC). The ADCs analyze OAT levels and pitotstatic inputs, and convert the information to data for the other components and displays of the G1000.

While most monitoring of engine and airframe systems is through the displays, control of these systems is mostly performed through conventional electrical and mechanical cockpit controls, or through cockpit controls that command the independently powered full authority digital engine controls (FADECs).

However, the standby altimeter and air data displays bypass the ADCs and receive reference inputs (static and ram-air pressures) directly from the pilot-side pitot-static system. The stall-warning computer processes signals from the stall-warning vane (on the copilot side of the fuselage). The stall warning computer sends normalized angle of attack (AOA) to the G1000 to

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display a reference approach cue speed, 1.3 VS1, represented as an open green circle on the airspeed tape. A red low-speed awareness range extends from the bottom of the airspeed tape up to the low speed velocity, VLSA. The stall warning computer also sends an impending stall signal to disconnect the autopilot (AP). An aural stall warning tone is heard in the speakers and headsets when airspeed is below VLSA. The stall warning system also provides three crew alerting system (CAS) messages: STALL WARN FAIL , STALL WARN HTR , and STALL WARN HI . The STALL WARN HI CAS message appears when the stall warning system is operating on the ice-contaminated schedule (refer to Chapter 10—“Ice and Rain Protection” for more details).

G1000 ARCHITECTURE The G1000 is a system of individual line-replaceable units (LRUs), which integrate into a modular avionics system that provides: • Flight instrumentation • Navigation and hazard avoidance • Flight guidance • Communications • Monitoring of aircraft systems The pilot and copilot monitor and operate the instruments and avionics, and some aircraft systems, through the displays and control panels (Figure 16-2 and Table 16-1). The G1000 provides system redundancy through the use of dual, parallel systems (one for pilot and one for copilot), with cross-side connections to provide maximum capability to both sides, and to ensure system redundancy if a failure occurs. Any one of the three displays is capable of displaying all critical flight information upon pilot command in the event of a display failure.

DATA COMMUNICATIONS LRUs communicate with each other through various types of data communication lines. Refer to Figure 16-2 for an interface diagram.

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GARMIN INTEGRATED AVIONICS UNITS The G1000 is regulated and coordinated by central processing computers in the two Garmin integrated avionics units (GIAs), which also contain the essential navigation and communications avionics: • NAV/COM • Instrument landing system (ILS) • Global positioning system (GPS) • Flight director (FD) Each GIA receives additional information from its onside ADC and attitude and heading reference system (AHRS). Finally, each GIA monitors engine/airframe sensors directly, or through Garmin engine/airframe (GEA) interface units. All outputs from the GIAs are displayed on the PFDs and/or MFD. In addition to the main processors, specific features include: • Wide-area augmentation system (WAAS)enabled, 12-channel parallel GPS receiver (simultaneously tracks and uses up to 12 satellites). • Very high frequency communication (VHF COM) transmitter providing frequencies from 118.00 to 136.990 MHz, in 25 kHz (760-channel) or 8.33 kHz (3040-channel) spacing. • Very-high frequency omnidirectional range/ILS localizer (VOR/LOC) receiver tuning 108.00 to 117.95 MHz, at 50 kHz increments. • ILS glide slope receiver tuning 328.6 to 335.4 MHz, as matched with the ILS frequency tuned in the VOR/LOC receiver. • FD processor, which interfaces with the GFC 700 automatic flight control system (AFCS) • Digital aural warnings

NOTE Marker beacon reception is in the audio panel, which connects to the GIAs.

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HIGH-SPEED DATA BUS

#1 AIR DATA COMPUTER (GDC 74B) • OAT • AIRSPEED • ALTITUDE • VERTICAL SPEED

#1 INTEGRATED AVIONICS UNIT (GIA 63W) • G1000 PROCESSORS • VHF COM • VHF NAV/LOC • GLIDESLOPE • GPS • FLIGHT DIRECTOR

#2 AIR DATA COMPUTER (GDC 74B)

XM WEATHER DATALINK (GDL 69A)

• OAT • AIRSPEED • ALTITUDE • VERTICAL SPEED

WEATHER RADAR (GWX 68)

#1 ATTITUDE /HEADING REFERENCE SYSTEM (GRS 77)

BACKUP GPS #2 ATTITUDE /HEADING REFERENCE SYSTEM (GRS 77)

#1 MAGNETOMETER (GMU 44) MAGNETIC HEADING

• ATTITUDE • RATE OF TURN • SLIP/SKID • HEADING

#2 MAGNETOMETER (GMU 44)

• ATTITUDE • RATE OF TURN • SLIP/SKID • HEADING

MAGNETIC HEADING

BACKUP GPS XPDR1 MODE S TRANSPONDER w/DIVERSITY (GTX33D) #1 ENGINE /AIRFAME ADAPTER (GEA 71)

ANALOG & DISCRETE ENGINE/ AIRFRAME SYSTEMS SENSORS

#2 INTEGRATED AVIONICS UNIT (GIA 63W) • G1000 PROCESSORS • VHF COM • VHF NAV/LOC • GLIDESLOPE • GPS • FLIGHT DIRECTOR

XPDR2 MODE S TRANSPONDER (GTX33) ANALOG & DISCRETE ENGINE/ AIRFRAME SYSTEMS SENSORS

#2 ENGINE /AIRFAME ADAPTER (GEA 71)

DISCRETE ENGINE/AIRFRAME SYSTEMS SENSORS

DISCRETE ENGINE/AIRFRAME SYSTEMS SENSORS

PITCH SERVO (GSA 81) PITCH TRIM SERVO (GSA 81) YAW SERVO (GSA 80) ROLL SERVO (GSA 80) LEFT EXTERNAL ARINC 429 LRUs • L/R FADEC • L FUEL • PRESSURIZATION • CESSNA DIAGNOSTICS • ELT

LEGEND HIGH-SPEED DATA BUS ARINC 429 DATA INTERFACE CABLE (RS-485) DATA INTERFACE CABLE (RS-232) REVERSIONARY CONTROL

RIGHT EXTERNAL ARINC 429 LRUs • L/R FADEC • R FUEL

Figure 16-2.   G1000 Integrated Avionics Architecture

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Table 16-1.   G1000 SYSTEM COMPONENTS ITEM

FUNCTION

GIA 63W

Interface Adapter

GDU 1040A

Primary Flight Display (PFD) Unit

Left and right color LCD displays. Each PFD displays flight instruments and flight guidance and basic avionics indications. MFD information may also appear on the PFDs and vice versa. TAWS is a built-in feature.

GDU 1500

Multifunction Display (MFD)

Center-panel color LCD display. The MFD displays a multifeature map. Also display EICAS.

GMC 710

Automatic Flight Control System (AFCS) Controller

GSA 80

High Torque Servo

Yaw and roll servos for AFCS. Part of GFC700.

GSA 81

Low Torque Servo

Pitch and pitch trim servos for AFCS. Part of GFC700.

GCU 475

MFD/FMS Controller

MFD and flight management system controller on lower cockpit pedestal.

GRS 77

Attitude and Heading Reference System (AHRS)

GMU 44

Magnetometer

GDC 74B

Air Data Computer (ADC)

Processes inputs from pitot-static system, temperature sensors to determine airspeed, altitude, and vertical speed.

GEA 71

Engine/Airframe Unit (GEA)

Receives and processes signals from various engine and airframe sensors.

GMA 1347D

Audio Amp and Marker Beacon

Audio panel for controlling audio sources; integrated marker beacon receiver.

KN63

Distance Measuring Equipment (DME) System

GTX 33

Transponder

Mode S transponder—Conventional Mode C transponder plus additional capability for data communications with ground radar and traffic information systems.

GTX 33D

Diversity Transponder

Mode S transponder—Conventional Mode C transponder plus additional capability for data communications with ground radar and traffic information systems and Mode S diversity.

GWX 68

Weather Radar

GDL 69A

Datalink

RA350 2

Automatic Direction Finder (ADF) System

Provides ADF bearing to the G1000 displays (optional).

C406-N

Emergency Locator Transmitter (ELT) System

When activated, the ELT sends out a distress signal of 121.5, 243.0, and 406.028 MHz. The 406.028 MHz signal includes important information about the airplane including latitude and longitude, which is detected by the CosPas Sarsat Satellite System for search and rescue operations.

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DESCRIPTION Main modules of the G1000 system. Coordinates information input from all sources to the cockpit displays. Also receives discrete data and ARINC 429 from various engine systems, and processes for presentation to other G1000 components. Each GIA includes a WAAS-enabled GPS receiver, VOR/ILS localizer receiver, ILS glide-slope receiver, FD system, and a VHF communications transceiver.

Located in top center panel. Part of GFC700 system.

Provides aircraft attitude and heading guidance to G1000 displays and AFCS.

Measures local magnetic field information.

Provides DME distance information to G1000 displays.

Provides airborne weather and ground map weather radar to MFD. Satellite radio receiver that provides real-time weather information to the MFD as well as digital audio entertainment. Communicates with the MFD via HSDB connection. A subscription service required to enable.

FOR TRAINING PURPOSES ONLY

DISPLAYS

OTHER UNITS

Two PFDs and one MFD provide a central display and crew interface for the G1000. Various knobs and softkeys provide system control.

Most elements of the Mustang avionics system are LRUs. Each LRU is a self-contained avionics module that can be removed from the airplane and replaced, independent of all other systems. Most LRUs are panel-mounted or in a rack immediately behind the MFD in the center panel. See Table 16-1 for a general overview of Mustang avionics modules, including LRUs.

The PFDs are two identical, 10.4-inch color liquid crystal displays (LCDs) in the instrument panel (see Figure 16-1). Each PFD provides flight instrument displays and basic avionics indications (NAV/ COM and transponder settings, course deviation indicator (CDI) and ILS indications. The 15-inch MFD provides a moving-map display and indications for most airframe and engine systems. The moving map display indicates current aircraft position relative to topography and surface features, terrain obstructions, airspace boundaries, airways, aviation facilities (including airports and navaids), and weather. The left side of the MFD provides indications for engine and aircraft systems and crew alerts.

Reversionary Mode In the event of a screen failure, the essential information from the PFDs and MFD can be combined onto the remaining screens by crew selection of the DISPLAY BACKUP button at the bottom of the audio panels. This ensures availability of adequate information for continued flight. With some software versions, the Garmin G-1000 system automatically switches into reversionary mode with the loss of the MFD screen.

Display Controls The crew uses controls on the bezels of each PFD and MFD to command various instrument, avionics, and aircraft system settings. Along the bottom of each display, variable-purpose “softkeys” provide multiple control functions, depending upon flight conditions or settings selected by the crew. The function of each softkey appears immediately above it on the display.

Attitude And Heading Reference System The remote-mounted GRS 77 AHRS and GMU 44 magnetometer combine to replace conventional gyros and magnetic compass systems with longlife, solid-state sensors. Each PFD has its own AHRS (AHRS1 for pilot PFD, AHRS2 for copilot PFD), which also connects to the same-side GIA. An AHRS combines the functions of an attitude gyro, directional gyro and turn-and-slip instrument. Each AHRS is electrically stabilized by retrieving information from three other sources besides itself. The magnetometer, ADC, and GIAs provide this supplemental data. These three external sources provide reference information that also enables the AHRS to function if powered on after power interruption in flight, and begin providing valid guidance within seconds.

Magnetometer Each GMU 44 magnetometer is a magnetic sensor that provides local magnetic field information to its corresponding AHRS. The magnetometers are in the vertical tail to minimize magnetic influence from aircraft structures and contents.

Additional control of avionics and systems is provided through the AFCS controller on the top center panel, and by the MFD/FMS controller on the lower cockpit pedestal.

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Air Data Computer Each GDC 74B air data computer (ADC1 for pilot, ADC2 for copilot) is a remote-mounted device that provides air data to the GIAs and the PFDs, including: • Air temperature • Pressure altitude • Density altitude • Vertical speed • Indicated airspeed • True airspeed • Mach number Each ADC measures aircraft static and impact pressure information from pressure transducers connected to the same-side (pilot or copilot) pitotstatic system and raw air temperature data from its own outside temperature probe. Using the raw data, each ADC unit computes the air data values, then sends them to its corresponding GIA and PFD. The system is reduced vertical separation minimum (RVSM) compliant. Each ADC also communicates with the AHRS to provide stabilization and orientation information.

Engine/Airframe Interface Unit Each GEA 71 interface unit is a computer that monitors analog and discrete (digital) sensors on airframe and engine systems, and translates these into system indications and alerting outputs to the GIAs. Each GEA interface unit supplies information to both GIAs. The GIAs process this information further, and distribute it to other systems, particularly to the engine indicating and crew alerting system (EICAS) display (normally presented on the MFD).

FMS and MFD Controller A flight management system (FMS) is integrated into the processors for the displays and GIAs. The system is controlled with the the GCU 475 FMS and MFD controller, which is on the cockpit pedestal below the throttles. The FMS controller also duplicates many of the functions of comparable controls on the PFDs when used to control settings and displays on the MFD.

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Automatic Flight Control System The Mustang has a GFC700 AFCS. The system functions are distributed across various units: • The AFCS controller is under the center glareshield and has the mode-select buttons. • Each GIA performs mode logic and FD computations. • The servos compute and monitor for the AP, yaw damper (YD), auto trim and manual electric pitch trim functions. • The PFDs display FD commands and mode annunciations. Yoke-mounted and throttle-mounted switches complete the system control inputs: • CWS switch • AP DISC switch • GA switch

GTX33 Mode S Transponder and GTX33D Mode S Diversity Transponder The G1000 includes two transponders: A GTX 33 Mode S transponder for the copilot side, and a GTX 33D Mode S diversity transponder for the pilot side. Each transponder connects (through its sameside GIA) to the PFDs for control and display. Both transponders provide mode A (normal), mode C (altitude encoding) and mode S (data communications) functions. The GTX 33D provides diversity capability. Both transponders are automated transceivers operating on radar frequencies, receiving ground radar and traffic alert and collision avoidance system (TCAS) interrogations at 1030 MHz, then transmitting a coded response to ground-based radar at 1090 MHz.

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Ground stations can interrogate mode S transponders individually using a 24-bit International Civil Aviation Organization (ICAO) mode S address, which is unique to the particular aircraft. In addition, ground stations may interrogate a GTX 33 for its transponder data capability and the aircraft flight identification, which is the registration number or other call sign. The GTX 33 makes the maximum airspeed capability (set during configuration setup) available to TCAS systems on board nearby aircraft to aid in the determination of TCAS advisories. The unit includes an altitude monitor and traffic information service (TIS). Altitude and traffic alerts are announced by a voice or tone audio output. The PFD displays the code, reply indication, and operating mode. The MFD displays TIS graphical information, which may also appear in the PFD inset map. A traffic alert causes the PFD inset maps to automatically appear.

Audio/Marker Beacon System The GMA 1347D audio amplifier and marker beacon receiver is a panel-mounted system. The unit has a microcontroller for processing front panel key commands, annunciator control, input/output functions, and communication. It includes an intercom system (ICS) with public-address (PA) function.

Weather Radar The GWX 68 weather avoidance radar provides real-time radar information, including precipitation and ground-mapping returns to the G1000. Returns are displayed on the MFD. The GWX 68 communicates though the high-speed data bus by way of the GDL69A.

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AVIONICS POWER SWITCHES Three switches control power to the Citation Mustang avionics and instruments: • Battery switch • Avionics power switch • Standby instruments switch

BATTERY SWITCH The battery toggle switch is in the DC POWER section of the left lower instrument panel and has three positions: BATT, OFF, and EMER (Figure 16-3). The switch controls DC power to the other switches, and directly supplies power to components required for EICAS operation. The EICAS display is needed by the pilot during all aircraft operations, including start-up. For this reason, some components are powered when the battery switch is set to BATT or EMER: • BATT—Both PFDs, MFD, GIAs, GEAs • EMER—PFD1, GIA1, GEA1

NOTE Flight instruments, and all navigation, communications, and flight guidance functions are not enabled by the battery switch alone. The AVN PWR switch must also be powered on (up).

XM Weather Datalink The GDL69A is a remote-mounted satellite-broadcast receiver that receives XM weather for display on the MFD (and/or PFD inset map). The GDL 69A can receive XM weather and XM radio services. It communicates to the G1000 through the high-speed data bus. Figure 16-3.   Battery Switch

FOR TRAINING PURPOSES ONLY

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CITATION MUSTANG PILOT TRAINING MANUAL 16 AVIONICS

AVIONICS POWER SWITCH The avionics power switch is in the AVIONICS section of the left lower instrument panel and has two positions: AVN PWR and OFF (Figure 16-4). This switch energizes the avionics solid-state relays (SSRs) closed. Each relay provides DC power from a powered bus (left or right electrical bus or emergency bus), through a corresponding avionics bus, to power specific units. All units receive from an avionics bus, except those powered directly from the battery switch.

30 minutes. The amber light does not illuminate when aircraft power is charging the battery and providing power to the standby instruments. In the BATT TEST position, the condition of the backup battery is tested. Illumination of the green light beside the switch indicates proper battery charge. If the green light does not illuminate, the backup battery is not properly charged, and standby flight instruments may not operate with the loss of normal DC power.

PRIMARY FLIGHT DISPLAY DESCRIPTION The main control display units for the G1000 are the PFDs. The Citation Mustang has two PFDs, one in front of the pilot and one in front of the copilot (Figure 16-5).

Figure 16-4.   Avionics Power Switch and Standby Instruments Switch

STANDBY FLIGHT INSTRUMENTS SWITCH The standby flight instruments switch is in the AVIONICS section of the left lower instrument panel and has three positions: STBY INST, OFF, and BATT TEST (Figure 16-4). When the switch is in the STBY INST position, it provides DC power (from a backup battery) to the standby flight instruments on the upper center instrument panel: • Airspeed indicator • Attitude gyro • Altimeter

The PFDs are 10.4-inch color LCDs, which are designed for visibility in bright sunlight. (Brightness is manually or automatically variable.) A bezel around each PFD contains controls for operating the PFD.

Display Color-Coding Color-coding on the displays indicates common meanings to the pilot. These colors and their meanings are: • Cyan—Pilot adjustable • Green—Active • White—Armed/standby • Amber—Caution • Red—Warning • Magenta—GPS derived

A small amber light emitting diode (LED) on the right side of the switch illuminates when in the STBY INST position, indicating the standby flight instruments are discharging power from the dedicated standby lead-acid battery pack. A fully charged battery supplies power for approximately

16-10

FOR TRAINING PURPOSES ONLY

FOR TRAINING PURPOSES ONLY

RAM AIR TEMPERATURE BOX

HORIZONTAL SITUATION INDICATOR ISA DEVIATION TEMPERATURE BOX

HEADING BOX

AIRSPEED INDICATOR

NAV FREQUENCY WINDOW

SOFTKEYS

SLIP/SKID INDICATOR NAVIGATION STATUS BAR

Figure 16-5.   PFD Graphical Callouts

ATTITUDE INDICATOR COM FREQUENCY WINDOW

SYSTEM TIME BOX

16 AVIONICS

TRANSPONDER STATUS BAR

TURN RATE INDICATOR

BAROMETRIC SETTING BOX

VERTICAL SPEED INDICATOR

ALTIMETER

SELECTED ALTITUDE BOX

CITATION MUSTANG PILOT TRAINING MANUAL

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CITATION MUSTANG PILOT TRAINING MANUAL 16 AVIONICS

GRAPHICAL FLIGHT INSTRUMENTATION

Airspeed Pointer

The following flight instruments are depicted graphically on the PFD: • Airspeed display • Altitude • Attitude display including turn coordination (slip/skid) • Horizontal situation indicator (HSI) (heading, turn rate, and radio navigation) • Vertical speed

NOTE If the aircraft exceeds normal flight attitudes (30° nose-up or 20° nose-down pitch, or roll greater than 65° bank), the PFD declutters leaving the primary flight instruments.

Airspeed Display The airspeed is displayed as a vertical scale on the left side of the PFD (Figure 16-6). AIRSPEED TREND VECTOR

The airspeed (in KIAS) appears as a rolling numeric display (white numbers on black) in the center of the airspeed tape. This tape has a pointer, which always points at the center of the airspeed display. Below 20 KIAS, the digits change to dashes. Once the airspeed reaches and/or exceeds VMO or MMO, the pointer changes to red with white numbers.

Trend Vector On the outside right edge of the tape, a magenta trend vector indicates predicted airspeed in 6 seconds at the current rate of airspeed acceleration. When the trend vector enters the barber pole range, the numbers inside the airspeed pointer change from white to amber.

Speed Awareness Airspeeds above the maximum operating speed appear in the high-speed awareness range represented on the airspeed tape by a red/white “barber pole” (Figure 16-7) The flap speed references represent the airspeed limits when the aircraft is below 18,000 feet. A red low-speed awareness range extends up from the bottom of the tape to indicate VLSA.

VSPEED REFERENCES

ACTUAL SPEED

REFERENCE APPROACH SPEED CUE

Figure 16-6.   Airpseed Indicator

Figure 16-7.   Red Pointer

Moving Tape Display Box

Airspeed References

White numerals and tick marks on a transparent rolling-tape display box indicate airspeeds currently above and below the current airspeed. Each small tick mark represents 5 knots, and each large tick mark (with a number) represents 10 knots.

Airspeed references, or flags, appear on the outside right edge of the tape display fixed to their corresponding airspeed (Table 16-2). Takeoff and landing speed flags can be displayed and adjusted in the reference window using the TMR/REF softkey. Takeoff references are automatically turned off when airspeed reaches 160 kt (see Figure 16-6).

16-12

FOR TRAINING PURPOSES ONLY

16 AVIONICS

CITATION MUSTANG PILOT TRAINING MANUAL

Table 16-2.   V-SPEED TABLE Vspeed

Flag

V1

1

VR

R

V2

2

VENR

E

VAPR

AP

VREF

RF

AIRSPEED REFERENCE

A reference approach speed cue appears as an open green circle on the right edge of the tape display. The open green circle represents 1.3 VS1 (see Figure 16-6).

Flap Speed References Maximum flap-extension speeds are indicated on the outside right edge of the tape display by reference flags similar to the V-speed flag (Figure 16-8). Maximum speed for takeoff/approach flaps is indicated by a flag labeled “TA”. Maximum speed for landing flaps is indicated by a flag labeled “LD.” The flap reference speed flag cannot be turned off or adjusted by the pilot.

FLAP SPEED REFERENCES

Figure 16-8.   Flap Speed

AIRSPEED REFERENCE BUG

Figure 16-9.   AFCS Reference

Mach Window At the bottom of the airspeed tape display, a window indicates the Mach number, depending upon aircraft altitude or airspeed.

Attitude And Turn Coordination At the center of the PFD is the attitude indication (Figure 16-10). The amber symbol represents the airplane, and the amber bars represent the wingtips. The horizon line is white, with brown (ground) below and blue (sky) above. The horizon line indicates 0° pitch. Pitch lines are at 2.5° intervals above and below the horizon line and are labeled at 10° intervals. Red chevrons indicate extreme pitch at greater than 50° nose-up or 30° nose-down (Figure 16-11). Roll indication is provided by an arc at the top of the display that rotates with the horizon line. The arc has major tick marks at 30° and 60°, and minor tick marks at 10°, 20° and 45°.

Turn Coordination (Slip/Skid)

AFCS Selected Airspeed Box and Bug When the AFCS is set to a selected airspeed during flight level change mode, the selected airspeed appears in a box at the top of the display, and a corresponding notched blue bug appears on the inside right edge of the airspeed display (Figure 16-9).

Turn coordination is indicated by a small sliding bar (slider) under the attitude display roll pointer (Figure 16-12). The slider deflects a full width left or right of the roll pointer to indicate the same condition as indicated by a ball positioned one ball width to the left or right of center in a conventional instrument.

FOR TRAINING PURPOSES ONLY

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CITATION MUSTANG PILOT TRAINING MANUAL 16 AVIONICS

9

1

ROLL POINTER

2

ROLL SCALE

3

HORIZON LINE

4

AIRCRAFT SYMBOL

5

LAND REPRESENTATION

6

PITCH SCALE

7

SLIP/SKID INDICATOR

8

SKY REPRESENTATION

9

ROLL SCALE ZERO

8

1

20

20

10

10

7

2 6

3

4

10

10

5

Figure 16-10.   Attitude Indication COMMAND BARS

60

60

50

50

40

40

30

30

NOSE HIGH

10

10

AIRCRAFT SYMBOL 20

20

30

30 40

40 NOSE LOW

Figure 16-11.   Pitch Attitude Warnings

20 Figure 16-12.   Slip/Skid Indication

Flight Director Command Bars The single-cue command bars vertically move together to indicate pitch commands and bank left or right to indicate roll commands (Figure 16-13). Command bars that display as a cross pointer move independently to indicate pitch (horizontal bar) and roll (vertical bar) commands. Both PFDs show the same command bar format (Figure 16-14).

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Figure 16-13.   Flight Director Single-Cue Command Bars

Altimeter Moving Tape Display Box White numerals and tick marks on a transparent rolling-tape display box indicate altitudes currently above and below the current altitude. Each minor tick mark indicates 20 feet, and each major tick mark (with a number) indicates 100 feet.

Altitude Pointer The current altitude (indicated in feet above mean sea level) appears as a rolling numeric display (white numbers on black) in the center of the altitude display box (Figure 16-15). This display has a pointer, which always points at the center of the altitude tape.

Trend Vector On the inside left edge of the tape, a magenta trend vector indicates predicted altitude in 6 seconds at the current rate of altitude change.

FOR TRAINING PURPOSES ONLY

Reference Altitude Alerting When the aircraft is more than 1,000 feet above or below the reference altitude, the digits in the selected altitude box are cyan on black (see Figure 16-15). When the aircraft is at the reference altitude (±1,000 feet), the colors reverse (black digits on cyan). When approaching within 1,000 feet of the reference altitude, the digits and their background flash (alternate colors) continuously for 5 seconds. When departing the reference altitude by ±200 feet, the digits flash amber on black for 5 seconds (Figure 16-16).

COMMAND BARS

AIRCRAFT SYMBOL

Figure 16-14.   Flight Director Cross-Pointer Command Bars

SELECTED ALTITUDE BUG

SELECTED ALTITUDE

14 1500000

WITHIN 200 FT

DEVIATION OF ±200 FT

4000

4000

4000

Figure 16-16.   Altitude Alerting Display

Barometric Setting

1400 1300 ALTITUDE TREND VECTOR

WITHIN 1,000 FT

20 2 1 00 80

CURRENT ALTITUDE

A box at the bottom of the altitude tape indicates the current altimeter setting in inches of mercury (or hectoPascals, if set to metric values) (Figure 16-17). To adjust the setting, rotate the BARO knob on the right side of each PFD bezel. The BARO settings on the pilot and copilot PFDs can be synchronized through the PFD setup menu window.

1100 BAROMETRIC MINIMUM DESCENT ALTITUDE BUG

NOTE

1000 900 IN 30.09

BAROMETRIC SETTING

If pilot and copilot PFD barometric settings differ by more than .02 inches (of mercury), the barometric setting boxes on both displays appear with amber digits. The baro settings for both PFDs can be synchronized.

Figure 16-15.   Altimeter

Selected Altitude Bug and Box A box at the top of the altitude display indicates the preselected altitude. A corresponding notched cyan bug appears on the right edge of the altitude display at the selected altitude, or at the end of the display if the selected altitude is off scale. The reference altitude is selected with the ALT SEL knob on the AFCS controller.

Barometric Altitude Minimums and Alerting For altitude awareness, a barometric minimum descent altitude (MDA, or decision height) can be set (Figure 16-18). When active, the MDA is displayed in a box labeled “BARO MIN” to the lower left of the altimeter and on the altitude tape with a bug (once the altitude is within range of the tape). This altitude can be adjusted in the BARO MIN field in the timer/references window from 0 to 16,000 feet (in 10-foot increments when using the small FMS knob). The MDA is reset any time the power is cycled.

FOR TRAINING PURPOSES ONLY

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CITATION MUSTANG PILOT TRAINING MANUAL

CITATION MUSTANG PILOT TRAINING MANUAL 16 AVIONICS

WITHIN 2500 FT

WITHIN 100 FT

2300 30

00

2200 50 2100

BAROMETRIC MINIMUM BOX

2000 BARO MIN 2000FT

20 00 19 80

2080

22 40

BAROMETRIC MINIMUM BUG

ALTITUDE REACHED

29.92IN

BARO MIN 2000FT

2000

1900

1900

1800

1800 29.92IN

BARO MIN 2000FT

1700 29.92IN

Figure 16-17.   Barometric MDA Displays

BARO MIN

1200FT

ON

Figure 16-18.   Barometric MDA Altitude Alert Setting

Metric/Standard Units of Measure After pressing the PFD softkey, an ALT UNIT softkey appears. It allows the display of the digital altitude and barometric pressure indications in metric values: altitudes in meters (MT) and barometric pressure in hectoPascals (hPA). The altitudes appear in boxes above the normal digital displays, which continue to display in feet. The moving-tape display continues to display in feet. The barometric pressure box at the bottom of the display changes to hPA indications.

Turn Rate Indication When the aircraft is yawing or turning, a magenta arc extends from the center pointer, left or right (depending on the direction of turn), a few degrees around the outer edge of the compass rose, to indicate the heading the aircraft will reach in 6 seconds (up to 24°). Two tick marks on either side of the center pointer, above the compass rose, indicate turn rate. If the magenta arc extends to the second tick mark (at 18° from center), a standard-rate turn (360° in 2 minutes) is indicated. The first mark indicates a half-standard-rate turn. At rates greater than 4°/second, an arrowhead appears at the end of the magenta trend vector and the prediction is no longer valid.

Horizontal Situation Indication

Selected Heading Bug

Heading

A rotatable cyan heading bug appears on the compass card at the selected heading, and the heading is presented with cyan digits in the black HDG box, immediately left of the magnetic heading digital readout.

The compass card appears with the current aircraft magnetic heading under the white pointer at the top of the compass card (Figure 16-19). Letters (N–S–E–W) appear at the four cardinal points and numbers indicate degrees at 30° intervals. Major tick marks are at 10° increments and minor tick marks at 5°. A digital readout of the current magnetic heading appears in a black box with large white digits immediately above the compass card pointer (lubber). For additional course awareness, a crosstrack error (XTK) indication appears near the bottom of the HSI whenever using GPS as the primary navigation source and course deviation indicator (CDI) exceeds maximum deviation. 16-16

Radio Navigation Selected Course, Course Pointer and Course Deviation Indicator The HSI can display three sources of navigation NAV1, 2, or GPS. The CDI softkey cycles through the navigation sources. The selected course appears as a digital readout in the black CRS box, immediately to the right of the magnetic heading readout. The selected course also appears as a long arrow across the face of the compass card. This arrow is

FOR TRAINING PURPOSES ONLY

16 1

CRS

30

W 21

GPS

TERM

N

4

013°

13

33

24

3

14

287°

013°

HDG

2

15

16 AVIONICS

CITATION MUSTANG PILOT TRAINING MANUAL

5

12

6

7

OBS

15

12

8

1

TURN RATE INDICATOR

2

SELECTED HEADING

3

CURRENT TRACK BUG

4

LATERAL DEVIATION SCALE

5

NAVIGATION SOURCE

6

AIRCRAFT SYMBOL

10 9

E

6

3

S

11

7

COURSE DEVIATION INDICATOR

12

FLIGHT PHASE

8

ROTATING COMPASS ROSE

13

SELECTED COURSE

9

TO/FROM INDICATOR

14

TURN RATE/HEADING TREND VECTOR

10

COURSE POINTER

15

CURRENT HEADING

11

HEADING BUG

16

LUBBER LINE

Figure 16-19.   Horizontal Situation Indicator (HSI)

360°

283°

33 N

3

W

GPS

ENR

GPS

LOI

30 TERN

180°

158°

S

15

21 12

15 LOC1

S

VOR2

SUSP

30

33

N

33

N

3

E

12

15

S

21

Figure 16-20.   Navigation Sources

the course pointer. The center section of the course pointer is the CDI, which disappears when there is no signal for navigation (Figure 16-20).

The color and depiction of the course indicator (and digits in the CRS box) varies as selected by the CDI softkey: • NAV1—Green single arrow and green selected course digits • NAV2—Green double arrow and green selected course digits • GPS—Magenta single arrow and magenta selected course digits The course is selected by the CRS 1 (for pilot PFD) or CRS 2 (for copilot PFD) knobs on the AFCS controller. The navigation source for the CDIs on the pilot and copilot PFDs can be synchronized through the PFD setup window. However, if the CDIs are not synchronized and PFDs are set to the same VHF navigation sources, the source reference (NAV or LOC) appears in amber lettering. Refer to Table 16-3 for information on flight phases and CDI scaling.

FOR TRAINING PURPOSES ONLY

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CITATION MUSTANG PILOT TRAINING MANUAL 16 AVIONICS

Table 16-3.   Flight Phases and CDI Scaling FLIGHT PHASE

ANNUNCIATION

AUTO CDI SCALING

DEPARTURE

DPRT

0.3 NM

OCEANIC

OCN

2.0 NM

ENROUTE

ENR

2.0 NM

TERMINAL MODE

TERM

1.0 NM

APPROACH (NON-PRECISION)

LNAV

1.0 NM DECREASING TO 350 FT DEPENDING ON VARIABLES

APPROACH (VERTICAL GUIDANCE)

LNAV +V

1.0 NM DECREASING TO 350 FT DEPENDING ON VARIABLES

APPROACH (LNAV/VNAV)

L/V NAV

1.0 NM DECREASING TO 350 FT DEPENDING ON VARIABLES

APPROACH (LPV)

LPV

MISSED APPROACH

MAPR

1.0 NM DECREASING TO 350 FT DEPENDING ON VARIABLES 0.3 NM

Bearing Pointers and Information Windows Two bearing pointers can be selected to appear on the face of the compass card (press the PFD softkey and select BRG1 and/or BRG2). The BRG1 pointer is a single cyan line with an open arrow pointer and the BRG2 pointer is a double cyan line with an open arrow pointer. If a bearing pointer is selected for display, a white circle appears in the center of the compass card to separate the bearing pointer(s) from the CDI. Bearing pointers never override the CDI. The BRG1 window is at the lower left corner of the compass display and the BRG2 window is at the lower right corner of the compass display. Each window indicates: • Distance to the station/waypoint • Station/waypoint identifier (or frequency) • Source for the bearing (NAV1, NAV2, GPS, or automatic direction finder [ADF]) • Arrow icon matching the bearing pointer

OBS Mode OBS mode is useful for holding and for intercepting a course to a waypoint. Enabling OBS mode with the OBS softkey suspends the automatic sequencing of waypoints in a GPS flight plan, but

16-18

retains the active-to waypoint as the navigation reference even after passing the waypoint. When OBS is disabled by pressing the OBS softkey again, the GPS returns to normal operation, with automatic sequencing of waypoints. OBS mode also allows a desired course TO/ FROM a waypoint to be set (with a CRS knob); pressing the CRS knob recenters the CDI and returns the course point TO the waypoint bearing.

SUSPend Mode SUSPend mode is automatically activated when appropriate during approach operations. As the aircraft crosses the missed approach point (MAP), automatic approach waypoint sequencing is suspended. SUSP appears on the HSI (to the lower right of the aircraft symbol) in place of OBS and the OBS softkey label changes to SUSP. SUSP mode is also automatically actuated when on vectors to an approach or while in a hold.

Glide Slope During precision approach operations, glide-slope position is indicated by a diamond on a glide-slope scale, which is on the left edge of the altitude display (Table 16-4 and Figure 16-21).

FOR TRAINING PURPOSES ONLY

Table 16-4.   VERTICAL DEVIATION DISPLAY G

G

V

(GLIDESLOPE)

(GLIDEPATH)

(VERTICAL DEVIATION)

MARKER BEACON ANNUNCIATION

M

1800

G

1700

INDICATOR

OUTER MARKER

MIDDLE MARKER

INNER MARKER

O

M

I

BUG ICON

0

M

1800

ALTIMETER

G Figure 16-22.   Marker Beacon Indications

1600

4

40 15 20 00

GLIDE-SLOPE INDICATOR

0

2 VERTICAL SPEED POINTER

1400

500

1300 2

29.92IN 4

Figure 16-21.   Glide-Slope Indicator

Marker Beacon When passing over a marker beacon, a small box appears at the top left of the altitude display (Figure 16-22). The box appears differently for different beacons:

Figure 16-23.   Vertical Speed Indicator

Vertical Speed Reference Box and Bug

• O uter marker—Blue box with black letter “O”

A reference vertical speed set through the AFCS indicates on the vertical speed scale by a notched cyan bug on the scale and a vertical speed indicated with cyan digits in a box at the top of the scale.

• M iddle marker—Amber box with black letter “M”

INSET MAP

• I nner marker—White box with black letter “I”

The pilot may activate a small moving map or inset in the lower left corner of the PFD. The PFD inset map can contain much of the same information as available in the full-size moving map on the MFD.

Vertical Speed Vertical Speed Indication The vertical speed indication is through a fixed index scale on the right side of the altitude scale, and through a moving black pointer box with white digits indicating current vertical speed (when greater than 100 fpm up or down) (Figure 16-23). A negative number indicates descent. The vertical speed indication displays in 50-foot increments.

AUXILIARY INFORMATION WINDOW In the lower-right corner of the PFD, an auxiliary information window can be set to display information from the system, including:

FOR TRAINING PURPOSES ONLY

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16 AVIONICS

CITATION MUSTANG PILOT TRAINING MANUAL

CITATION MUSTANG PILOT TRAINING MANUAL 16 AVIONICS

• N earest airports, with basic information as currently recorded in the navigation database (NRST softkey) • ADF and distance measuring equipment (DME) tuning (ADF/DME softkey) • Flight plan display and entry (FPL key) • Garmin system messages (MSG softkey) • PDF setup (MENU key) • References (TMR/REF softkey) • Flight plan procedures (PROC key) • Direct-To (D Key) Also use this window to select a waypoint from the database and/or to display information about the waypoint.

REVERSIONARY MODE To provide emergency backup of the MFD and PFDs, and to provide flexibility in display modes, the G1000 has a reversionary mode that causes both the PFD and the MFD to display the essential information for continued flight. In the event of display failure, the display modes are as follows: • PFD1 failure—MFD enters reversionary mode; PFD2 remains in normal mode. • MFD failure—PFDI and PFD2 enter reversionary mode. • PFD2 failure—PFD1 and the MFD remain in normal mode. Reversionary mode can be manually selected for the onside display and the MFD by pressing the large red DISPLAY BACKUP button at the bottom of either audio control panel. Pressing the red DISPLAY BACKUP button a second time returns both displays to normal mode. In reversionary mode, the left edge of the display includes a single-column of EICAS information condensed from the normal two-column EICAS display on the MFD. Additionally, a CAS message window appears on the right edge of the PFD and expands as necessary to display any active CAS messages (up to 14 in a window) (Figure 16-24). If more than 14 messages appear, they can be scrolled.

16-20

HDG NO COMP ROL NO COMP PIT NO COMP ALT NO COMP

0

5700

4

5600 5500

2

BOTH ON GPS1 BOTH ON AHRS1 BOTH ON ADC2

COMPARATOR WINDOW REVERSIONARY SENSOR WINDOW

20

54 20 00

5300

2

5200

4

30.07IN

SYSTEM MESSAGES WINDOW

MESSAGES

GWX FAIL - GWX IS INOPERATIVE. GMX FAIL - GMX IS INOPERATIVE.

CNFG MODULE - PFD1 CONFIGURATION MODULE IS INOPERATIVE. XPDR1

IDENT

5537 IDNT TMR/REF

UTC

NRST

18:07:21 MSG

MESSAGES SOFTKEY

Figure 16-24.   G1000 System Messages

NAV/COM FREQUENCIES AND NAVIGATION DATA WINDOWS Across the top of the PFD are avionics settings and data indications, including VHF NAV and COM frequencies (active and standby), and navigation status data.

PFD CONTROLS Each PFD is controlled by knobs and pushbutton keys on its bezel, and/or by pilot operation of the AFCS or MFD/FMS controller. The following general discussion of controls is also applicable to the other controllers.

Controls Garmin displays (PFD and MFD) and control panels (audio panel, AFCS controller, and FMS controller) have various types of knobs, buttons and switches. Refer to Figure 16-25 for PFD controls locations and Table 16-5 for a description on those controls.

FOR TRAINING PURPOSES ONLY

Figure 16-25.   Primary Flight Display

10

11

6

3

9

7

2

FOR TRAINING PURPOSES ONLY

16 AVIONICS

12

15

14

13

4

5

8

1

CITATION MUSTANG PILOT TRAINING MANUAL

16-21

16-22 Sets the altimeter barometric pressure. Press to enter standard pressure (29.92) or return to the previous setting. Tunes the standby frequencies for the COM transceiver (large knob for MHz; small knob for kHz). Press to switch the tuning box (cyan box) between COM1 and COM2. Toggles the standby and active COM frequencies. Press and hold this key for two seconds to tune the emergency frequency (121.5 MHz) automatically into the active frequency field. Controls COM audio volume level. Press to turn the COM automatic squelch ON and OFF. Volume level is shown in the COM frequency field as a percentage.

Erases information, cancels entries, or removes page menus. Flight management system knob. Press the FMS knob to turn the selection cursor ON and OFF. When the cursor is ON, data may be entered in the applicable window by turning the small and large knobs. The large knob moves the cursor on the page, while the small knob selects individual characters for the highlighted cursor location.

Dual NAV Knob

Joystick

BARO Knob

Dual COM Knob

COM Frequency Transfer Key

COM VOL/SQ Knob

Direct-to Key

FPL Key

CLR Key

Dual FMS Knob

MENU Key

PROC Key

ENT Key

3

4

5

6

7

8

9

10

11

12

13

FOR TRAINING PURPOSES ONLY

14

15

Validates or confirms a menu selection or data entry.

Gives access to IFR departure procedures (DPs), arrival procedures (STARs) and approach procedures (IAPs) for a flight plan. If a flight plan is used, available procedures for the departure and/or arrival airport are automatically suggested. These procedures can then be loaded into the active flight plan. If a flight plan is not used, both the desired airport and the desired procedure may be selected.

Displays a context-sensitive list of options. This list allows the user to access additional features or make setting changes that relate to particular pages.

Displays the active flight plan page for creating and editing the active flight plan.

Allows the user to enter a destination waypoint and establish a direct course to the selected destination (the destination is either specified by the identifier, chosen from the active route, or taken from the map pointer position).

Changes the map range when rotated. Activates the map pointer for pan when pressed.

Tunes the standby frequencies for the NAV receiver (large knob for MHz; small knob for kHz). Press to switch the tuning box (cyan box) between NAV1 and NAV2.

Toggles the standby and active NAV frequencies.

NAV Frequency Transfer Key

2

Controls NAV audio volume level. Press to toggle the Morse code identifier audio ON and OFF. Volume level is shown in the NAV frequency field as a percentage.

DESCRIPTION

NAV VOL/ID Knob

CONTROL

1

#

16 AVIONICS

Table 16-5.   PRIMARY FLIGHT DISPLAY CONTROL DESCRIPTIONS

CITATION MUSTANG PILOT TRAINING MANUAL

FOR TRAINING PURPOSES ONLY DCLTR

OFF

TRAFFIC

TOPO

AHRS1

OFF

VERSION .10 AND PRIOR

TERRAIN STRMSCP NEXRAD

AHRS2

OPTN3

BRG2 (OFF)

BRG1 (OFF)

XM LTNG

IN

BRG2 (ADF)

BRG1 (ADF)

METERS

BRG2 (GPS)

BRG1 (GPS)

IDENT

IDENT

MSG

BACK

MSG

MSG

MSG MSG

BACK BACK

LEGEND

MSG

BACK

IF MESSAGES REMAIN AFTER ACKNOWLEDGEMENT, THE MSG SOFTKEY IS BLACK ON WHITE.

ON

OFF

16 AVIONICS

SELECT THE STD BARD OR OFF SOFTKEY TO RETURN TO THE TOP-LEVEL SOFTKEYS.

HPA

BACK

DISPLAY MSG WINDOW

MSG

NOTES: TOP LEVEL SOFTKEY NAMES ARE DISPLAYED.

MSG

BACK

BACK

DISPLAY NEAREST AIRPORT WINDOW

BACK

BKSP

DISPLAY TIMER/ REFERENCES WINDOW

ALT UNIT STD BARO

7

CODE

BRG2 (NAV1)

6

VFR

BRG1 (NAV1)

5

GND

SPECIAL AIRCRAFT POSITION IDENTIFICATION

BEZEL-MOUNTED SOFTKEYS (PRESS)

BRG2

4

ALT

DME TUNING WINDOW (OPTIONAL)

BRG1

3

ON

OPTN2

DME

X POINTR

OPTN1

WIND

2

STBY

CDI (GPS)

CDI (NAV2)

CDI (NAV1)

SWITCH NAVIGATION SOURCES FOR CDI

Figure 16-26.   Softkey Chart for Garmin Software (Sheet 1 of 2)

DCLTR-3

DCLTR-2

DCLTR-1

ADC2

SNGL CUE

DFLTS

1

0

ADC1

FD FRMT

XPDR2

XPDR1

OBS MODE WHEN NAVIGATING WITH GPS

CITATION MUSTANG PILOT TRAINING MANUAL

16-23

16-24

FOR TRAINING PURPOSES ONLY OPTN1

WIND

DCLTR

OFF

TRAFFIC

TOPO

AHRS1

VERSION .16

TERRAIN STRMSCP NEXRAD

AHRS2

OFF

XM LTNG

MSG

MSG

MSG

BACK

DISPLAY MSG WINDOW

MSG

MSG

MSG

BACK

LEGEND

MSG

MSG BACK

BACK

IF MESSAGES REMAIN AFTER ACKNOWLEDGEMENT, THE MSG SOFTKEY IS BLACK ON WHITE.

ON

OFF

SELECT THE STD BARD OR OFF SOFTKEY TO RETURN TO THE TOP-LEVEL SOFTKEYS.

HPA

BACK

BACK

BACK

BACK

DISPLAY NEAREST AIRPORT WINDOW

NOTES: TOP LEVEL SOFTKEY NAMES ARE DISPLAYED.

Figure 16-26. Softkey Chart for Garmin Software (Sheet 2 of 2)

DCLTR-3

DCLTR-2

DCLTR-1

ADC2

OPTN3

BRG2 (OFF)

BRG1 (OFF)

IN

BRG2 (ADF)

BRG1 (ADF)

METERS

BRG2 (GPS)

IDENT

IDENT BKSP

DISPLAY TIMER/ REFERENCES WINDOW

ALT UNIT STD BARO

7

CODE

BRG1 (GPS)

6

VFR

BRG2 (NAV1)

5

GND

BRG1 (NAV1)

4

ALT

SPECIAL AIRCRAFT POSITION IDENTIFICATION

BEZEL-MOUNTED SOFTKEYS (PRESS)

BRG2

3

ON

DME TUNING WINDOW (OPTIONAL)

BRG1

OPTN2

DME

2

STBY

SYNTERR HRZN HDG APTSIGNS

DFLTS

1

0

ADC1

SYN VIS

XPDR2

XPDR1

CDI (GPS)

CDI (NAV2)

CDI (NAV1)

SWITCH NAVIGATION SOURCES FOR CDI

16 AVIONICS OBS MODE WHEN NAVIGATING WITH GPS

CITATION MUSTANG PILOT TRAINING MANUAL

Softkeys Along the bottom of the PFD, the bezel contains 12 “softkeys” marked with upward-pointing triangles. These keys do not have a single, specific, permanent function. They have different purposes at different times, as determined by the G1000 software. Some (or all) softkeys have labels appearing immediately above them on the display. The labels change depending upon pilot settings and/or current conditions. Navigating the lower level menus is done through the top level softkey menu. When the label for a specific feature is toggled off, the text is white on a black background. When the label is toggled on, the text is black on a light gray background. When the BACK softkey is available (on the right end of the softkeys), this key can be pressed to escape the current menu and return to the previous menu display. For details on the menus for PFD softkeys refer to Figure 16-26.

TRANSPONDER DISPLAY AND CONTROL The PFD depicts current transponder status in the black XPDR box on the bottom right side of the PFD. With digits and letters, the box indicates current transponder code and mode of operation: • STBY—Transponder in standby mode • ON—Normal mode (mode A) operating • A LT—Altitude encoding (mode C) operating • IDENT • GND—Ground To indicate the transponder is replying to interrogations from radar or other sources, a small white letter “R” appears at the right end of the XPDR box. The transponder code and mode are set by the pilot, through the softkey XPDR menu (appears when pressing the XPDR softkey). When in the XPDR softkey menu, the pilot selects modes (as listed above) by pressing the corresponding mode softkey. Pressing the VFR softkey automatically selects the appropriate VFR country code as preset at the factory.

Revision 1.0

Pressing the CODE softkey causes digits 0–7 to appear above the softkeys. Enter the transponder code. To change the code before confirming it, press the BKSP softkey, which backs the cursor through the code (erasing a digit at a time), and reenter the erased digits. Setting the code in active transponder also sets the same code in the inactive transponder. Five seconds after the fourth digit is entered, the transponder code becomes active. Pressing BACK returns the pilot to the previous level of softkeys. The IDENT key is always visible from the main softkey menu or the XPDR menu, and can be pressed at any time it is visible. Pressing the IDENT key always makes the transponder squawk IDENT, then always returns the softkeys to the main menu. IDENT can also be selected from the control wheel.

SUPPLEMENTAL FLIGHT DATA SENSOR Source Selection Each PFD normally presents data from the sensors associated with its respective-side ADC and AHRS. To select cross-side ADC or AHRS, use the SENSOR softkey on the main softkeys menu. In flight only, if a sensor fails, automatic conversion occurs after 2 seconds.

DME Information Window The DME information window appears to the left of the HSI directly above the BRG 1 information (Figure 16-27).

Wind Data Box The wind data box appears above the DME window. There are three wind presentations, which indicate the calculated wind velocity. When the window is selected for display, but window information is invalid or unavailable, the window displays NO WIND DATA (Figure 16-28).

FOR TRAINING PURPOSES ONLY

16-25

16 AVIONICS

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CITATION MUSTANG PILOT TRAINING MANUAL

DME INFORMATION WINDOW

TUNING MODE FREQUENCY

–.–– NM

HDG

BEARING 1 POINTER

178°

223°

15

E

6 3

BEARING SOURCE

DME NAV2 113.00 9.9 NM 39.3NM TOP NAV1

VOR1

BEARING 1 INFORMATION WINDOW

POINTER ICON

NO WAYPOINT SELECTED

30

54.6NM BUM NAV1

STATION IDENTIFIER

BEARING 2 POINTER

21

W

GPS DISTANCE TO BEARING SOURCE

356°

CRS

24

12

DME DISTANCE

S

N

DME LABEL

33

NO DATA GPS

9.9NM OJC NAV2

BEARING 2 INFORMATION WINDOW

POINTER ICON

BEARING SOURCE

Figure 16-27.   HSI with Bearing and DME Indication NORMAL DISPLAY 130

M120 .411

305

035°

30

W

OPTION 2

33

24

N

OPTION 1

HDG

RAT

GPS

21

3

0°C

S

6

RAT

0°C ISA +15°C

15

Figure 16-28.   Wind Data Box

REVERSIONARY MODE 130

CABIN PRESS ALT FT 7500 5.0 DIFF PSI FLAPS

The AFCS presents alerts in the system status field near the top of the airspeed scale.

Air Temperatures (RAT And ISA) The ram air temperature (RAT) display is at the lower-left corner of the PFD display area (Figure 16-29). The ISA (International Standard Atmosphere) temperature display is immediately right of the RAT box. The ISA temperature box indicates the difference between ISA temperature and the OAT.

E GPS

UP

System Time Box A box in the lower-right corner of the PFD indicates time as referenced by satellite data (Figure 16-30). The time is displayed as local time (LCL) in 12-hour or 24-hour format, or as Coordinated Universal Time (UTC).

Generic Timer A generic timer may be set and operated by the pilot (Figure 16-31). It displays hours/minutes/ seconds, counting up or counting down. Access is through the TMR/REF softkey.

FOR TRAINING PURPOSES ONLY

S

W

27 DC VOLTS 27 200 DC AMPS 200 BATT VOLTS 27

15

12

6

750 1000

145

035°

3

750 1000

RAT 0°C ISA +15°C

HDG

24

ISA +15°C

125 60

120

Figure 16-29.   Ram Air Temperature Box

AFCS Alerts

16-26

125 60

142.8

21

When the TIS detects a traffic hazard, an amber/ black TRAFFIC box appears to the right of the top of the airspeed scale. If a traffic alert occurs, the inset map scales to the appropriate range. If the range on the map is set appropriately, an amber symbol appears on the map with a track vector.

N2% OIL PSI °C FUEL LBS PPH 0 °C

N

142.8

33

Traffic Alerts

30

OFF

12

OPTION 3

E

16 AVIONICS

DME NAV1 117.95

14900

117° 30

W

CRS

300°

The MFD has many possible settings and displays. For additional details, refer to the supplemental G1000 manuals and documents supplied with the Citation Mustang.

4

30.04IN

33

N

24

TERM

17:41:40 21

3

LCL

S

6

MFD CONTROLS

E XPDR1

1259 ALT

R LCL

17:12:20

12

15

Figure 16-30.   System Time Box

REFERENCES TIMER

01:04:45

UP

STOP?

The MFD is controlled by softkeys across the bottom of the MFD bezel, and by the MFD/FMS controller on the cockpit pedestal, below the throttle quadrant (Figure 16-33, Tables 16-6 and 16-7).

MFD Softkeys

Figure 16-31.   Timer Reference Window

MFD softkey functions are variable depending on user inputs. Each softkey has a label above it, which indicates its current function.

MULTIFUNCTION DISPLAY

MFD/FMS Controller

DESCRIPTION

Dual FMS Knob

This section provides a basic discussion of MFD displays and controls, and is intended to provide the pilot a useful understanding of the purpose and organization of the MFD (Figure 16-32).

The dual FMS knob on the MFD/FMS controller is the main control for selecting most MFD functions. The large knob selects page groups, while the small knob selects individual pages within the selected group.

The MFD/FMS controller includes an alphanumeric keypad for direct entry of numbers and text.

The G1000 includes a single 15-inch MFD in the center of the instrument panel. The MFD provides indications for: • EICAS • Moving-map displays (with depictions of navigation references and flight hazards) • Information pages on waypoints, instrument flight rule (IFR) procedures, airports, airways, and navaids • Flight planning • Navigation status indications • System status indications The MFD also provides an alternate display for essential flight instrumentation from either PFD through the use of reversionary mode.

For data entry, pressing the dual FMS knob activates the cursor, which allows the user to input new data. While the cursor is active, rotating the large FMS knob allows the user to move to different data entry locations on the display, as indicated by cursor movement. While the cursor is active, rotating the small FMS knob allows the user to change the text or data at the cursor location. Rotating the FMS knob causes each possible character to display sequentially. When the correct entry at the cursor location is selected with the small FMS knob, confirm by pressing the ENT key. It moves to the next position when you press ENT.

FOR TRAINING PURPOSES ONLY

16-27

16 AVIONICS

CITATION MUSTANG PILOT TRAINING MANUAL

CITATION MUSTANG PILOT TRAINING MANUAL

Figure 16-32.   Multifunction Display

16 AVIONICS

16-28

FOR TRAINING PURPOSES ONLY

Revision 1.1

Figure 16-33.   MFD Controller

14

1

13

3

2

12

4

11

5

10

9

8

7

6

16 AVIONICS

CITATION MUSTANG PILOT TRAINING MANUAL

FOR TRAINING PURPOSES ONLY

16-29

CITATION MUSTANG PILOT TRAINING MANUAL 16 AVIONICS

Table 16-6.   MFD SOFTKEYS MFD SOFTKEYS

FUNCTION

CAS

Scroll up (enabled only when a sufficient number of items are displayed in the crew alerting system display to warrant scrolling).

CAS

Scroll down (enabled only when a sufficient number of items are displayed in the crew alerting system display to warrant scrolling). Enables second-level Navigation Map softkeys

MAP TRAFFIC

Displays traffic information on Navigation Map

PROFILE

Displays/removes Profile View on Navigation Map Page Displays topographical data (e.g., coastlines, terrain, rivers, lakes) and elevation scale on Navigation Map.

TOPO TERRAIN

Displays terrain information on Navigation Map (not available with TAWS-A)

AIRWAYS

AIRWAYS: Displays airways on the map when next level softkeys are pressed (default label is dependent on map setup option selected); cycles through the following:

AIRWY ON

AIRWY ON: All airways are displayed

AIRWY LO

AIRWY LO: Only low altitude airways are displayed

AIRWY HI

AIRWY HI: Only high altitude airways are displayed (Default label is dependant on map setup option selected)

NEXRAD

Displays NEXRAD weather and coverage information on Navigation Map (optional feature)

XM LTNG

Displays XM WX lightning information on Navigation Map (optional feature)

METAR

Displays METAR flaps on airport symbols shown on the Navigation Map

LEGEND

Displays the legend for the selected weather products. Available only when NEXRAD, XM LTNG, METAR and/or PROFILE softkeys are selected.

BACK

Selects desired amount of map detail; cycles through declutter levels:

DCLTR DCLTR-1

DCLTR-1: Declutters land data

DCLTR-2

DCLTR-2: Declutters land and SUA data

DCLTR-3

DCLTR-3: Removes everything except the active flight plan

SW CHRT

16-30

Returns to top-level softkeys

When available, displays optional airport and terminal procedure charts

FOR TRAINING PURPOSES ONLY

Revision 1.1

Direct-to Key

FPL Key

MENU Key

PROC Key

Joystick

Alphanumeric Keys

Plus (+) Minus (-) Key

Decimal Key

SEL Key

ENT Key

CLR Key

SPC Key

BKSP Key

2

3

4

5

6

7

8

9

10

11

12

13

14

CONTROL

Dual FMS Knob

1

#

FOR TRAINING PURPOSES ONLY Moves the cursor back one character space.

Adds a space character.

16 AVIONICS

Erases information, cancels entries, or removes page menus. Pressing and holding this key displays the navigation map page.

Validates or confirms a menu selection or data entry.

The center of this key activates the selected softkey, while the right and left arrows move the softkey selection box to the right and left, respectively.

Enters a decimal point.

Toggles a (+) or (-) character.

Allow entry of airports, waypoints, etc.

Changes the map range when rotated. Activates the map pointer when pressed. Pans the map or cursor when moved.

Gives access to IFR departure procedures (DPs), arrival procedures (STARs) and approach procedures (IAPs) for a flight plan. If a flight plan is used, available procedures for the departure and/or arrival airport are automatically suggested. These procedures can then be loaded into the active flight plan. If a flight plan is not used, both the desired airport and the desired procedure may be selected.

Displays a context-sensitive list of options. This list allows the user to access additional features or make setting changes that relate to particular pages.

Displays the active flight plan page for creating and editing the active flight plan, or for accessing stored flight plans.

Allows the user to enter a destination waypoint and establish a direct course to the selected destination (the destination is either specified by the identifier, chosen from the active route, or taken from the map pointer position).

Flight management system knob. This knob selects the MFD page to be viewed; the large knob selects a page group (MAP, WPT, AUX, NRST), while the small knob selects a specific page within the page group. Pressing the FMS Knob turns the selection cursor ON and OFF. When the cursor is ON, data may be entered in the applicable window by turning the small and large knobs. In this case, the large knob moves the cursor on the page, while the small knob selects individual characters for the highlighted cursor location.

DESCRIPTION

Table 16-7.   MFD CONTROLLER DESCRIPTIONS

CITATION MUSTANG PILOT TRAINING MANUAL

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CITATION MUSTANG PILOT TRAINING MANUAL 16 AVIONICS

ENGINE INDICATING AND CREW ALERTING SYSTEM The left side of the MFD displays the two-column EICAS display. The EICAS display includes the engine indicating system (EIS) and the CAS.

Engine Indicating System The EIS is a group of graphical and digital indications of the condition and performance of the aircraft system. These include indications for: • Engines and oil system • Fuel system • Electrical system • Pressurization • Rudder and aileron trim • Flap position

Crew Alerting System The CAS is described in detail in Chapter 4—“Master Warning,” along with a summary explanation of each CAS message.

MAIN MFD DISPLAY AREA The main MFD area presents different pages from four main page groups: • MAP—Moving-map displays • WPT—Waypoint information pages • AUX—Auxiliary information pages • NRST—Nearest facilities The selected display page group is noted by a cyan abbreviation (MAP, WPT, AUX, or NRST) in the lower right corner of the main display area. Each page group contains more than one page, indicated by the group of boxes to the right of the page groups list. To select a page group, use the large knob of the dual FMS knob on the MFD/FMS controller. Rotating clockwise selects a page group farther right on the list. The page being depicted is represented by a cyan box among the other boxes. The top right corner always indicates the page group and page name.

16-32

MAP Page Group This page group displays various map presentations (Figure 16-34), which include: Standard maps: • NAVIGATION—Displays navigation map with user-selected overlays (with variations including TRK UP, DTK UP, and NORTH UP). • TRAFFIC—Displays only traffic, as reported by TIS or the optional traffic advisory system (TAS). • WEATHER RADAR—Displays precipitation intensity. • WEATHER DATA LINK—Displays any weather depiction or data available from XM Weather satellite downlink. Combinations of display overlay can be selected (subscription required). • TERRAIN PROXIMITY—Displays only terrain/obstruction hazards, as indicated in terrain database.

MAP Softkey Menu When the NAVIGATION MAP is selected, the MAP softkey is normally visible on the bottom-left of the MFD. Pressing the MAP softkey brings up secondary level softkeys, which provide the option to select/deselect various combinations of overlays for the navigation map display, including: • TRAFFIC—Shows traffic indications (visible with any other overlay) • PROFILE - Winds Aloft data inside the Profile View is enabled by default when the Profile View is displayed on the Navigation Map. • TOPO—Shows topographic shading (not visible with NEXRAD) • TERRAIN—Shows terrain and obstruction hazards, color coded by proximity to aircraft altitude (not visible with NEXRAD) • AIRWAYS—Shows low and/or high altitude airways • NEXRAD—Shows NEXRAD downloaded from optional XM weather satellite broadcast (also visible with TRAFFIC and/or XM LTNG)

FOR TRAINING PURPOSES ONLY

Revision 1.1

NAVIGATION STATUS WINDOW

MINIMUM SAFE ALTITUDE

GROUND SPEED (CURRENT)

16 AVIONICS

CITATION MUSTANG PILOT TRAINING MANUAL

TRACK ANGLE

ESTIMATED TIME ENROUTE

CURRENT MFD PAGE GROUP

CURRENT MAP PAGE TITLE

MAP ORIENTATION

“TERRAIN DISPLAY ENABLED” ICON (NOTE)

MAP RANGE SETTING

TERRAIN MAPPING LEGEND (PROXIMITY TO AIRCRAFT ALTITUDE)*

TERRAIN HAZARDS 100–1,000 FEET BELOW AIRCRAFT WITHIN 100 FEET OF AIRCRAFT ALTITUDE OR CLOSER

MAP SOFTKEY (TO SELECT MAP SOFTKEYS) SOFTKEY LEGEND INDICATING CURRENT FUNCTION OF CORRESPONDING SOFTKEY (“DECLUTTER” KEY SHOWN HERE)

MFD PAGE GROUP MAP = MAP GROUP WPT = WAYPOINT INFO AUX = AUXILIARY INFO NRST = NEAREST AIRPORTS

MFD PAGE (SPECIFIC PAGE IN CURRENT PAGE GROUP)

NOTE—DISPLAY-ENABLED ICONS TERRAIN

XM LIGHTNING

TRAFFIC INFORMATION SYSTEM (TIS)

STORMSCOPE (OPTIONAL)

NEXRAD RADAR IMAGES

Figure 16-34.   MFD with MAP Displays

FOR TRAINING PURPOSES ONLY

16-33

CITATION MUSTANG PILOT TRAINING MANUAL 16 AVIONICS

• X M LTNG—Shows lightning strikes as downloaded from optional XM weather satellite broadcast • BACK—Returns to the top level softkey menu If the pilot has selected a specific GPS course to navigate (using the Direct-To key, the flight plan features, automatic flight control, or other methods), the NAVIGATION MAP, and other MFD map displays, depict the selected course by a magenta line from the aircraft to the currently selected destination or next scheduled waypoint.

Map Display Options Range The RANGE/PAN joystick knob on the MFD/FMS controller allows the user to zoom the map scale. To increase the map range, rotate the joystick knob clockwise. To decrease the map range, rotate the joystick knob counterclockwise.

Pan Moving maps can also be panned (moved off-center to view other areas away from the aircraft). To pan, press the RANGE/ PAN joystick knob. This causes an arrow to appear over the aircraft depiction. Pan the map by moving the joystick in the desired direction. As the arrow nears the edge of the display, it becomes stationary and the map moves (pans). Note that a MAP POINTER information box appears at the top of the map showing: • Distance and bearing of the pointer from the aircraft • Elevation at the pointer • Current latitude and longitude of the pointer If no further joystick input is provided by the pilot for 60 seconds, the map automatically returns to center on the current aircraft position. Pressing the joystick again deselects the panning arrow and returns the map to the aircraft position.

WPT Page Group The WPT page group depicts information about specific waypoints. Any of the waypoints found in the WPT page group can be entered into a flight plan or used for direct-to navigation.

16-34

The following are the pages found within the WPT page group: • Airport • Intersection • NDB • VOR • User WPT

Selecting Waypoints The user selects the waypoint of interest by entering its identifier, name, or location. As the selection is being made, the database automatically displays information for the first waypoint in the database that matches the selection criteria. The actual selection may not appear until the user has entered all of the selection criteria. Verify the correct waypoint page is displayed before using the information from that page.

Automatic Frequency Entry Some waypoint information pages display associated frequencies. To automatically enter one of these frequencies into a standby NAV or COM, select the frequency with the dual FMS knob, then press the ENT key. The selected frequency does not change to the active until the user presses the COM frequency transfer key.

AUX Page Group The AUX page group provides auxiliary information and data entry pages for the pilot. These include: • WEIGHT PLANNING • TRIP PLANNING • UTILITY • GPS STATUS • SYSTEM SETUP • XM INFORMATION • SYSTEM STATUS • DIAGNOSTICS (available on ground only for maintenance purposes) For details on the functions of each of these pages, refer to the supplemental G1000 documents supplied with the aircraft.

FOR TRAINING PURPOSES ONLY

NRST Page Group The NRST page group displays a moving map showing aircraft position relative to the following: • AIRPORTS • INTERSECTIONS • NDB • VOR • USER WPTS • FREQUENCIES • AIRSPACES The moving map initially displays the course to the nearest resource as a dotted/dashed white line.

FPL Page Group The FPL page group includes these pages: • A CTIVE FLIGHT PLAN—Including vertical navigation. • F LIGHT PLAN CATALOG—Allows the pilot to store several flight plans for future use.

NAVIGATION STATUS BOX The navigation data bar is the center area across the top of the MFD. It displays four fields of real time, GPS derived, navigation data. These fields can be pilot selected from a wide range of options found on the AUX-SYSTEM SETUP page under the MFD DATA BAR FIELDS box. Listed below are some examples of the information that can be displayed. • BRG—Bearing (to the active waypoint) • DIS—Distance (to the active waypoint) • DTK—Desired track (to the active waypoint) • E TA—Estimated time of arrival (to the active waypoint) Refer to the appropriate Garmin manual for a listing of options available for the software version being used.

FLIGHT MANAGEMENT SYSTEM WEIGHT PLANNING Before flight, when the G1000 initializes, the AUX–WEIGHT PLANNING page appears. The pilot uses the dual FMS knob to move through the windows of the page, and enter current data for the flight. Pressing the EMPTY WT softkey sets the cursor to that entry, and the pilot corrects as appropriate, then enters data for other weights. Fuel weights may be automatically entered from the current EICAS indications by pressing the FOB SYNC softkey. During aircraft operation, fuel flow and ETE can be automatically determined from actual operation, and the blank calculated fields display corresponding calculated data.

FMS AND FLIGHT PLANS A FMS provides for flight navigation planning and enroute status monitoring. The FMS primarily operates through the flight plan pages of the MFD, which allow the pilot to enter a flight plan with an entire flight profile, from takeoff to landing. Both lateral and vertical navigation courses may be entered into the flight plan. The flight plan may be used to guide the flight by the: • Pilot viewing the FPL and MAP pages for general reference • Pilot using the FD to follow the flight plan • Aircraft using the AP to follow the flight plan Flight plans changed or terminated in flight may be stored for future use and may be deleted. Computers for these functions are in the displays. GPS navigation data (provided from the GIAs) is the primary navigation source that is automatically selected flight plan status tracking. Any one of the displays can perform the FMS functions. Airspeed and altitude information from the ADC may also be used, depending on flight plan configuration.

FOR TRAINING PURPOSES ONLY

16-35

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CITATION MUSTANG PILOT TRAINING MANUAL 16 AVIONICS

MFD/FMS FLIGHT PLAN CONTROLS AND INDICATIONS

These waypoints may be selected, as needed, through the FPL (flight plan) pages.

The MFD/FMS controller provides primary control of these functions. PFDs may also be set to display smaller windows containing FMS information and maps. Corresponding controls on the PFD bezels may be used to control the FMS.

IFR PROCEDURES

NOTE Most explanations in this section are based on the assumption the user is using the MFD and the MFD/FMS controller. Some (but not all) FMS operations can also be performed with a PFD and its controls. This section provides limited operating information for training purposes only. For more detailed and current instructions, refer to the AFM, the Garmin G1000 Pilot’s Guide, and the Garmin G1000 Cockpit Reference Guide. Flight plans and their associated map and information pages are normally presented on the MFD. The FPL (flight plan) page group includes two types of pages: • A CTIVE FLIGHT PLAN—Flight plan currently active in the G1000, including VNAV • F LIGHT PLAN CATALOG—List of all flight plans stored in the G1000 Both FPL pages may be viewed on the MFD. Courses and waypoints in an active flight plan are depicted on the MAP–NAVIGATION MAP page (and other maps) of the MFD. The course for the active leg of an active flight plan is displayed as a magenta line. Other legs of the course display as white lines.

USER-DEFINED WAYPOINTS The user may define new waypoints for use in navigation. This is helpful in FMS operations to allow maximum flexibility in flight planning. Userdefined waypoints are necessary when a waypoint is not stored in the navigation database. The pilot enters user-defined waypoints into the WPT–USER WPT INFORMATION page. 16-36

IFR procedures are stored in the database and can be included in flight plans. The procedures include: • Standard instrument departures (SIDs) • Arrival procedures (standard terminal arrival routes [STARs]) • Approach procedures

VERTICAL NAVIGATION The MFD/FMS controller provides for vertical navigation (VNAV) capability in-flight planning (Figure 16-35). This feature is available for enroute/ terminal cruise and descent operations when VNAV has been enabled and a VNAV flight plan (with at least one vertical waypoint) has been activated. Vertical DIRECT TO also provides VNAV functions. The flight director/AFCS may be armed for VNAV at any time; however, no target altitudes are captured during a climb. VNAV indications include the following: • V NAV TARGET ALTITUDE—As preset by the pilot, that altitude in MSL to be reached by flying a programmed vertical speed. • V ERTICAL DEVIATION INDICATOR— This symbol indicates deviations from the correct flight path; thereby allowing the pilot (or AFCS) to make adjustments in pitch, as required, to maintain the selected vertical profile. • REQUIRED VERTICAL SPEED—The vertical speed rate required to achieve the preselected altitude within that distance defined by the programmed flight plan. • For more information on vertical navigation, refer to the Garmin manuals and guides supplied with your Citation Mustang.

FOR TRAINING PURPOSES ONLY

0 V

13400 13300

VERTICAL DEVIATION INDICATOR

00 1 13 80 20

13200

3600

VNAV TARGET ALTITUDE

4 2

-1250

REQUIRED VERTICAL SPEED

2

AFCS indications appear at the top of the PFD below the navigation status box (Figure 16-36). This status box contains information about the current status of the AFCS, including whether FD, AP, or YD are active, which guidance mode the AFCS is currently following, and which target values are being observed. ROL

13000 12900 29.92IN

Various controls on the yokes and throttles will also disengage the AP. FD can be selected using the FD key.

AP YD PIT

ALTS

Figure 16-36.   Automatic Autopilot

4

Flight Director

Figure 16-35.   VNAV Indications

AUTOMATIC FLIGHT CONTROL SYSTEM DESCRIPTION The Mustang includes an AFCS that provides flight guidance and automatic flight control. The automatic flight control system includes three primary functions: • Flight director • Autopilot • Yaw damper • Manual electric pitch trim The AFCS controller (above the MFD) provides control of these functions, and the PFDs provide necessary indications (Figure 16-37 and Table 16-8). Computers for these functions are in the GIAs and the servos. If either GIA fails, the other GIA performs the FD functions. If both GIAs fail, the AFCS is not operational. The FD and AP follow the same sources of data for AFCS guidance. If AP is engaged, the information is used to control the airplane. If the FD is engaged, the information is used to guide the pilot. The pilot commands AP engagement/disengagement with the AP key on the AFCS controller.

The AFCS FD function causes magenta command bars to appear on the PFDs, which indicates the attitude required to correctly navigate the selected horizontal and/or vertical flight path. The pilot maneuvers the aircraft to maintain the delta symbol on the PFD as closely as possible to the underside of the command bars. Each GIA has a FD; one is active and the other is standby. The active FD is selected with the XFR key. The active FD analyzes the selected flight profile and compares it to current aircraft position. It then computes such functions as attitude, heading and roll rate as necessary to maneuver to the selected flight path. Using these calculations, the active FD moves the command bars on both PFDs. FD commands are limited to: • Pitch—±20° • Vertical acceleration—0.1g • Bank angle—30° • Bank Rate—5°/second

Autopilot The AP maneuvers the airplane to follow the FD by providing signals to the four flight-control servos, which are: • Pitch • Roll • Yaw • Pitch trim

FOR TRAINING PURPOSES ONLY

16-37

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10 FOR TRAINING PURPOSES ONLY

Figure 16-37.   AFCS Controller

12 13 16 17 18 19

1

2

3

4

15

5

14

6

7

11

8

9

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16-38

Adjusts the reference in Pitch Hold, Vertical Speed, and Flight Level Change Modes.

NOSE UP/DN Wheel

VNV Key

ALT SEL Knob

YD Key

AP Key

BANK Key

BC Key

HDG Knob

11

12

13

14

FOR TRAINING PURPOSES ONLY

15

16

18

19

Adjusts the Selected Heading and bug in 1° increments on the HSI (both PFDs). Press to synchronize the Selected Heading to the current heading.

Selects/deselects Backcourse Mode.

Selects/deselects Low Bank Mode.

Engages/disengages the autopilot.

Engages/disengages the yaw damper.

16 AVIONICS

Controls the Selected Altitude in 100-ft increments (the Baro minimum altitude is also available).

Selects/deselects Vertical Path Tracking Mode for Vertical Navigation flight control.

Toggles Airspeed Reference between IAS and Mach for Flight Level Change Mode.

SPD Key

10

Adjusts the Selected Course in 1° increments on the HSI of the corresponding PFD. Press to center the course deviation indicator and return the course pointer directly to the bearing of the active waypoint/station.

Selects/deselects Flight Level Change Mode.

Selects/deselects Vertical Speed Mode.

Selects/deselects Altitude Hold Mode.

Transfers between the active flight director and standby flight director.

Activates/deactivates the flight director only. Pressing once turns on the pilot-side flight director in the default vertical and lateral modes. Pressing again deactivates the flight director and removes the command bars. If the autopilot is engaged, the key is disabled.

Selects/deselects Navigation Mode.

Selects/deselects Approach Mode.

Selects/deselects Heading Select Mode.

DESCRIPTION

CRS Knobs

CONTROL

9 and 17

FLC Key

8

XFR Key

5

VS Key

FD Key

4

7

NAV Key

3

ALT Key

APR Key

2

6

HDG Key

1

#

Table 16-8.   AFCS CONTROLLER DESCRIPTIONS

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Yaw Damper

Control Yoke Switches

The YD stabilizes the aircraft in flight to prevent yaw instability. It can be engaged independently of the AP.

On each control yoke, three switches control various AFCS functions, particularly disabling the AP, and returning manual flight control to the crew (Figure 16-40):

The YD must be selected off during takeoff and landing.

CONTROLS AND INDICATIONS The AFCS is manipulated by various knobs and keys on the controller and by additional controls on the yoke and throttles.

AFCS Status Box On PFD The AFCS status box indicates the settings and status of active and pending AFCS functions: • White—Armed modes • Green—Active modes • Amber (flashing)—Canceled modes • Red (flashing)—Abnormal AP disconnect

AFCS Controller The AFCS controller is at the top of the instrument panel, above the MFD. It contains controls for the FD, AP, YD, and for making associated settings or selections on the PFDs (including target headings, courses, altitudes, vertical speeds, and airspeeds). The pushbutton keys are momentary-contact on/off toggle switches. Most keys and knobs affect both FD and AP functions. However, the FD and AP keys select or deselect all FD and AP operation. On the right side of most keys on the AFCS controller, small white LEDs indicate when the key is selected on, and the corresponding function is active or enabled (Figures 16-38 and 16-39).

• D OWN–UP trim switches—When operated to adjust pitch, trim also disengages the AP. • A P TRIM DISC switch—Immediately disconnects both the AP and the YD. • C WS (control wheel steering) switch— Momentarily disengages the AP and synchronizes the FD command bars with the current aircraft pitch (if not in glide slope or vertical navigation mode) and roll (if in roll hold mode). A CWS button is on each control wheel. Upon release of the CWS button, the FD may establish new reference points, depending on the current vertical and lateral modes. The CWS display presents in the AFCS status box (Figure 16-41).

GA Switches On the outboard side of each throttle is a recessed go-around (GA) pushbutton switch (Figure 16-40). Either switch sets the AFCS for optimum singleengine climb configuration during a takeoff or goaround. Pressing the GA switch: • Disengages the AP—The pilot can reengage the AP manually by pressing the AP key on the AFCS. The YD remains engaged. • Engages the FD—Command bars appear on the PFDs, and immediately direct 8° nose-up pitch (Figure 16-43). • On the ground, enables takeoff mode, which puts the command bars 10° pitch up. PILOT-SIDE PFD SELECTED

XFR

XFR

COPILOT-SIDE PFD SELECTED

Figure 16-38.   AFCS Status Bar

16-40

FOR TRAINING PURPOSES ONLY

LATERAL MODES

GPS ROL ARMED

ACTIVE

AUTOPILOT STATUS

YAW DAMPER STATUS

AP YD VS FLIGHT DIRECTOR INDICATOR ARROW

VERTICAL MODES

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ARMED

100FPM ALTS VPTH ACTIVE

MODE REFERENCE

Figure 16-39.   AFCS Status Box

During go-around mode, any attempt to modify pitch attitude (using the CWS button or the DN/ UP thumbwheel knob) causes the AFCS to revert to default pitch and roll modes.

OPERATION Emergency Descent Mode The AFCS has an emergency descent mode (EDM) that enables the AP, when pressurization is lost, to automatically descend the aircraft to 15,000 feet at VMO/MMO, regardless of pilot physiological condition (Figure 16-44). The aircraft must be above 30,000 MSL for EDM to arm. To exit EDM, disconnect the AP. If the AP is engaged, and the onboard cabin pressure sensors detect a cabin altitude greater than 14,500 feet, the AP automatically enters EDM, and displays EDM (white letters in a red box) on the PFDs.

Figure 16-40.   Control Yoke Switches CONTROL WHEEL STEERING

ROL

CWS YD PIT

ALTS

Figure 16-41.   CWS Display

During EDM, the AFCS: • Selects HDG mode and sets heading (heading bug on the PFD HSI) to 90° left of the current heading • Selects FLC mode, 0.63 m and sets altitude preselect (altitude bug on the PFD altitude display) to 15,000 feet To enable the maximum possible descent rate when in EDM, the pilot should immediately: • Reduce throttles to IDLE • Extend speedbrakes

Figure 16-42.   GA Switch

FOR TRAINING PURPOSES ONLY

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GO-AROUND MODE ACTIVE

AUTOPILOT DISCONNECT ANNUNCIATION FLASHES AMBER 5 SEC

NAV1 111.90 NAV2 113.80

114.10 SJC RH30L ??? 110.90 ISJC GA

SURNE ?? AP YD GA

DOS

4.6NM

BRG

303°

136.975 COM1 118.000 COM2

4000

130 120 110

4 103 100 2

COMMAND BARS INDICATE CLIMB

118.200 136.975

20

20

10

10

G

2500

4

2400

2

2300 80

22 60

2200

90

10

80

HDG

10

2100

306°

276°

30

W

-900 2

303°

CRS

4

2000 29.92IN

33

24

N

LOC2

21

3

S

6

15 ISA

12

0°c

E

RAT

+0°c

XPDR1

1263 ALT

R LCL

10:34:16

Figure 16-43.   Go-Around Mode HEADING SELECT MODE ACTIVE

EMERGENCY DECENT MODE ANNUNCIATION

NAV1 111.90 NAV2 113.80

114.10 110.90

AUTOPILOT ENGAGED

SJC

M .630

FLIGHT LEVEL CHANGE MODE ACTIVE

--.-NM BRG ---° 0.630 ALTS

DIS

HDG

AP YD FLC

210

20

20

10

10

4

30200 30100

190

10

180 HDG

29800

150° 12

15

GPS

CRS

283°

6

24

3

W 30

33

N +44°c

XPDR1

1253 ALT

Figure 16-44.   Emergency Descent Mode

16-42

4

29.92IN

S ENR

2

21

E

29900

10

060°

M170 .538

ISA

2

0 20 2200 30 00

200

0°c

SELECTED ALTITUDE SET TO 15,000 FT

30300

220

RAT

136.975 COM1 118.000 COM2

15000

EDM

230

SELECTED HEADING SET 90° LEFT OF CURRENT HEADING

118.200 136.975

FOR TRAINING PURPOSES ONLY

R LCL

9:44:38PM

SYNTHETIC VISION SYSTEM (SVS) The optional Synthetic Vision System (SVS) is a visual enhancement to the G1000 Integrated Flight Deck. SVS depicts a forward-looking attitude display of the topography immediately in front of the aircraft. The field of view is 30 degrees to the left and 35 degrees to the right. SVS information is shown on the Primary Flight Display (PFD), or on the Multifunction Display (MFD) in Reversionary Mode. The depicted imagery is derived from the aircraft attitude, heading, GPS three-dimensional position, and a nine arc-second database of terrain, obstacles, and other relevant features. The terrain data resolution of nine arc-seconds, meaning that the terrain elevation contours are stored in squares measuring nine arc-seconds on each side, is required for the operation of SVS. Loss of any of the required data, including temporary loss of the GPS signal, will cause SVS to be disabled until the required data is restored. The SVS terrain display shows land contours, large water features, towers, and other obstacles over 200’ AGL that are included in the obstacle database. Cultural features on the ground such as roads, highways, railroad tracks, cities, and state boundaries are not displayed even if those features are found on the MFD map. The terrain display also includes a north–south east–west grid with lines oriented with true north and spaced at one arc-minute intervals to assist in orientation relative to the terrain. The colors used to display the terrain elevation contours are similar to that of the TOPO map display. The Terrain Awareness and Warning System (TAWS) is integrated within SVS to provide visual and auditory alerts to indicate the presence of terrain and obstacle threats relevant to the projected flight path. Terrain alerts are displayed in red and yellow shading on the PFD. The terrain display is intended for situational awareness only. It may not provide the accuracy or fidelity on which to base decisions and plan maneuvers to avoid terrain or obstacles. Navigation must not be predicated solely upon the use of the TAWS terrain or obstacle data displayed by the SVS.

Revision 1.1

The following SVS enhancements appear on the PFD: • Pathways • Flight Path Marker • Horizon Heading Marks • Traffic Display • Airport Signs • Runway Display • Terrain Alerting • Obstacle Alerting

OPERATION SVS is activated from the PFD using the softkeys located along the bottom edge of the display. Pressing the softkeys turns the related function on or off. When SVS is enabled, the pitch scale increments are reduced to 10 degrees up and 7.5 degrees down. SVS functions are displayed on three levels of softkeys. The PFD Softkey leads into the PFD function Softkeys, including synthetic vision. Pressing the SYN VIS Softkey displays the SVS feature softkeys. The softkeys are labeled PATHWAY, SYN TERR, HRZN HDG, and APTSIGNS. The BACK Softkey returns to the previous level of softkeys. Synthetic Terrain must be active before any other SVS feature may be activated. HRZN HDG, APTSIGNS, and PATHWAY Softkeys are only available when the SYN TERR Softkey is activated (gray with black characters). After activating the SYN TERR Softkey, the HRZN HDG, APTSIGNS, and PATHWAY softkeys may be activated in any combination to display desired features. When system power is cycled, the last selected state (on or off) of the SYN TERR, HRZN HDG, APTSIGNS, and PATHWAY softkeys is remembered by the system.

Pathways Pathways provide a three-dimensional perspective view of the selected route of flight shown as colored rectangular boxes representing the horizontal and vertical flight path of the active flight plan. The box size represents 700 feet wide by 200 feet tall during enroute, oceanic, and terminal flight

FOR TRAINING PURPOSES ONLY

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phases. During an approach, the box width is 700 feet or one half full scale deviation on the HSI, whichever is less. The height is 200 feet or one half full scale deviation on the VDI, whichever is less. The altitude at which the pathway boxes are displayed is determined by the selected altitude during climb, cruise, and when the active leg is the final approach course prior to intercepting the glidepath/ glideslope. During a descent (except while on the approach glidepath/glideslope), the pathway boxes are displayed at the selected altitude, or the VNAV altitude programmed for the active leg in the flight plan, or the published altitude constraint, whichever is higher. Just prior to intercepting the glidepath/ glideslope, the pathway boxes are displayed on the glidepath/glideslope, or the selected altitude, whichever is lower.

Flight Path Marker The Flight Path Marker (FPM), also known as a Velocity Vector, is displayed on the PFD at groundspeeds above 30 knots. The FPM depicts the approximate projected path of the aircraft accounting for wind speed and direction relative to the three-dimensional terrain display. The FPM is always available when the Synthetic Terrain feature is in operation. The FPM represents the direction of the flight path as it relates to the terrain and obstacles on the display, while the airplane symbol represents the aircraft heading. The FPM works in conjunction with the Pathways feature to assist the pilot in maintaining desired altitudes and direction when navigating a flight plan. When on course and altitude the FPM is aligned inside the pathway boxes as shown.

Traffic WARNING Intruder aircraft at or below 500 ft. AGL may not appear on the SVS display or may appear as a partial symbol.

Traffic symbols are displayed in three dimensions, appearing larger as they are getting closer, and smaller when they are further away. Traffic within 250 feet laterally of the aircraft will not be displayed on the SVS display. Traffic symbols and coloring are consistent with that used for traffic displayed in the Inset map or MFD traffic page. If the traffic altitude is unknown, the traffic will not be displayed on the SVS display.

Airport Signs Airport Signs provide a visual representation of airport location and identification on the synthetic terrain display. When activated, the signs appear on the display when the aircraft is approximately 15 nm from an airport and disappear at approximately 4.5 nm. Airport signs are shown without the identifier until the aircraft is approximately 8 nautical miles from the airport. Airport signs are not shown behind the airspeed or altitude display. Airport signs are activated and deactivated by pressing the APTSIGNS Softkey.

TAWS Alerting Terrain alerting on the synthetic terrain display is triggered by Forward-looking Terrain Avoidance (FLTA). When an obstacle becomes a potential impact point the color of the obstacle matches the red or yellow X on the Inset map and MFD map displays. Obstacles are represented on the synthetic terrain display by standard two-dimensional tower symbols found on the Inset map and MFD maps and charts. Obstacle symbols appear in the perspective view with relative height above terrain and distance from the aircraft. Unlike the Inset map and MFD moving map display, obstacles on the synthetic terrain display do not change colors to warn of potential conflict with the aircraft’s flight path until the obstacle is associated with an actual FLTA alert. Obstacles greater than 1000 feet below the aircraft altitude are not shown. Obstacles are shown behind the airspeed and altitude displays.

Traffic symbols are displayed in their approximate locations as determined by the related traffic systems.

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FOR TRAINING PURPOSES ONLY

Revision 1.1

Figure 16-45.   SVS on the PFD

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Figure 16-46.   Terrain Alerts with SVS on the PFD

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AUDIO PANEL DESCRIPTION The Mustang avionics system includes an audio panel for each crewmember (pilot and copilot). The audio panels are visible as vertical switch panels on the instrument panel immediately outboard of each PFD. Each audio panel includes an audio amplifier, a marker beacon receiver, and controls for selecting and managing audio sources. These panels manage all audio sources, including COM transceivers, NAV, ADF, DME and marker beacon receivers, aircraft intercom and aural warning systems. They also control the switching of microphones and headsets. The audio panel also includes a digital voice recorder, which holds up to 2 minutes and 30 seconds of recorded audio to assist the pilot in recording ATC clearances.

CONTROLS AND INDICATIONS The audio panels use small keys for each item. Above each key is a small, triangle-shaped LED indicator, which illuminates when the device for the corresponding button is selected on that audio panel (Figure 16-47 and Table 16-9). Each audio panel may have different selections from the other panel.

Power-Up During aircraft power-up, all audio panel annunciators illuminate for 2 seconds. Then all the audio switch selections (and annunciators) return the setting in effect when the aircraft was powered down. There are two exceptions: The speaker and intercom are always activated during power-up, and remain active until deselected.

Fail-Safe COM Operation If both GIAs fail, the audio panels directly connect the pilot headset and microphone to the COM1 transceiver, and the copilot headset and microphone to the COM2 transceiver. (No audio is available to the speaker in this situation.) If only one GIA fails, the crewmember on that side will only have audio access to the corresponding COM transceiver.

1

2

3

4

5

6

7

8

9

10

11

12

13

14

15

16

17

18

19

20

21

22

23

24

Figure 16-47.   Audio Panel

FOR TRAINING PURPOSES ONLY

16-47

CONTROL

COM1 MIC

COM1 COM2 MIC

COM2 COM3 MIC COM3 PA

TEL MUSIC SPKR

MKR/MUTE

HI SENS DME NAV1 ADF NAV2 AUX REC

PLAY

INTR COM MAN SQ

ICS Knob

MSTR Knob Reversionary Mode Button

#

1

16-48

2 3

4 5 6 7

8 9 10

11

12 13 14 15 16 17 18

19

20 21

22

23 24

Selects the #1 transmitter for transmitting. COM1 receive is simultaneously selected when this key is pressed, allowing received audio from the #1 Com receiver to be heard. When selected, audio from the #1 Com receiver can be heard. Selects the #2 transmitter for transmitting. COM2 is simultaneously selected when this key is pressed allowing received audio from the the #2 Com receiver to be heard. When selected, audio from the #2 Com receiver can be heard. Not used on the Cessna Citation Mustang. Not used on the Cessna Citation Mustang. Selects the passenger address system (if installed). The selected Com transmitter is deselected when the PA key is pressed. There are no cabin speekers. Headset jacks for passagers are an option. Not used on the Cessna Citation Mustang. Not used on the Cessna Citation Mustang. Pressing this key selects and deselects the corresponding cockpit speaker. All audio will be heard on the speaker, to include audio warnings. Activates the marker beacon receiver audio. Pressing mutes the currently received marker beacon receiver audio. Push again to turn off all marker audio. Press to increase marker beacon receiver sensitivity. Press again to return to normal. Pressing turns DME audio on or off. When selected, audio from the #1 Nav receiver can be heard. Pressing turns on or off the audio from the ADF receiver. When selected, audio from the #2 Nav receiver can be heard. Not used on the Cessna Citation Mustang. Press to start the recording up to 2.5 minutes of COM receiver audio. When no audio is being received, nothing is recorded. Press again to stop recording. Press once to play the last recorded audio. Press again to stop playing. Press twice quickly while audio is playing and the previous block of recorded audio will be played. Each subsequent two presses will skip back to the previously recorded block. Pressing selects the pilot/copilot intercom on both audio panels. Press again to deselect the intercom. Press to enable manual squelch for the intercom. When active, press the ICS knob to illuminate ‘SQ’. Turn the ICS knob to adjust squelch. Turn to adjust intercom volume or squelch. Press to switch between volume and squelch control as indicated by the ‘VOL’ or ‘SQ’ being illuminated. The MAN SQ key must be selected to allow squelch adjustment. The master volume control adjusts volume for all headset audio. Pressing manually selects reversionary mode.

DESCRIPTION

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Table 16-9.   AUDIO PANEL DESCRIPTIONS

CITATION MUSTANG PILOT TRAINING MANUAL

FOR TRAINING PURPOSES ONLY

COM Radio Priority The G1000 includes two COM radios. COM1 is in GIA#1 and COM2 is in GIA#2. Either pilot may communicate using either COM radio. Each crew microphone is connected to a COM radio through a COM MIC key. For each crewmember, their COM MIC selection (COM MIC1 or COM MIC2) makes the selected COM radio their active transceiver, connected to their microphone. On each PFD, this is indicated by green digits for the active frequency of the primary transceiver. All other frequencies use white digits. The pilot may listen to the other COM radio by selecting the COM1 or COM 2 button. Both crewmembers may select the same primary transceiver, or each may select a different primary transceiver. If both crewmembers select the same primary transceiver, and each crewmember keys their microphone at the same time, the microphone that transmits through the COM radio is the microphone of the crewmember with priority for that COM unit (the pilot has priority on COM1, and the copilot has priority on COM2). The hand microphone is connected to the pilot audio panel.

TRAFFIC INFORMATION SERVICE The TIS provides the pilot with limited information about nearby potential air-traffic hazards in terminal areas that have TIS-capable ground-based radar.

NOTE TIS is an advisory service only, to help the pilot locate traffic visually. It is the pilot responsibility to see and avoid traffic. The service has three basic requirements: • TIS-equipped aircraft must have an operating mode-S transponder (the G1000 includes two). • Conflicting aircraft must have an operating mode-A, mode-C or mode-S transponder. • Both aircraft must be within range (approximately 55 miles) of an air traffic control

radar that has TIS enabled. These radars are most likely to be in congested terminal areas. (Refer to the FAA Airman’s Information Manual for current TIS-radar coverage status). The ground-based radar, when providing TIS, detects transponder-operating aircraft that are in proximity to other aircraft, and transmits traffic information about those aircraft to other aircraft nearby. If the client aircraft are operating a modeS transponder, ground radar transmits the location, direction, speed, and vertical proximity of the other nearby aircraft. Ground radar only reports a maximum of eight traffic hazards, primarily those within 7 NM horizontally and+3,500/–3,000 feet vertically of the client aircraft. The mode-S transponder receives this information, and depicts it on the TRAFFIC display of the MFD map and the PFD inset map.

CAUTION Traffic Information Service is only effective when within range of a TIS-capable terminal radar site. It may operate intermittently, or not at all due to interference with transmission or reception (by obstacles, terrain or the aircraft itself, or condition of the ground radar). TIS does not provide information on aircraft without an operating transponder. TIS is a “line-of-sight” system that uses tracking to report and update traffic notifications (every 5 seconds) on the MFD traffic map page or on the PFD map inset. Traffic alert messages appear on the PFD to the right of the top of the airspeed tape. The GTX 33 transponder has selective addressing or MODE SELECT (mode-S) capability. Mode-S functions include a data link capability that allows information to be exchanged between aircraft and various air traffic control facilities. TIS uses data acquired by surveillance of the modeS radar system. During turns or other maneuvering, TIS data may become random as a result of interference between aircraft and ground antennae.

FOR TRAINING PURPOSES ONLY

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Additionally, reception may be interrupted by terrain (e.g. mountains). TIS performs an automatic test during power-up and displays a standby screen on the traffic map page. If the power-up test fails, a NO DATA, DATA FAILED or FAILED message displays on the second page in the map group. The traffic map page displays the following TIS information: • Interrogating aircraft location, surrounding traffic locations, and range marking rings • C ur rent traff ic mode (OPERATE, STANDBY) • Traffic alert messages (FAILED, DATA FAILED, NO DATA, UNAVAILABLE) • Traffic display banner (AGE, TRFC COAST, TA OFF Range, TRFC RMVD, TRFC FAIL, NO TRFC DATA, TRFC UNAVAIL, TRAFFIC) When the aircraft is airborne, TIS switches from standby to operating mode. TIS OPERATING displays in the upper left corner of the traffic map page. A traffic advisory (TA) symbol displays when traffic is within the following range: • ±500 feet • 0.5 NM • 34 seconds A traffic advisory is accompanied by an aural “traffic” callout as well as when traffic becomes unavailable. In addition, an alert box appears on the PFD and the inset map is automatically displayed with traffic information. Altitude deviation from the interrogating aircraft is shown above or below the target symbol. Target climb or descent is shown as an up or down arrow. TIS assists pilots in the visual acquisition of other aircraft in visual meteorological conditions only. Do not use as a collision avoidance system. Pilots still have “see and avoid” responsibility for possible traffic conflicts. During instrument meteorological conditions, TIS must not be used for maneuvering when no visual contact exists with other aircraft.

16-50

TRAFFIC ADVISORY SYSTEM As a substitute for the standard G1000 traffic information service (TIS), an optional traffic advisory system (TAS) is available. TIS and TAS have similar functions and indications. However, while TIS is dependent upon TIS-enabled ground stations to detect and report traffic, TAS directly detects and identifies traffic with an onboard interrogator. As a result, TAS can operate anywhere, not only in places with TIS-enabled ground stations. Also, because it is direct, TAS is faster, updating every half-second (TIS updates every 5 seconds).

WARNING TAS cannot detect all traffic hazards. TAS is only effective at identifying traffic with an operating transponder. Only aircraft with a transponder operating in Mode C (altitude encoding) or Mode S (data link) can be identified by relative altitude. It is the pilot’s responsibility to see and avoid traffic. Do not rely on TAS to avoid traffic in instrument meteorological conditions (IMCs). When equipped with TAS, the Citation Mustang uses the Honeywell KTA 870 (refer to the Honeywell KTA 870 Pilot’s Guide).The unit connects to the MFD through GIA. Indications are provided on the MFD and PFD, and aural alerts through the audio panel.

TAS Traffic Symbols Traffic indications are similar to those previously described for TIS. On MFD/PFD maps that display traffic, the following symbols appear to indicate traffic that the TAS unit has identified: • S olid yellow dot—Aircraft with TAS range, and a traffic hazard/threat • Y ellow/black dot (on outer ring)—Aircraft outside TAS range; whether traffic hazard threat or not is unknown • S olid white diamond—Proximity traffic not an apparent hazard/threat, but approaching your altitude

FOR TRAINING PURPOSES ONLY

• H ollow white diamond—Proximity traffic, all other • “ TA” text with data—Nonbearing traffic (bearing unknown), text includes distance, altitude, and trend, if known. (On the TRAFFIC MAP page, this message displays near the center of the map). Digits appear above or below the symbol, indicating the difference between the aircraft’s altitude and the other aircraft’s altitude. The digits are above the symbol when the aircraft is above, and below the symbol when the aircraft is below. Also, a minus (–) symbol before the digits indicates the aircraft is below, and a plus (+) symbol before the digits indicates the aircraft is above. An up arrow beside the symbol indicates that the other aircraft is climbing (at 500 fpm or more).

MAP–TRAFFIC MAP Page And TAS Controls To view TAS indications exclusively or to adjust TAS settings, switch to the MAP–TRAFFIC MAP page (use the large FMS knob to select the MAP page group and the small FMS knob to select the TRAFFIC MAP page). The system self-test verifies operation of the aural warning when the STANDBY softkey is pressed, followed by pressing the TEST softkey; in approximately 8 seconds, three different traffic symbols appear and the aural message “TAS system test OK” is heard. (If the test fails, a different message is heard). On the MAP–TRAFFIC MAP page, pressing the OPERATE softkey activates TAS, which begins displaying transponder-operating traffic in the area. The alert “TAS OPERATING” appears in the upper left corner of the TAS display. To adjust the altitude-based traffic altering function, use the ALT MODE softkey. Select the threat zone to monitor: BELOW (current altitude), NORMAL (within ____ feet of current altitude), ABOVE, or UREST (unrestricted). Rotate the RANGE/PAN joystick knob to change viewing range, and move the joystick knob vertically or horizontally to pan the display.

Revision 1.1

TAS On Other MFD/PFD Maps TAS indications can be viewed on other MFD maps, and on the PFD inset map, by pressing the MAP softkey below the corresponding page, followed by the TRAFFIC softkey.

NOTE The TRAFFIC softkey label is black letters on a gray background when TAS is active on that map.

AIRBORNE WEATHER RADAR A GWX 68 airborne weather radar provides precipitation returns and ground-mapping returns. The radar primarily assists the pilot in detecting dangerous thunderstorms along the flight path. It also clarifies the boundaries and intensities of storm cells and locations of severe precipitation. The GWX 68 has both horizontal and vertical scan capability, which provides greater detail on the position of storms, including their heights. The groundmapping feature of the onboard radar assists the pilot in identifying landmarks and bodies of water. The GWX 68 radar unit is in the nose of the aircraft. The antenna sweeps across an arc of 90° horizontally (45° each side of center), and 60° vertically (30° above and below the horizon). The radar transmits a very-high-energy radio beam, which is reflected from precipitation and surface features.

WARNING The radar beam is dangerous, and close exposure can cause severe injury. When operating the radar on the ground, ensure that no persons or objects are within 11 feet of the antenna. (People inside the aircraft are safe, if no object is close enough to the antenna to reflect the entire beam into the aircraft.) Do not begin to transmit, until the aircraft is clear of all persons and objects on the ground within 11 feet.

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WEATHER RADAR PAGE AND CONTROLS On the MFD, the MAP–WEATHER RADAR page provides access to the weather radar depiction and control. On the ground, the weather radar must be activated to operate. Pressing the MODE softkey causes other softkeys to appear, providing control of the radar. The STANDBY key initiates the 1-minute warm-up period required before operating the radar. After 1 minute, ensure the aircraft is clear of all persons and objects within 11 feet of the antenna and forward of the aircraft, then select the WEATHER softkey, which activates the weather radar antenna.

NOTE If the aircraft is already airborne, the STANDBY mode is not required before energizing the antenna. After landing, the radar automatically returns to STAND-BY mode. Pressing the BACK key exits the MODE softkey menu, and returns to the main radar softkey menu.

RADAR DISPLAY AND INDICATIONS The GWX 68 digital radar utilizes a four-color display capable of scanning airspace ahead of the aircraft through various pilot adjustable angles. Specific sectors may be monitored through a horizontal plane of 20°, 40°, 60°, or 90°. A vertical scanning function provides for scanning through 8° of coverage, selectable by the pilot to assist in analyzing storm tops, gradients, and cell buildup activity at various altitudes. The following additional features provide an added margin of safety: • E xtended sensitivity time control (STC)— Logic that automatically correlates distance of the return echo with intensity, so that cells do not suddenly appear to get larger, as they get closer.

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• W ATCH (weather attenuated color highlight)—Helps identify possible “shadowing” effects of short-range cell activity. Radar return signals are weakened, or attenuated, by intense precipitation. The same is true for extensive areas of lesser precipitation. Under these circumstances, the potential exists for “the storm behind the storm” to be masked from viewing with airborne radar units. • W eather alert looks ahead for intense cell activity in the 80–320 NM range, even if these ranges are not being monitored. Refer to Table 16-10. Table 16-10.   PRECIPITATION INTENSITIES WEATHER MODE COLOR

INTENSITY

BLACK


2 IN./HR.

Range Adjustment The range of the radar display is adjustable. Rotate the RANGE/PAN joystick knob to adjust the display range. The current distance to each dotted arc is noted at one end of each arc. Range adjustment only affects the display; it does not change the beam intensity or sensitivity (gain) of the radar.

WATCH Softkey The GWX 68 radar includes a WATCH (weatherattenuated color highlight), which projects a colored “shadow” to indicate a warning to the pilot of possible attenuation. The WATCH softkey selects or deselects this feature.

Gain Control Radar sensitivity may be adjusted by pressing the GAIN softkey, then rotating the small FMS knob to adjust the radar gain. This could change the color indications and give a false indication of actual precipitation and surface returns. However, adjusting radar-return gain may increase clarity of some situations.

FOR TRAINING PURPOSES ONLY

Table 16-11.   GROUND TARGET RETURN INTENSITY LEVELS

CAUTION When finished viewing at an adjusted gain, press the GAIN softkey again to turn off the gain adjustment, and return the radar to calibrated-gain mode, so that colors are again accurately indicating the strength of the returns.

ANTENNA STABILIZATION The GWX 68 radar antenna is stabilized through gyro data to minimize low quality images as the pilot maneuvers around weather, particularly in turbulence. Through control key inputs, the STAB feature may be selected ON or OFF as desired.

ANTENNA TILT Adjusting the vertical tilt angle of the antenna selects the vertical direction to scan. This enables the pilot to focus attention on a particular area, such as the tops of a storm, or the ground. To adjust the antenna tilt from the MAP–WEATHER RADAR page, press the FMS knob, and a cursor/highlight appears in the TILT field of the display. Adjust the antenna (radar beam) tilt as desired, using the small FMS knob. If the VERTICAL display is presented, a line appears indicating the angle of tilt.

GROUND MAP MODE COLOR

INTENSITY

BLACK

0 DB

CYAN

> 0 DB TO < 9 DB

YELLOW

9 DB TO < 18 DB

MAGENTA

18 DB TO < 27 DB

BLUE

27 DB AND GREATER

TERRAIN AWARENESS AND WARNING SYSTEM An integral terrain awareness and warning system (TAWS) provides the pilot with aural and visual warnings of terrain and obstacles near, at or above the aircraft altitude or flight path (Figure 16-48). TAWS relies on current terrain and obstacle databases.

TRAFFIC 96.7

117.95 117.95 240

TERRAIN DIS

28.7NM

BRG

TERRAIN

TRAFFIC

20

20

300°

136.075 136.975

15200 15180 4

Figure 16-48.   Traffic and Terrain Display

GROUND MAPPING Ground mapping mode highlights surface features. Cities and other built-up areas typically provide strong returns, while water typically provides little or no return. Ground mapping may assist the pilot in navigation. To select ground mapping from the MAP – WEATHER RADAR page, press the MODE softkey. Press the GROUND softkey. Adjust antenna tilt to select the specific angle desired for ground mapping. Refer to Table 16-11.

HAZARD DEPICTIONS AND ALERTS Red areas on the terrain map indicate terrain within 100 feet (approximately 33 meters) of the current aircraft altitude or higher, including terrain above the current aircraft altitude (Figure 16-49). Amber areas on the terrain map indicate terrain between 100 feet and 1,000 feet below the aircraft altitude. Black areas on the map indicate terrain greater than 1,000 feet below the aircraft. Obstacles are similarly colored, and depicted with standard obstacle symbols. If TAWS detects an imminent impact (based on aircraft course, ground speed, and vertical speed), an X is depicted in the current aircraft course at the point of expected impact.

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The data link is remotely mounted; its controls are on the G1000 panels and displays.

TERRAIN

NOTE For information on activating XM Weather service subscription, refer to the Garmin GDL69/-69A XM Satellite Radio Activation Instructions, or refer to “Activating XM Radio Services” in the Garmin G1000 Pilot’s Guide for the Cessna Citation Mustang.

–100 FT –1000 FT

NOTE Figure 16-49.   Terrain Colors

Color-coded warnings (for red areas) and cautions (for amber areas) are signaled through warning flags appearing on the PFD, and through aural warnings. If any page, except the TERRAIN PROXIMITY page is selected on the MFD, a popup warning message appears in the lower right corner of the display. In this situation, pressing the ENT (enter) key immediately switches the MFD to the MAP–TERRAIN PROXIMITY page; pressing the CLR (clear) key causes the current map to remain visible on the MFD. TAWS hazard depictions may also be displayed as overlays on other maps by selecting the TERRAIN softkey. However, TAWS hazard depictions cannot be displayed simultaneously with XM Weather NEXRAD images or airborne weather radar indications, because they use similar color coding to indicate weather threat areas.

XM WEATHER AND GDL 69/69A DATA LINK The MFD can display satellite-broadcast weather information. From any location in North America, the MFD can receive aviation weather information from the XM Weather™ satellite broadcasts. The XM Weather service also provides temporary flight restriction (TFR) reports. The system can download XM Weather data from satellite broadcasts, which is received from a commercial subscription service. The GDL 69/69A data link receiver enables the G1000 to receive XM Weather (and TFR) data.

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This section contains basic, limited information on the features and operation of the XM Weather service data option. For more information on XM Weather and the Garmin GDL 69/69A Data Link LRU, refer to the Garmin manuals and guides supplied with your Citation Mustang.

XM Weather Information Available The user level (class) of XM Weather service and specific weather-information products available are listed on the AUX–XM INFORMATION page on the MFD, along with the associated identification codes. Complete XM Weather service can include the following information: • Graphical information depictions: °° NEXRAD data (NEXRAD) °° METAR data (METAR) °° Wind data (WIND) °° Echo tops (ECHO TOP) °° Cloud tops (CLD TOP)

°° Lightning strikes (XM LTNG) °° Storm cell movement (CELL MOV)

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• Graphic and text information: °° SIGMETs/AIRMETs (SIG/AIR) °° Surface analysis including city forecasts (SFC) °° County warnings (COUNTY) °° Freezing levels (FRZ LVL) °° Hurricane track (CYCLONE) °° Temporary flight restrictions (TFR) • Text-only information: °° METAR data °° Terminal aerodrome forecasts (TAF)

XM Weather Display Options And Limitations On MFD map pages (and PFD inset maps), softkeys allow the user to select which XM Weather information (if any) to display. Multiple XM Weather information products can be displayed simultaneously on some pages. Note that some XM Weather information products cannot display with certain other XM Weather products. Likewise, some XM Weather products cannot display with certain information from other sources. The NEXRAD legend is shown in Figure 16-48.

MAP–XM WEATHER DATA LINK Page In the MAP page group, the WEATHER DATA LINK page provides the maximum number of views of XM Weather information products. The range/pan joystick provides an interface with many of the weather products. When the map pointer is pointing to, or in, a specific feature on the map, the center box of the map pointer information bar gives information on that feature. When pointing to a specific symbol representing a weather hazard, information is given about that hazard. When the hazard is an area, and the map pointer is on the boundary, or inside the boundary of the area, information about the hazard appears in the center box. When the map pointer is in a place where multiple hazards exist, only one is identified in the center box; SIGMETs have priority.

For boundaries of hazard areas, refer to the Garmin manuals. On the main softkey menu, the MORE WX softkey leads to more weather products. Selecting some of these features may require a more detailed softkey selection, such as specific altitudes for winds aloft depiction, or forecast periods (current, 12-hour, 24-hour, etc.) for surface weather depictions.

Viewing METARs and TAFs METARS and TAFs are available in text format through XM Weather. METARs and TAFs can be viewed by selecting the WPT–AIRPORT INFORMATION page and pressing the WX softkey. This changes the page to the WPT–WEATHER INFORMATION page (for the selected airport). On the right side of the page are windows for METAR reports and TAF forecasts.

STANDBY FLIGHT INSTRUMENTS The Mustang has four standby flight instruments that can function independent of the electrical system or any other system, including the avionics. These instruments are: • Standby attitude indicator • Standby airspeed indicator • Standby altimeter • Standby magnetic compass Each 2-inch indicator is a self-contained unit. Electrical power for these units comes through the STBY INST switch from the right avionics electrical bus. The bus connects to normal DC power, and also has a dedicated 1.2 ampere-hour standby battery. The standby battery is in the nose and connects to the main electrical bus for charging power (refer to Chapter 2—“Electrical System”). Circuit breakers for the standby battery charging circuit, and each of the standby instruments (except the compass), are in the AVIONICS section of the right CB panel.

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Weather Product

Symbol

Expiration Time Refresh Rate (Minutes) (Minutes)

NEXRAD

30

5

Cloud Top (CLD TOP)

60

15

Echo Top

30

7.5

XM Lightning (LTNG)

30

5

Cell Movement (SCIT)

30

12

SIGMETs/AIRMETs (SIG/AIR)

60

12

METARs

90

12

City Forecast (CITY)

60

12

Surface Analysis (SFC)

60

12

Freezing Levels (FRZ LVL)

60

12

Winds Aloft (WIND)

60

12

County Warnings (COUNTY)

60

5

Cyclone Warnings (CYCLONE)

60

12

Radar Coverage

no product image

30

5

TFRs

no product image

60

12

TAFs

no product image

60

12

Figure 16-50.   NEXRAD Legend

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The pneumatic instruments (airspeed and altitude) bypass both ADCs and are directly connected (pneumatically) to the pilot-side pitot-static system. The magnetic compass operates without external power. However, an internal light in the compass receives DC power from the same source as the other standby instruments

STANDBY ATTITUDE INDICATOR

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STANDBY AIRSPEED INDICATOR The standby airspeed indicator displays aircraft airspeed (Figure 16-52). It measures the ram and static air pressures directly from the pilot-side pitot-static system, and presents the airspeed on a single pointer indicator. The pointer is referenced against a dial marking to display the indicated airspeed. VMO is marked with a red line at 250 KIAS.

The standby attitude indicator is a self-contained unit that provides visual pitch and roll aircraft attitude information (Figure 16-51). The indicator contains an electrically powered gyroscope, which maintains vertical orientation. The instrument is internally lighted. A PULL TO CAGE knob allows the gyro to be aligned prior to flight.

Figure 16-52.   Standby Airspeed Indicator

To determine MMO limits, compare indicated altitude to the placarded altitudes, and determine the corresponding limiting calibrated airspeed (KCAS) from the table.

WARNING When relying only on standby airspeed, use caution to remain below VMO and MMO.

Figure 16-51.   Standby Attitude Indicator

CAUTION The attitude indicator may be damaged if the PULL TO CAGE knob is released with a “snap.” Release pull to cage knob slowly in order to avoid a “snap” release.

Green and red LEDs on the upper left corner of the bezel provide self-test indications during power-up.

NOTE When power is applied to the standby airspeed indicator, the green and red LEDs flash and the standby airspeed indicator needle rotates clockwise to the maximum limit, then counterclockwise to the zero-park position. After completion of the needle travel test, the needle returns to the measured pressure, and normal operation of the instrument begins.

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If the red LED illuminates or the green LED is dark, the instrument is not operational and maintenance is required. The instrument pointer may be tested by pushing the built-in test (BIT) switch on the lower right bezel.

STANDBY ALTIMETER DISPLAY The standby altimeter displays aircraft altitude (Figure 16-53). The instrument measures the static air pressure directly from the pilot-side static system. It presents the baro-corrected altitude on a digital readout at the top center of the instrument dial, and a pointer displays the precise altitude on the dial markings. A barometric-setting knob is on the lower-left corner of the altimeter. The setting appears in a digital readout as hectoPascals (MB) on the lower-left face and inches of mercury (INHG) on the lower-right face of the instrument dial.

When the power-up self-test is complete, verify the green LED illuminates, and the indicator needle reads the current altimeter barometric setting. If the red LED illuminates, the instrument is not operational, and maintenance is required. The instrument pointer may be tested by pushing the BIT switch on the lower right bezel.

STANDBY MAGNETIC COMPASS The standby magnetic compass is on the windshield post above the assist handle (Figure 16-54). It uses a standard magnetic compass wheel inside a kerosene-filled chamber, viewed through a window in the front of the instrument. The compass has a calibration placard on the windshield post. Aircraft heading appears in the window under the lubber line in the center of the window.

Figure 16-53.   Standby Altimeter

Green and red LEDs on the upper left corner of the bezel provide self-test indications during power-up.

NOTE When power is applied to the standby altimeter, the green and red LEDs flash. The standby altimeter needle moves to the zero park position and then returns to the measured pressure once normal operation of the instrument begins.

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Figure 16-54.   Magnetic Compass

CAUTION Avoid placing metal or magnetic objects near the compass, because these can cause errors.

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NOTE Magnetic compass is influenced by the windshield heat, cockpit fan, and fresh air fan. These items must be off prior to referencing magnetic compass heading, then may be reselected on. The items must then be reselected off prior to each referencing of the magnetic compass.

LIMITATIONS For specific limitations, refer to the FAA-approved AFM.

EMERGENCY/ ABNORMAL For specific information on emergency/abnormal procedures, refer to the appropriate checklist or FAA-approved AFM.

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QUESTIONS 1. The two GDC 74B air data computers are responsible for supplying information to what instruments? A. Standby airspeed indicator, standby altimeter, and standby attitude indicator B. PFD attitude indicator and horizontal direction indicator C. PFD airspeed indicators, altimeters, and vertical speed indicators D. MFD XM weather information system 2. The open green circle that can appear on the airspeed indicator indicates: A. The stall warning computer has indicated a reference approach speed cue of 1.3 VS1 B. The stall warning computer has indicated a reference approach speed cue of 1.1 VS0 C. The stall warning computer is indicating an approach speed of 1.5 VLSA D. The airspeed indicator has a fault to determine the final approach speed 3. The standby flight instruments on the Mustang aircraft are: A. Vacuum and ram-air driven; instruments indicate airspeed, altitude, and attitude regardless whether power is lost to the aircraft B. Electrically driven instruments that have their own dedicated standby battery pack, which can power the airspeed, altitude, and attitude instruments for 30 minutes C. Powered with the avionics power switch set to the on position D. Being electrically charged even with a loss of all DC power.

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4. If the pilot forgets to turn off the takeoff V speeds in the timer/reference box on the PFD, the pilot should: A. Immediately turn off the takeoff V speeds as this can cause a hazardous condition when trying to land the aircraft B. Press the TMR/REF and MENU bezel key on the PFD and choose to turn off all takeoff V speeds, or accelerate the aircraft above 160 knots C. Press the MENU bezel key on the PFD and choose to turn on the landing V speeds D. Shutdown the avionics power switch and reboot the PFD 5. When setting the BARO MIN on the PFD, the pilot should keep in mind that they are setting: A. The height above ground level for the minimums of an approach B. The altitude that they want the aircraft to level to when climbing to a selected altitude C. The altitude that they want the aircraft to descend to when reaching the minimums of an approach D. The altitude that is the decision height or minimum descent altitude on an approach 6. If the MFD fails: A. The pilot and copilot PFD displays normally and the display backup button at the bottom of the pilot or copilot audio panel has to be selected to display the necessary flight information, engine information, and CAS window. B. The pilot and copilot PFDs automatically go into reversionary mode. C. The pilot PFD remains unaffected and the copilot screen automatically switches to reversionary mode. D. The pilot screen only automatically switches into reversionary mode.

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7. For flight planning purposes, the pilot can enter the weight and balance information by going to the: A. MAP page on the MFD B. NRST page group on the PFD C. AUX page group on the MFD D. PROC page group on the PFD 8. When the pilot chooses a lateral or vertical mode on the AFCS controller, the AFCS status box indicates: A. Green for standby and white for active B. Magenta for standby and white for active C. White for standby and green for active D. Both white for standby and for active 9. The CWS button on the control yoke will: A. Momentarily disengage the autopilot, but the servos remain engaged B. Momentarily disengages the AP and synchronizes the FD command bars with the current aircraft pitch and roll attitude C. Disengage the yaw damper D. Discontinue the pitch mode of the autopilot 10. The PLAY key on the audio panel allows the pilot to: A. Play back the previous recorded 2-minute block of a received audio transmission B. Play the last audio transmission on the No. 2 NAV C. Receive the last PA request from the passengers D. Play the XM radio song that is being uploaded 11. While operating the GWX weather radar system on the ground, make sure that: A. The gain is properly calibrated B. The radar system has been tested C. The No. 2 COM is set to the ATIS D. No one is within 11 feet of the radar antenna if the weather is turned on while on the ground

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CHAPTER 17 OXYGEN SYSTEM CONTENTS GENERAL............................................................................................................................ 17-1 OXYGEN SYSTEM............................................................................................................. 17-3 Description.................................................................................................................... 17-3 Components................................................................................................................... 17-3 Controls and Indications................................................................................................ 17-5 Operation....................................................................................................................... 17-8 Limitations.................................................................................................................... 17-9 LIMITATIONS................................................................................................................... 17-10 EMERGENCY/ABNORMAL........................................................................................... 17-10 QUESTIONS..................................................................................................................... 17-11

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ILLUSTRATIONS 17-1.

Oxygen System Schematic.................................................................................... 17-2

17-2.

Standard Crew Masks............................................................................................ 17-3

17-3.

Crew Oxygen Mask, Stowed (Pilot Side).............................................................. 17-3

17-4.

Flow Indicator........................................................................................................ 17-4

17-5.

Pilot and Copilot Consoles.................................................................................... 17-4

17-6.

Passenger Oxygen Mask........................................................................................ 17-5

17-7.

OXYGEN CUTOFF Knob.................................................................................... 17-5

17-8.

Oxygen Control Valve Knob.................................................................................. 17-5

17-9.

Mic Switches......................................................................................................... 17-6

17-10. Oxygen Pressure Gauge......................................................................................... 17-7 17-11. Overboard Discharge Indicator.............................................................................. 17-7 17-12. Crew Mask Controls.............................................................................................. 17-7 17-13. Oxygen Bottle........................................................................................................ 17-9

TABLES Table Title Page 17-1.

OXYGEN SUPPLY DURATION....................................................................... 17-10

17-2.

CAS MESSAGE................................................................................................. 17-10

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17 OXYGEN SYSTEM

CHAPTER 17 OXYGEN SYSTEM

INTRODUCTION This chapter covers the oxygen system and squat switch (weight-on-wheels sensing) systems on the Citation Mustang. Oxygen is available to the crew and passengers during pressurization system malfunctions or when required. One squat switch on each landing gear indicates when weight is on the wheel. The squat switches provide signals to various aircraft systems.

GENERAL The oxygen system includes the crew and passenger distribution systems (Figure 17-1). Oxygen is available to the crew at all times and is available to the passengers either automatically above a cabin altitude of approximately 14,800 feet, or manually at any altitude by the oxygen control valve.

The squat switches provide signals to various aircraft systems, controls and indications to adjust them for different operation, depending on whether the aircraft is in flight or on the ground. The specific role of the squat switch in each aircraft system is described in detail in that system chapter of this manual.

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17-1

17-2

LP

FILL PORT

FOR TRAINING PURPOSES ONLY

CHECK VALVE

OXYGEN CONTROL VALVE SELECTOR

LH MULTI-FUNCTION PCB WITH TRANSDUCER

RH MULTI-FUNCTION PCB WITH TRANSDUCER

PILOT OXYGEN MASK

CABIN DROP BOX ASSEMBLY (SINGLE MASK)

COPILOT OXYGEN MASK

Figure 17-1.   Oxygen System Schematic

REGULATED PASSENGER OXYGEN (CONTROLLED BY OXYGEN CONTROL VALVE)

REGULATED CREW OXYGEN (AVAILABLE TO CREW AT ALL TIMES)

BOTTLE PRESSURE

LEGEND

NORMAL (SOLENOID VALVE NORMALLY CLOSED)

OXYGEN CONTROL VALVE

DROP MASK

CREW ONLY

LOW OXYGEN PRESSURE SWITCH (OXYGEN OFF - CAS MESSAGE)

OXYGEN CUTOFF VALVE

OXYGEN BOTTLE

REGULATOR

OXYGEN BLOWOUT DISC

OXYGEN SUPPLY HANDLE

CABIN DROP BOX ASSEMBLY (DOUBLE MASK)

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DESCRIPTION The Mustang oxygen system is primarily for emergency use, but also allows limited-duration nonemergency use. It provides breathable low-pressure oxygen (at approximately 70 psi) to crew and passengers through individual oxygen masks. The system uses a single bottle of compressed oxygen to supply both crew masks and passenger masks. A regulator controls overall system pressure, and a shutoff valve (controlled by an oxygen supply valve labeled OXYGEN SUPPLY in the cockpit) enables or disables the system. Another cockpit control labeled OXYGEN CONTROL VALVE selects distribution modes.

The shutoff valve on the bottle is normally open in flight. It is mechanically controlled in the cockpit by the OXYGEN SUPPLY control knob.

Crew Oxygen Masks Each crewmember is supplied with a quick-donning mask with a built-in microphone and regulator (Figure 17-2). Each oxygen mask is stowed immediately outboard and aft of each crewmember in a container above the outboard shoulder of each crewmember (Figure 17-3) and is equipped with an inline flow indicator (Figure 17-4). A flow indicator indicates to the crew that oxygen is received.

An oxygen gauge indicates the pressure (and indirectly, volume) of oxygen in the bottle. A crew alerting system (CAS) message indicates when insufficient oxygen is available. If oxygen supply is shut off or if pressure in the system is too low, an amber OXYGEN OFF message appears. Individual controls on crew masks adjust their oxygen flow.

COMPONENTS The system includes: • Oxygen bottle (with integral shutoff valve and pressure regulator) • Oxygen masks (crew and passenger) • Oxygen control valve

Figure 17-2.   Standard Crew Masks

Oxygen Bottle A single bottle holds all compressed oxygen for the system. It is on the right side of the nose storage compartment and has a 623-liter (22-cubic-feet) useable capacity with 1,133-liter (40-cubic-feet) option. Oxygen is stored in the bottle at a pressure between 1,600 and 1,800 psig. A shutoff valve and pressure regulator on the bottle control the flow of oxygen to the distribution system (Figure 17-1). The regulator reduces oxygen system pressure to 70 psi downstream of the bottle.

Figure 17-3.   Crew Oxygen Mask, Stowed (Pilot Side)

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The crew masks plug into OXYGEN MASK receptacles on the pilot and copilot side consoles (Figure 17-5). The mask oxygen line plugs into the large valve port, and the mask microphone plugs into the MIC jack, both of which are under the OXYGEN MASK section of the console. Ensure both plugs are fully inserted before flight. 17 OXYGEN SYSTEM

If the aircraft is to be parked outside and the temperature is colder than 0°C, the masks must be removed from the aircraft and kept warm. Figure 17-4.   Flow Indicator

The mask is quick-donning by pressing the red sides of the nosepiece, which causes the harness to inflate and easily slip over the head. The mask is a diluter/pressure-demand type with 100% oxygen provided by pushing a lever/tab on the bottom-right corner of the mask to the 100% position. To qualify as a quick-donning mask, the crew oxygen mask must be properly stowed in the receptacle behind, above, and outboard of each crewmember on the forward cabin divider, and must be set to 100%.

Passenger Oxygen Masks In the cabin, passenger masks are in overhead containers and drop automatically or manually (Figure 17-6). A lanyard attached to the mask aids in pulling the mask down if it does not drop clear of the box. A short lanyard physically connects the mask to a pin in a valve inside the overhead oxygen line. Pulling this lanyard pulls out the pin to start the oxygen flow to the mask. The act of lowering the oxygen mask to the face also pulls free the lanyard and pin, enabling oxygen to flow. Passenger masks have no flow indicator.

NOTE Headsets, eyeglasses, or hats worn by the crew will interfere with the quickdonning capability of the oxygen masks.

PILOT CONSOLE

COPILOT CONSOLE

Figure 17-5.   Pilot and Copilot Consoles

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The OXYGEN CONTROL VALVE knob is on the far left edge of the instrument panel (Figure 17-8). It controls oxygen flow to the passenger cabin. Its three positions actuate a control valve for passenger oxygen distribution as desired. The knob positions are: • CREW ONLY • NORMAL • DROP MASK

Figure 17-6.   Passenger Oxygen Mask

CONTROLS AND INDICATIONS OXYGEN CUTOFF Knob The OXYGEN CUTOFF cutoff knob is on the lower-right corner of the instrument panel, below the OXYGEN pressure gauge (Figure 17-7). It closes the regulator at the bottle for delivery of oxygen to the crew and passengers. When the OXYGEN SUPPLY knob is placed in the CUTOFF position, line pressure is vented overboard.

Figure 17-8.   Oxygen Control Valve Knob

CREW ONLY Mode When the control knob of the oxygen control valve is placed in the CREW ONLY position, oxygen is not available to the passengers. In this position, oxygen is only available to the crew. After donning the mask, the crew must set the lever under each mask to NORMAL or 100%. For pressure breathing, rotate the mask emergency select knob to EMERGENCY.

NOTE

Figure 17-7.   OXYGEN CUTOFF Knob

Oxygen masks are certified to 40,000 feet cabin altitude for the crew only. The CREW ONLY mode operates with or without DC power.

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17 OXYGEN SYSTEM

OXYGEN CONTROL VALVE Knob

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NORMAL Mode When the oxygen system is enabled, if the OXYGEN CONTROL VALVE knob is selected to the NORMAL position (see Figure 17-8), passenger oxygen masks automatically drop down from the cabin ceiling anytime cabin pressure altitude is greater than 14,800 feet. 17 OXYGEN SYSTEM

Normally, the pressurization system maintains an 8,000-foot cabin altitude up to the maximum certified altitude. However, if cabin altitude exceeds approximately 14,800 feet, a cabin-altitude sensor energizes the passenger oxygen solenoid valve open. Oxygen flows into the passenger distribution system and releases latches on the mask compartment doors. This allows the doors to open and the masks to fall out. After restoration of the cabin pressure to normal values, the solenoid valve deenergizes closed at approximately 11,500 feet cabin altitude, shutting off oxygen flow to the passengers. The pilot can bypass the solenoid valve by selecting the OXYGEN CONTROL VALVE knob to DROP MASK.

DROP MASK Mode The pilot can supply oxygen to the passengers at any cabin altitude by placing the OXYGEN CONTROL VALVE selector to the DROP MASK position (see Figure 17-8). When this position is selected, all masks in the cabin to immediately drop from the cabin overhead. This mode operates with or without DC power. When oxygen flow to passengers is not desired, shut off oxygen flow to passenger masks by selecting the OXYGEN CONTROL VALVE knob to the CREW ONLY position at any time, or the NORM position when below 11,500 feet cabin altitude.

MIC Switches (Left And Right) The left and right MIC switches are immediately below and inboard of each control yoke shaft on the lower instrument tilt panel. There is one switch for each crewmember. Each switch has two positions: OXY MASK and HEADSET. Normally, each switch is set to the HEADSET position, which selects crew audio input to the avionics system from the microphones in the crewmember headset (Figure 17-9).

17-6

Figure 17-9.   Mic Switches

Selecting a MIC switch to OXY MASK selects audio input from the microphone in that crew oxygen mask, and disables audio input from that crewmember headset microphone. Depressing the microphone button on the respective control wheel allows the crewmember to transmit through the headset microphone or through the oxygen mask microphone, as selected by the respective MIC switch. When the switch is in the OXY MASK position, the cockpit speaker turns on and cannot be turned off using the audio panel button.

OXY CONTROL Circuit Breaker The OXY CONTROL circuit breaker in the ENVIRO section of the pilot CB panel protects the passenger oxygen solenoid valve. Pulling this circuit breaker disables the spring-loaded solenoid, which closes the valve.

OXYGEN Pressure Gauge The OXYGEN pressure gauge (Figure 17-10) is on the right side of the copilot instrument panel, below and to the right of the copilot primary flight display (PFD), and above the OXYGEN CUTOFF knob (see Figure 17-7). The gauge illuminates internally. The range markings are as follows: • Yellow arc .................................... 0–400 psi • Redline .......................................... 2,000 psi

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Crew Oxygen Mask Controls And Indications Each crew mask has the following controls (Figure 17-12): • Harness inflation plate 17 OXYGEN SYSTEM

• N–100% diluter rocker switch • Emergency select knob • PRESS to TEST button • Vent valve Figure 17-10.   Oxygen Pressure Gauge

• Flow indicator

Service anytime gauge indicates insufficient volume. Refer to the Airplane Flight Manual (AFM); normal indication is between 1,600 and 1,800 psig.

Overboard Discharge Indicator A green overboard discharge indicator (disc) is on the right side of the nose section directly below the nose access door (Figure 17-11). If the disc is ruptured, the oxygen bottle has experienced overpressure and is now empty. If the disc ruptures, perform maintenance before flight.

Figure 17-12.   Crew Mask Controls

Harness Inflation Plate

Figure 17-11.   Overboard Discharge Indicator

The red harness inflation plate is a mechanical valve control on the lower-left corner of the mask, which controls inflation of the harness. Squeezing the plate against the mask causes a momentary flow of pressurized oxygen to the harness. This inflates the harness, which expands to allow the crewmember to slip the mask harness over their head. When the plate is released, the pressure is released from the harness, which then contracts to hold the mask firmly to the face of the crewmember.

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N–100% Diluter Rocker Switch The red N–100% diluter rocker switch is a mechanical valve on the lower-right corner of the mask. The switch controls the dilution of oxygen supplied by the mask to the crewmember:

17 OXYGEN SYSTEM

• N (normal diluted oxygen)—Forward switch position. Reduces the rate of oxygen usage by mixing oxygen with normal cockpit air at a ratio determined by cabin altitude. • 1 00% (pure oxygen)—Aft switch position. Provides only pure oxygen from the oxygen bottle. No cockpit air is mixed with the flow. The mask is required to be set to 100% and checked prior to flight in order to qualify as a quick-donning mask.

EMERGENCY–PRESS TO TEST Knob The red EMERGENCY–PRESS TO TEST knob/ button is a mechanical valve on the underside of the mask. The knob controls the pressure of oxygen supplied by the mask to the crewmember and the button is pressed to test and check if oxygen is available: • E MERGENCY position (clockwise, toward crewmember)—Provides oxygen under positive pressure, regardless of crewmember breathing. • D emand-breathing position (not labeled; counterclockwise, away from crewmember)— Provides oxygen on demand as determined by crewmember breathing. This is the normal setting. • P RESS TO TEST function—Spring-loaded button in the center of the knob. Pressing the button on the knob causes a positive pressure and flow of oxygen to the mask until the button is released.

Vent Valve When smoke goggles are worn, they fit over the vent on the top of the mask nosepiece. A vent valve control on the front of the mask nosepiece slides forward to open the vent to allow oxygen to enter and clear the smoke goggles.

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NOTE Open the vent valve only if pressure breathing (EMERGENCY position) has been selected. If the vent is opened when the mask is in the other position (demand breathing), smoke may be drawn into the mask.

Flow Indicator A flow indicator (slide in the mask hose near the connector to the oxygen panel) shows clear when oxygen is available to the mask and is flowing. It shows black when there is no flow.

OPERATION For specific, current instructions on normal operations, refer to the AFM. Where the following information differs from the AFM, use the AFM information and follow the AFM instructions. The following information is only for training and background information.

WARNING Strictly obey the procedures for the use of oxygen equipment. Do not use oil, grease, or other lubricants made from petroleum in the area of oxygen equipment. This can cause a dangerous fire hazard.

Preflight During preflight, ensure the OXYGEN SUPPLY control knob is fully pushed in (forward) to open the shutoff valve on the oxygen bottle. Check that proper pressure is indicated on the OXYGEN gauge. Test each crew mask before flight using the PRESS TO TEST button to be sure that it is receiving oxygen from the system. Ensure that oxygen flows into the mask and to the pilot under positive pressure. Before takeoff, check that the OXYGEN OFF message is not displayed and the OXYGEN VALVE is in the NORMAL position.

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• T he OXYGEN gauge indicates adequate supply (refer to AFM) • T he OXYGEN SUPPLY knob is pushed in (forward) fully • T he OXYGEN OFF message does not appear When those conditions are met, the oxygen system can be operated in one of three modes as selected by the pilot using the OXYGEN CONTROL VALVE knob.

Crew Oxygen Mask Remove the crew oxygen mask from its container and squeeze the mask so the harness inflation plate is pressed against the mask to inflate the harness. Place the harness over the head and position the mask over the face and nose, then release the harness inflation plate. The harness contracts to hold the mask in place. The crewmember is assured that oxygen is being received when no restriction to breathing is present with the mask donned and the red N–100% diluter rocker switch is set to 100% (aft position). If the cabin altitude is at or below 25,000 feet, to conserve oxygen when using the mask, the diluter rocker switch may be set to normal (N).

NOTE On crew masks, select 100% oxygen above 25,000 feet cabin altitude. At cabin altitudes of 25,000 feet and below, select normal (N). For pressure breathing or smoke/fumes protection, rotate the emergency select knob on the underside of the mask clockwise toward the crewmember to the EMERGENCY position (see Figure 17-10). This position provides a steady flow of pressurized oxygen to the face cone and the smoke goggles (if installed).

Maintenance Considerations Service the oxygen system any time the pressure gauge indicates inadequate supply, or when the overboard discharge indicator shows an overpressure event has occurred. If the oxygen bottle depletes to empty or if the oxygen discharge indicator ruptures, the system must be purged and the oxygen bottle replaced before the next flight. The original oxygen bottle must be returned to the supplier for refurbishment or replacement before further use. Service the oxygen bottle through the filler port near the forward bulkhead, inside the right nose baggage door (Figure 17-13). Only use aviator oxygen (MIL-O-27210, Type 1) for servicing. The fill valve incorporates a check valve and filter. A pressure sealing cap prevents contaminants from entering the oxygen system.

Figure 17-13.   Oxygen Bottle

LIMITATIONS Table 17-1 indicates approximate normal duration of oxygen supply with different numbers of users.

WARNING Due to human physiological limitations, the passenger oxygen system is not satisfactory for continuous operation above 25,000 feet cabin altitude. The crew oxygen system is not satisfactory for continuous operation above 40,000 feet cabin altitude. Individual physiological limitations may vary. If crew or passengers experience hypoxia symptoms, descend to a lower cabin altitude.

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17 OXYGEN SYSTEM

In flight to operate the oxygen system, ensure there is adequate pressure in the system as indicated by these three conditions:

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Table 17-1.   OXYGEN SUPPLY DURATION

OXYGEN SUPPLY CHART

22 FT3

AVAILABLE TIME IN MINUTES 2 COCKPIT

2 COCKPIT, 1 CABIN

2 COCKPIT, 2 CABIN

8000 10,000 15,000 20,000

196 225 220 178

74 78 79 73

46 47 48 46

33 34 35 34

26 27 27 27

25,000 30,000 35,000 40,000

98 129 175 246

55

39

30

24

17 OXYGEN SYSTEM

CABIN ALTITUDE

2 2 COCKPIT, COCKPIT, 3 CABIN 4 CABIN

AVAILABLE TIME IN MINUTES CABIN ALTITUDE

1 COCKPIT

1 COCKPIT, 1 CABIN

1 COCKPIT, 2 CABIN

1 COCKPIT, 3 CABIN

1 COCKPIT, 4 CABIN

8000 10,000 15,000 20,000

392 450 440 356

91 95 96 92

52 53 54 53

36 37 37 37

28 28 29 29

25,000 30,000 34,000 40,000

197 258 350 492

77

48

35

27

WARNING No smoking when oxygen is being used or following use of passenger oxygen until lanyards have been reinstalled.

CAUTION Oil, grease, soap, lipstick, lip balm, and other fatty materials constitute a serious fire hazard when in contact with oxygen. Oxygen use limitations are further governed by the applicable regulations. In the U.S.A., the pilot must have the oxygen mask on his face during normally pressurized flight for single-pilot Part 135 operations above FL 250 or above single-pilot Part 91 operations above FL 350.

EMERGENCY/ ABNORMAL For specific information on emergency/abnormal procedures, refer to the appropriate checklist or FAA-approved AFM. Table 17-2.   CAS MESSAGE OXYGEN OFF DESCRIPTION

INHIBITS

LIMITATIONS

This message indicates the oxygen system pressure-sensor switch detects system pressure below approximately 45–50 psig. The message extinguishes if system pressure rises above 50–55 psig. This message also displays when the OXYGEN SUPPLY knob is in the PULL TO CUTOFF position. EMER, LOPI, TOPI

For specific limitations, refer to the FAA-approved AFM.

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1. When using the oxygen mask for smoke protection, the mask regulator should be set for: A. EMER B. NORM C. 100% oxygen D. STANDBY

5. After donning the oxygen mask, the mask microphone should be set for: A. L or R MIC switches—MASK B. NORM C. STANDBY D. EMER

2. With a total DC power failure, oxygen is: A. Available to passengers via the OXYGEN CONTROL VALVE in the DROP MASK position B. Available to passengers via the OXYGEN CONTROL VALVE in the NORMAL mode C. Available to passengers via the OXYGEN CONTROL VALVE in the CREW ONLY position D. Available to occupants regardless of the position of the OXYGEN CONTROL VALVE

6. The cockpit oxygen pressure gauge reads: A. The oxygen pressure, which is present at the crew masks B. Electrically derived system low pressure C. Bottle pressure E. Electrically derived system high pressure

3. With the amber CAS message OXYGEN OFF displayed, required pilot action is to: A. OXYGEN SUPPLY knob—PUSH IN B. OXYGEN SUPPLY knob—PULL OUT C. Rotary TEST knob—To ANNU D. Push the TMR/REF soft key 4. Pilot action required with a red CABIN ALT message displayed is to: A. Oxygen masks—DON and 100% OXYGEN B. OXYGEN SUPPLY VALVE—PULL OUT C. OXYGEN SUPPLY VALVE—CREW ONLY D. OXYGEN SUPPLY VALVE— Reversionary mode

7. During the walkaround, the pilot notices the green overboard discharge indicator is missing, this indicates: A. The oxygen bottle has discharged itself due to overpressure B. The nitrogen blowdown bottle has thermally discharged itself C. The hydrogen bottle has thermally discharged itself D. The halon bottle has thermally discharged itself 8. Passenger masks are dropped when: A. The OXYGEN CONTROL VALVE is in NORMAL, normal DC power available, and cabin altitude exceeds 14,800 feet B. The cabin altitude exceeds 13,500 feet, regardless of OXYGEN selector position C. The OXYGEN selector is in DROP MASK, regardless of altitude D. When cabin altitude exceeds 10,500 feet

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17 OXYGEN SYSTEM

QUESTIONS

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17 OXYGEN SYSTEM

9. On descent, oxygen flow to the passenger masks is automatically: A. Established B. Set to the NORMAL mode C. Terminated as cabin altitude descends through 11,500 feet D. Terminated as cabin altitude descends through 14,800 feet 10. The crew oxygen masks are certified up to: A. 25,000 feet cabin altitude B. 30,000 feet cabin altitude C. 35,000 feet cabin altitude D. 40,000 feet cabin altitude 11. The flight crew oxygen masks need to be removed from the airplane if the temperature is forecast to be below: A. 10°C B. 0°C C. –10°C D. –20°C 12. Prolonged use of passenger oxygen masks above ________ feet cabin altitude is not allowed: A. 20,000 B. 25,000 C. 30,000 D. 35,000

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CHAPTER 18 MANEUVERS AND PROCEDURES CONTENTS INTRODUCTION................................................................................................................ 18-1 PERFORMANCE................................................................................................................. 18-2 Takeoff and Landing Speeds.......................................................................................... 18-2 Weights.......................................................................................................................... 18-2 FLIGHT OPERATIONS....................................................................................................... 18-3 Preflight and Taxi........................................................................................................... 18-3

AIRWORK MANEUVERS.................................................................................................. 18-6 Steep Turns.................................................................................................................... 18-6 Miscellaneous................................................................................................................ 18-6 LIMITATIONS...................................................................................................................... 18-7 EMERGENCY/ABNORMAL.............................................................................................. 18-7 APPROACH TO STALL TRAINING REQUIREMENTS .................................................. 18-8 Training Scenarios......................................................................................................... 18-8 Checking / Testing Requirements.................................................................................. 18-8 STALL RECOVERY PROFILES......................................................................................... 18-9 STALL RECOVERY RATIONALE.................................................................................. 18-12 SIMULATOR TRAINING GUIDANCE........................................................................... 18-13 Initial Training Course................................................................................................ 18-13 Simulator Session # 1................................................................................................. 18-13 Simulator Session # 5................................................................................................. 18-13 Simulator Session #7.................................................................................................. 18-13

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18 MANEUVERS AND PROCEDURES

Takeoff........................................................................................................................... 18-5

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RecurrentTraining Course.......................................................................................... 18-13 Simulator Session # 1................................................................................................. 18-13 Simulator Session # 2................................................................................................. 18-13

18 MANEUVERS AND PROCEDURES

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ILLUSTRATIONS 18-1.

Takeoff and Landing Card...................................................................................... 18-3

18-2.

Approach to Stall - Enroute Configuration............................................................ 18-9

18-3.

Approach to Stall - Takeoff Configuration.......................................................... 18-10

18-4.

Approach to Stall - Landing Configuration........................................................ 18-11

18-5.

Takeoff - Normal................................................................................................. 18-14

18-6.

Takeoff - Engine Failure (Speed Below V1)....................................................... 18-15

18-7.

Takeoff - Engine Failure (At or Above V1)......................................................... 18-16

18-8.

VFR Approach - Normal.................................................................................... 18-17

18-9.

VFR Approach - Single Engine.......................................................................... 18-18

18-10. ILS Approach - Normal...................................................................................... 18-19 18-11. Nonprecision Approach - Normal....................................................................... 18-20 18-12. Nonprecision Approach - Single Engine............................................................ 18-21 18-13. Missed Approach - 2 Engine (Precision/Nonprecision)..................................... 18-22 18-14. Missed Approach - Single Engine (Precision/Nonprecision)............................. 18-23 18-15. Visual Approach................................................................................................. 18-24 18-16. Steep Turns.......................................................................................................... 18-25

TABLES Table Title Page 18-1.

MINIMUM MANEUVERING SPEEDS.............................................................. 18-2

18-2.

EXAMPLE CALLOUTS (IFR AND VFR)........................................................... 18-4

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18 MANEUVERS AND PROCEDURES

Figure Title Page

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18 MANEUVERS AND PROCEDURES

INTENTIONALLY LEFT BLANK

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18 MANEUVERS AND PROCEDURES

CHAPTER 18 MANEUVERS AND PROCEDURES

INTRODUCTION This chapter contains information and flight profiles likely to be encountered during training and in most daily flight operations. These procedures are consistent with the Cessna Model 510 Citation Mustang Aircraft Flight Manual (AFM) and pilot abbreviated checklists.

GENERAL The flight profiles in this chapter show some normal and emergency operating procedures. They are a general guide for training purposes. Actual in-flight procedures may differ due to aircraft con-

figuration, weight, weather, traffic, ATC instructions, etc. Procedures are consistent with the AFM. If a conflict develops between these procedures and the AFM, then AFM procedures must be followed.

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PERFORMANCE The Mustang performance is certified to Part 23 Normal Category with FAA Special Conditions to meet Commuter Category (14CFR Part 23) takeoff and landing performance requirements. The following areas will help to familiarize the pilot with terms in the AFM and to help the pilot understand the capabilities of the aircraft.

TAKEOFF AND LANDING SPEEDS Refer to the FAA-approved Cessna Model 510 Citation Mustang pilot abbreviated checklists or AFM for takeoff and landing speeds.

18 MANEUVERS AND PROCEDURES

V1 (takeoff decision speed)—The distance to continue the takeoff to 35 feet will not exceed the scheduled takeoff field length if recognition occurred at V1 (accelerate-go). The distance to bring the aircraft to a full stop (accelerate-stop) will not exceed the scheduled takeoff field length provided that the brakes are applied at V1. VR (rotation speed)—The speed at which rotation is initiated during takeoff to attain the V2 climb speed at or before a height of 35 feet above runway surface has been reached. V2 (takeoff safety speed)—This climb speed is the actual speed at 35 feet above the runway surface as demonstrated in flight during takeoff with one engine inoperative. V ENR (enroute climb speed)—Single-engine enroute climb speed. VREF (minimum final approach speed)—The airspeed equal to the landing 50-foot point speed with flaps LAND (anti-ice OFF) or with flaps 15 (antiice ON) and landing gear extended. For emergency and abnormal procedures, the airspeed equal to the 50-foot point speed with flaps in landing position as defined and landing gear extended. VAPP (missed approach climb speed)—The landing approach climb speed (1.3 VS1) with 15 flap position, landing gear up.

18-2

Speeds are posted on the airspeed tape for reference during takeoff, approach, and landing. The V-speed flags on the G-1000 may be selected ON or OFF at the discretion of the pilot. Minimum maneuvering speeds provide a safety margin above stall speed (for current flap setting and weight) when maneuvering prior to establishing a stabilized final approach. Table 18-1 lists minimum maneuvering speed. Table 18-1.   MINIMUM MANEUVERING SPEEDS FLAP CONFIGURATION

CITATON CE-510

UP

VREF + 30

TO/APR

VREF + 20

LAND

VREF + 20

WEIGHTS Maximum takeoff weight is limited by the most restrictive of: 1. Maximum cer tif ied takeoff weight ( F l a p s — TA K E O F F / A P P R O A C H ) 8,645 pounds 2. Maximum weight permitted by climb requirements 3. Maximum weight permitted by takeoff field length Takeoff weight may be further limited by obstacle clearance requirements of a departure runway or procedure, or by the landing weight restrictions at destination. Maximum landing weight is limited by the most restrictive of: 1. Maximum certified landing weight—8,000 pounds 2. Maximum weight permitted by climb requirements or brake energy limits 3. Maximum weight permitted by landing distance

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Sample Takeoff Briefing

PREFLIGHT AND TAXI

“This will be a static (or rolling) takeoff with flaps at TO/APR. Check takeoff power and call “speed alive, 70 knots, V1 and rotate.” I will call for gear up, flaps, and yaw damper. The departure is _____. Call abort for any malfunction below V1. I will control the aircraft and extend the speedbrakes and notify the tower. Between 70 and V1 we will only abort for red lights, loss of directional control or loss of major displays. After V1 we will handle all problems in flight. We will climb to _____ feet before doing any actions. I will fly and talk to ATC, and you can then get into the checklist. If I do not respond to you or I do something dangerous, assume controls and we will sort it out later. Any questions or comments?”

If flying as a crew, the pilot-in-command ensures that the copilot understands the normal and emergency procedures to be used for that takeoff. This includes verbal callouts during takeoff roll and initial climb (refer to Table 18-2).

If flying as a single pilot, the pilot in command (PIC) does not perform any checklist items while the aircraft is taxiing. The only flight instrument check to perform while moving is a check of heading changes and movement of the slip indicator.

Some flight departments use preprinted cards for computations, ATIS and clearances. Sample takeoff and landing data (TOLD) cards are shown in Figure 18-1.

FLIGHT OPERATIONS Sample flight profiles are shown in Figures 18-2 through 18-17.

FlightSafety international

F lightS afety international

Figure 18-1.   Takeoff and Landing Card

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18 MANEUVERS AND PROCEDURES

Landing weight may be further limited by obstacle clearance requirements of a missed approach procedure.

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Table 18-2.   EXAMPLE CALLOUTS (IFR AND VFR) PHASE

CONDITION

CALLOUT

Takeoff

All airpseed indicators moving

“Airpseed alive”

All airspeed indicators indicating 70 KIAS

“70 knots”

Airspeed indicators at computer V1

“V1”

Airpseed indicators at computed VR

“Rotate”

Airspeed indicators at computed V2

“V2”

Departure/Enroute/Approach

Prior to intercepting an assigned course

“Course alive”

Climb and descent

Approaching transition altitude (IFR and VFR)

“Transition altitude altimeters reset”

1,000 feet above/below assigned altitude (IFR)

State altitude leaving and assigned leveloff altitude

At final approach fix

(Fix) altimeters and instruments check (NOTE 1)

500 feet above minimums

“500 above minimums”

100 feet above minimums

“100 above minimums”

Runway acquisition

“Runway at (clock position)” or “Approach lights at (clock position)” (NOTE 2)

Final

18 MANEUVERS AND PROCEDURES

After pilot flying reports “visual,” pilot not flying “VREF” reverts to instruments and callouts “Sink (rate of descent)” “On,” “Above,” or “Below glide slope,” if required At decision altitude (DA)

“Minimums, runway, not in sight” or “Minimums, runway at (clock position)” or “Minimums, approach lights, at (clock position)” (NOTE 2)

At minimum descent altitude (MDA)

“Minimums” (NOTE 2)

At missed-approach point (MAP

“Missed-approach point, runway not in sight” or “Missed-approach point, runway at (clock position)” or “missed-approach point, approach lights, at (clock position)”

NOTES: 1. CHECK FOR APPEARANCE OF WARNING FLAGS AND GROSS INSTRUMENT DISCREPANCIES 2. CARE MUST BE EXERCISED TO AVOID MAKING AMBIGUOUS CALLOUTS THAT COULD NEGATIVELY INFLUENCE THE PILOT FLYING, RESULTING IN A COMPROMISE OF SAFETY. 3. PILOTS FLYING UNDER A SINGLE-PILOT TYPE RATING MAY WANT TO MENTALLY OR VERBALLY ANNOUNCE THE RESPONSES TO THE RESPECTIVE “CONDITIONS” AS MENTIONED ABOVE AS A MATTER OF COURSE REGARDLESS OF WHETHER THEY ARE OPERATING AS A SINGLE PILOT OR AS A CREW.

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TAKEOFF Normal It is recommended to use the flight director during takeoff. Press the TO/GA button on the left throttle, then select the HDG mode. After lining up on centerline, press the heading knob. Advance power to takeoff detent. At V1, remove your hand from the throttles to the yoke and rotate at VR toward the command bars. With a positive rate of climb, raise the gear; raise flaps no earlier than V2 + 12.

Rejected (Before V1) Simultaneously apply the brakes, retard the throttles to idle, and extend the speedbrakes. Move the throttles to cutoff if runway departure is imminent. Notify the tower and accomplish any other memory items needed.

Engine Failure (After V1) Maintain directional control, accelerate to and rotate at VR, climb at V2 and with a positive rate of climb, retract the landing gear. At 1,500 feet AGL, retract the flaps at V2 + 10 and accelerate to VENR. A small amount of aileron into the good engine (pick up the dead engine) is needed to keep the wings levels (the yoke will be displaced). Use minimum safe, minimum enroute, or ATC assigned altitudes. Rudder trim may be used. If further climbs are needed, use computed VENR. Retrim rudder and aileron as needed as speed increases.

Climb Ensure gear and flaps are up, set power to climb detent and select autopilot (if desired). Continue the climb at desired climb speed until nearing the assigned cruise altitude. Once level, allow the aircraft to accelerate to the desired cruise airspeed/ Mach. Complete appropriate checks (refer to the AFM).

Cruise NOTE The throttles should be reduced to the CRU detent or below within 10 minutes after reaching an intermediate or final cruise altitude. The use of CLB during normal operations beyond 10 minutes after reaching cruise altitude will significantly decrease engine life and increase operator costs.

Descent Complete the appropriate descent checklist to include checking ATIS and, programming the G1000 for the arrival, approach and landing runway. Review the planned approach and missed approach and cross-check the flight plan page on the MFD to include headings, courses, altitudes, DA/MDA and MAP procedures. Complete the actual approach briefing before the top of descent. Begin arrival/approach tasks. Complete appropriate checks.

Approach and Landing Ensure proper navigation aids are set for planned approach. Ensure that proper navigation aids and navigation presentations are set, tuned, and identified for the planned approach.

Sample Approach Briefing “We are flying the _____ approach to RWY _____. CDIs and bearing pointers are set to _____ and____. V speeds and DA/MDA are set in the PFD. Gear and flaps to be set by the FAF. Call out 1,000 feet, 500 feet, and 100 feet above minimums. Advise location of the runway. I’ll call visual or missed. MAP is ____ to ____ and hold at ____. Any questions?” Single-pilot operators should consider review of the same procedures.

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18 MANEUVERS AND PROCEDURES

The single-pilot operator should mentally review the briefing as though working in a crew environment.

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When leveling off at an intermediate approach altitude, lead the level-off height with enough power to maintain the desired speed for the maneuver. As you configure the aircraft, speed will decrease. Plan to reach the glideslope (GS) intercept or final approach fix (FAF) with the landing gear down, flaps set, and speed set. If flying a straight-in twoengine approach, plan to have flaps set at LAND by the FAF; this permits a stabilized approach throughout final. If flying a one-engine approach, use flaps TO/APR on final. Decide early if the landing will be with flaps TO/APR or LAND; ensure sufficient runway is available for reduced flaps. Landing with flaps TO/APR allows for a stabilized approach throughout final. If circling to land, plan to fly the approach with flaps TO/APR until you decide landing is assured; then select LAND.

18 MANEUVERS AND PROCEDURES

Plan to arrive over the threshold at VREF for the flap setting desired at 50 feet above the runway with the yaw damper off. Idle power can then be selected. Following a normal landing, deploy speedbrakes and apply wheel brakes simultaneously. When clear of the runway, accomplish the after landing checks.

After Landing If flying as a crew, the checks may be performed while taxiing. If flying as a single pilot, taxi the aircraft clear of the runway, stop, and complete the after-landing check. No checklists are to be read while the aircraft is taxiing.

AIRWORK MANEUVERS STEEP TURNS Steep turns are flown at 45° of bank and 200 knots. Establish a base heading and altitude. Maintain the altitude during the maneuver and use the base heading for the turn reversal and final roll out. Use of the flight director and elevator trim is an option for the pilot. A pitch attitude of about 2.5° should hold level flight in the turns. A slight increase in N1 (approximately 2%–4%) is required to maintain target airspeed. If a moderate roll in rate is used to begin

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the maneuver, plan to use a approximately 20° heading lead point for reversing the turn and for the final roll out.

MISCELLANEOUS Takeoff and Landing For takeoff, line up as close to the end of the runway as possible. While holding the brakes, advance the throttles toward the TO detent, ensure that both engines are accelerating together, and release the brakes. As the aircraft accelerates, monitor the V speeds and rotate the aircraft for takeoff VR. The landing maneuver is preceded by a steady three-degree approach down to the 50-foot height point with airspeed at VREF in the landing configuration. At 50-feet, idle thrust is selected and the descent is continued into the flare, establishing a landing attitude. Once on the ground, maximum wheel braking is initiated after nose wheel contact and continued until the aircraft is stopped. While speedbrakes would normally be extended after landing, they were not utilized during the certification process.

Touch-and-Go Landings If doing touch-and-go landings. Consider using only TO/APR flaps on those landings; no need to change flaps on roll. If using LAND flaps for the landings, consider just holding the nose wheel off the runway while the other pilot sets the flaps to TO/APR. If power is added before the flaps are reset, airspeed will be higher than normal at liftoff.

Wheel Fusible Plug Considerations Brake application reduces the speed of an aircraft by means of friction between the brake stack components. The friction generates heat, which increases the temperature of the brake and wheel assembly, resulting in an increased tire pressure. Each main wheel incorporates three fuse plugs, which melt at a predetermined temperature to prevent a possible tire explosion due to excessively high tire pressure. Flight crews must take precautions when conducting repetitive traffic patterns, including multiple landings/or multiple rejected takeoffs, to prevent

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CITATION MUSTANG PILOT TRAINING MANUAL

overheating the brakes, which could melt the fuse plugs and cause loss of all tire pressure and possible tire and wheel damage. During such operations, available runway permitting, minimize brake usage and consider cooling the brakes in flight with the landing gear extended. Extending speedbrakes will assist in bringing the aircraft to a stop.

Adverse Runway Conditions Ensure the proper performance charts are used when taking off or landing on runways with adverse conditions. If the chart does not cover your particular situation, strongly consider not doing it. Hydroplaning occurs at 9.0 times the square root of the tire pressure for a water-covered runway. Approximate speeds equate to 85–90 knots.

18 MANEUVERS AND PROCEDURES

If landing or taxiing on slush, inspect drains, control surfaces, and wheels after shutdown.

Cold Weather Comply with the cold weather operations outlined in the AFM, Section 3.

Servicing Comply with fluid requirements outlined in the AFM, Section 2.

LIMITATIONS For specific limitations, refer to the FAA-approved AFM.

EMERGENCY/ ABNORMAL For specific information on emergency/abnormal procedures, refer to the appropriate checklist or FAA-approved AFM.

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APPROACH TO STALL TRAINING REQUIREMENTS Approach to stalls shall be done with and without the autopilot, in both VMC and actual or simulated IMC conditions, with and without a bank, and in realistic scenarios at different altitudes. When possible, it should be accomplished so that the client is surprised by the stall. Only the client’s ability to recognize and properly recover from an impending stall should be evaluated.

TRAINING SCENARIOS

It should be noted that smooth aircraft control on the entry should be maintained as an evaluation of the client’s general aircraft handling.

• Autopilot disengaged

It should also be noted that stall training should be conducted in a variety of different aircraft configurations and under a number of different flight scenarios. 18 MANEUVERS AND PROCEDURES

Stall recovery procedures are based on aircraft configuration; the recovery profiles in this training package include: • Enroute (Clean) Configuration Stall

1. Enroute (Clean) Configuration Stall A. High Altitude • Conducted within 5000 ft of the operations ceiling for the aircraft B. Manual Flight Conditions C. Automated Flight Conditions • Autopilot engaged 1. Takeoff Configuration Stall A. If there are multiple take off flap settings for the aircraft, stalls training should include different flap settings B. Aircraft bank • 15 to 30 degrees of bank 2. Landing Configuration Stall A. Aircraft descent

• Takeoff Configuration Stall

DEMONSTRATION SCENARIOS

• Landing Configuration Stall In order to best prepare pilots for inadvertent stall events during normal operations, the training of these configuration stalls should be conducted as maneuvers training and scenario based training.

1. AOA Reduction Recovery Demonstration A. Demonstration of stall recovery using AOA reduction only, without use of power.

CHECKING / TESTING REQUIREMENTS As outlined in the PTS and/or FSB Report 1. Enroute (Clean) Configuration Stall 2. Takeoff Configuration Stall 3. Landing Configuration Stall

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STALL RECOVERY PROFILES

1. MINIMUM ALTITUDE – AS REQUIRED 2. THRUST – IDLE 3. AUTOPILOT AND FD – AS DESIRED

APPROACH AND RECOVERY

AERODYNAMIC BUFFET, AND/OR ROLL-OFF, WHICHEVER OCCURS FIRST 1. 2. 3. 4. 5.

COMPLETION OF MANEUVER

1. RETURN AIRCRAFT TO DESIRED FLIGHTPATH

AUTOPILOT – DISCONNECT PITCH ATTITUDE – REDUCE ROLL ATTITUDE – LEVEL THROTTLES – TO DETENT AIRSPEED – INCREASE

AT STALL INDICATION

18 MANEUVERS AND PROCEDURES

TRAINING SET UP FOR MANEUVER

AT OR ABOVE VREF

Figure 18-2.   Approach to Stall - Enroute Configuration

Training execution: 1. The instructor sets up the stall scenario. 2. The entry altitude should be consistent with the expected operational environment for the stall configuration. 3. For training and evaluation, the maneuvers may be accomplished with the autopilot on or off as directed by the instructor. 4. The standard is based on the demonstration of smooth, positive control during entry, approach to stall, and recovery. The aim of these stall profiles is to familiarize the pilot with the stall characteristics and to train recognition and recovery procedures in accordance with the ATP Practical Test Standards in flight simulator training only. These stall profiles are not intended for maintenance test flights or aircraft training.

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TRAINING SET UP FOR MANEUVER

1. MINIMUM ALTITUDE – AS REQUIRED 2. THRUST – IDLE 3. FLAPS – TAKEOFF & APPROACH 4. AUTOPILOT AND FD – AS DESIRED 5. INITIATE BANK – 15° TO 30°

APPROACH AND RECOVERY

AERODYNAMIC BUFFET, AND/OR ROLL-OFF, WHICHEVER OCCURS FIRST

COMPLETION OF MANEUVER

1. RETURN AIRCRAFT TO DESIRED FLIGHTPATH

1. AUTOPILOT – DISCONNECT 2. PITCH ATTITUDE – REDUCE 3. ROLL ATTITUDE – LEVEL 4. THROTTLES – TO DETENT 5. AIRSPEED – INCREASE 6. VAPP +10, FLAPS - UP

18 MANEUVERS AND PROCEDURES AT STALL INDICATION

Figure 18-3.   Approach to Stall - Takeoff Configuration

Training execution: 1. The instructor sets up the stall scenario. 2. The entry altitude should be consistent with the expected operational environment for the stall configuration. 3. For training and evaluation, the maneuvers may be accomplished with the autopilot on or off as directed by the instructor. 4. The standard is based on the demonstration of smooth, positive control during entry, approach to stall, and recovery. The aim of these stall profiles is to familiarize the pilot with the stall characteristics and to train recognition and recovery procedures in accordance with the ATP Practical Test Standards in flight simulator training only. These stall profiles are not intended for maintenance test flights or aircraft training.

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TRAINING SET UP FOR MANEUVER

AERODYNAMIC BUFFET, AND/OR ROLL-OFF, WHICHEVER OCCURS FIRST

COMPLETION OF MANEUVER

1. RETURN AIRCRAFT TO DESIRED FLIGHTPATH

1. AUTOPILOT – DISCONNECT 2. PITCH ATTITUDE – REDUCE 3. ROLL ATTITUDE – LEVEL 4. THROTTLES – TO DETENT 5. AIRSPEED – INCREASE 6. FLAPS – TAKEOFF & APPROACH 7. POSITIVE RATE, GEAR – UP 8. VAPP +10, FLAPS – UP

18 MANEUVERS AND PROCEDURES

1. MINIMUM ALTITUDE – AS REQUIRED 2. SET VAPP & VREF 3. THRUST – 40-50% N1 4. FLAPS – TAKEOFF & APPROACH 5. GEAR – DOWN 6. FLAPS – LAND 7. AUTOPILOT AND FD – AS DESIRED

APPROACH AND RECOVERY

AT STALL INDICATION

Figure 18-4.   Approach to Stall - Landing Configuration

Training execution: 1. The instructor sets up the stall scenario. 2. The entry altitude should be consistent with the expected operational environment for the stall configuration. 3. For training and evaluation, the maneuvers may be accomplished with the autopilot on or off as directed by the instructor. 4. The standard is based on the demonstration of smooth, positive control during entry, approach to stall, and recovery. The aim of these stall profiles is to familiarize the pilot with the stall characteristics and to train recognition and recovery procedures in accordance with the ATP Practical Test Standards in flight simulator training only. These stall profiles are not intended for maintenance test flights or aircraft training.

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STALL RECOVERY RATIONALE Autopilot..................................................................................................................... Disconnect Rationale While maintaining the attitude of the airplane, disconnect the autopilot. Ensure the pitch attitude does not increase when disconnecting the autopilot. This may be very important in out-of-trim situations. Manual control is essential to recovery in all situations. Leaving the autopilot connected may result in inadvertent changes or adjustments that may not be easily recognized or appropriate, especially during high workload situations. Nose down pitch control................................................ Apply until stall warning is eliminated Nose down pitch trim..................................................................................................As Needed Rationale

18 MANEUVERS AND PROCEDURES

Reducing the angle of attack is crucial for recovery. This will also address autopilotinduced excessive nose up trim. If the control column does not provide sufficient response, pitch trim may be necessary. However, excessive use of pitch trim may aggravate the condition, or may result in loss of control or high structural loads. Bank..........................................................................................................................Wings Level Rationale This orients the lift vector for recovery. Power...........................................................................................................................As Needed Rationale During a stall recovery, maximum power is not always needed. A stall can occur at high power or at idle power. Therefore, the power is to be adjusted accordingly during the recovery. For airplanes with engines mounted above the wings, thrust application creates a helpful pitch-down tendency. For propeller-driven airplanes, power application increases the airflow around the wing, assisting in stall recovery. Return to the desired flightpath. Rationale Apply gentle action for recovery to avoid secondary stalls then return to desired flightpath.

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SIMULATOR TRAINING GUIDANCE

For the CE-510 Series Pilot Initial Training Course, the scenarios will be incorporated into the simulator lesson plans as follows:

SIMULATOR SESSION # 1 1. Stall Prevention Briefing 2. AOA Reduction Recovery Demonstration 3. Enroute (Clean) Configuration Stall A. Manual Flight Conditions B. IMC Conditions C. Low Altitude (Approx 5000’AGL)

RECURRENT TRAINING COURSE For the CE-510 Series Pilot Recurrent Training Course, the scenarios will be incorporated into the simulator lesson plans as follows:

SIMULATOR SESSION # 1 1. AOA Reduction Recovery Demonstration 2. Enroute (Clean) Configuration Stall A. Manual Flight Conditions 3. Takeoff Configuration Stall 4. Landing Configuration Stall

4. Takeoff Configuration Stall

SIMULATOR SESSION # 2

5. Landing Configuration Stall

1. Enroute (Clean) Configuration Stall

SIMULATOR SESSION # 5 1. Enroute (Clean) Configuration Stall

18 MANEUVERS AND PROCEDURES

INITIAL TRAINING COURSE

A. High Altitude B. Automated Flight Conditions

A. High Altitude B. Automated Flight Conditions C. VMC Conditions 2. Stall with System Malfunction A. Stall system related malfunction B. Stall with reduced pilot warning

SIMULATOR SESSION #7 Using different aircraft conditions, weights and CG loading than trained in previous sessions: 1. Enroute (Clean) Configuration Stall 2. Takeoff Configuration Stall 3. Landing Configuration Stall

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V1

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V2 + 12 KT

3

2. POSITIVE RATE OF CLIMB—GEAR UP

Figure 18-5.   Takeoff - Normal

1. AT VR—ROTATE SMOOTHLY TO 10° NOSE UP ATTITUDE

VR

1

18 MANEUVERS AND PROCEDURES 3. AT A PREDETERMINED SAFE ALTITUDE CONSIDERING THE TERRAIN AND OBSTACLES AT A MINIMUM AIRSPEED OF V2 + 12 KT, RETRACT THE FLAPS, ACCELERATE TO NORMAL CLIMB SPEED, AND COMPLETE THE AFTER TAKEOFF-CLIMB CHECKLIST

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2. • BRAKES—MAXIMUM PILOT EFFORT • THROTTLES—IDLE • SPEEDBRAKES—EXTEND • THROTTLES—CUTOFF (IF RUNWAY DEPARTURE IS IMMINENT)

2

18 MANEUVERS AND PROCEDURES

Figure 18-6.   Takeoff - Engine Failure (Speed Below V1)

1. ENGINE FAILURE PRIOR TO V1

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1. ENGINE FAILURE AT OR ABOVE V1, MAINTAIN DIRECTIONAL CONTROL AND ACCELERATE TO VR

1

Figure 18-7.   Takeoff - Engine Failure (At or Above V1)

2. AT VR—ROTATE TO 10° NOSE UP ATTITUDE AND CLIMB AT V2

2

3, 4, 5

3. GEAR UP WHEN POSITIVE RATE OF CLIMB IS ESTABLISHED. MAINTAIN V2 UNTIL 1,500 FEET AGL OR CLEAR OF OBSTACLES, WHICHEVER IS HIGHER; ACCELERATE TO V2 + 10 KT, AND RETRACT THE FLAPS.

18 MANEUVERS AND PROCEDURES

18-16 4. ACCELERATE TO VENR AND CLIMB IF NEEDED

5. COMPLETE THE AFTER TAKEOFF, CLIMB, AND ENGINE FAILURE CHECKLISTS

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1. DOWNWIND LEG (1,500 FT AGL): • AIRSPEED—150 KIAS • FLAPS—TAKEOFF AND APPROACH ABEAM MIDFIELD

2. ABEAM TOUCHDOWN: • GEAR—DOWN*

18 MANEUVERS AND PROCEDURES

4. FINAL APPROACH: • FLAPS—LAND • AIRSPEED—VREF TO VREF + 10 KT • REDUCE TO VREF SPEED WHEN LANDING IS ASSURED (NOTE)

3. BASE LEG: • BEGIN DESCENT • AIRSPEED—VREF +20 (MINIMUM) • BEFORE LANDING CHECKLIST COMPLETED

NOTE: IN GUSTY WIND CONDITIONS, INCREASE VREF BY 1/2 OF THE GUST FACTOR IN EXCESS OF 5 KNOTS * IF BEING RADAR-VECTORED TO A VISUAL APPROACH, LOWER THE GEAR ON BASE LEG OR NO LATER THAN 3 MILES FROM THE THRESHOLD ON A STRAIGHT-IN APPROACH.

Figure 18-8.   VFR Approach - Normal

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1. DOWNWIND LEG (1,500 FT AGL): • AIRSPEED—150 KIAS • FLAPS—TAKEOFF AND APPROACH ABEAM MIDFIELD

2. ABEAM TOUCHDOWN: • GEAR—DOWN*

4. FINAL APPROACH: • FLAPS—TO/APR • AIRSPEED—VAPP

18 MANEUVERS AND PROCEDURES

3. BASE LEG: • BEGIN DESCENT • AIRSPEED—VREF +20 KTS

NOTE: IN GUSTY WIND CONDITIONS, INCREASE VREF BY 1/2 OF THE GUST FACTOR IN EXCESS OF 5 KNOTS * IF BEING RADAR-VECTORED TO A VISUAL APPROACH, LOWER THE GEAR ON BASE LEG OR NO LATER THAN 3 MILES FROM THE THRESHOLD ON A STRAIGHT-IN APPROACH. ** SINGLE ENGINE VAPP MINIMUM AND MAINTAIN FLAPS TO/APR

Figure 18-9.   VFR Approach - Single Engine

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2. ABEAM FAF OR PROCEDURE TURN OUTBOUND: • FLAPS—TO/APR • AIRSPEED—VAPP +20

1. DOWNWIND ON VECTORS OR APPROACHING THE INITIAL APPROACH FIX: • DESCENT CHECKLIST—COMPLETE • AIRSPEED—VREF +30 MINIMUM

18 MANEUVERS AND PROCEDURES

5. MISSED APPROACH: • REFER TO MISSED APPROACH SINGLE ENGINE

3. PROCEDURE TURN INBOUND: • GEAR—DOWN • AIRSPEED—VAPP • BEFORE LANDING CHECKLIST—COMPLETE

4. RUNWAY IN SIGHT

NOTE: IN GUSTY WIND CONDITIONS, INCREASE VREF BY 1/2 OF THE GUST FACTOR IN EXCESS OF 5 KNOTS. FOR CIRCLING APPROACHES, MAINTAIN MANEUVERING SPEED CONSISTENT WITH FLAP POSITION. TURN FINAL, SELECT FLAPS TO LAND, AND REDUCE TO VREF SPEED WHEN LANDING IS ASSURED.

Figure 18-10.   ILS Approach - Normal

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2. ABEAM FAF OR PROCEDURE TURN OUTBOUND: • BEFORE LANDING CHECKLIST—INITIATE • FLAPS—TO/APR • AIRSPEED—VREF +20

1. DOWNWIND ON VECTORS OR APPROACHING THE INITIAL APPROACH FIX: • DESCENT CHECKLIST—COMPLETE • AIRSPEED—VREF +30

5. MISSED APPROACH: • REFER TO MISSED APPROACH NORMAL

18 MANEUVERS AND PROCEDURES

3. FIX INBOUND: • GEAR—DOWN • AIRSPEED—VAPP • FLAPS—LAND • AIRSPEED—VREF • BEFORE LANDING CHECKLIST—COMPLETE

4. RUNWAY IN SIGHT

NOTE: IN GUSTY WIND CONDITIONS, INCREASE VREF BY 1/2 OF THE GUST FACTOR IN EXCESS OF 5 KNOTS. FOR CIRCLING APPROACHES, MAINTAIN MANEUVERING SPEED CONSISTENT WITH FLAP POSITION. TURN FINAL, SELECT FLAPS TO LAND, AND REDUCE TO VREF SPEED WHEN LANDING IS ASSURED.

Figure 18-11.   Nonprecision Approach - Normal

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2. ABEAM FAF OR PROCEDURE TURN OUTBOUND: • FLAPS—TO/APR • AIRSPEED—VAPP +20

1. DOWNWIND ON VECTORS OR APPROACHING THE INITIAL APPROACH FIX: • DESCENT CHECKLIST—COMPLETE • AIRSPEED—VREF +30 MINIMUM

18 MANEUVERS AND PROCEDURES

5. MISSED APPROACH: • REFER TO MISSED APPROACH SINGLE ENGINE

3. PROCEDURE TURN INBOUND: • GEAR—DOWN • AIRSPEED—VAPP • BEFORE LANDING CHECKLIST—COMPLETE

4. RUNWAY IN SIGHT

NOTE: IN GUSTY WIND CONDITIONS, INCREASE VREF BY 1/2 OF THE GUST FACTOR IN EXCESS OF 5 KNOTS. FOR CIRCLING APPROACHES, MAINTAIN MANEUVERING SPEED CONSISTENT WITH FLAP POSITION. TURN FINAL, SELECT FLAPS TO LAND, AND REDUCE TO VREF SPEED WHEN LANDING IS ASSURED.

Figure 18-12.   Nonprecision Approach - Single Engine

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Figure 18-13.   Missed Approach - 2 Engine (Precision/Nonprecision)

1. FINAL APPROACH: • GEAR—DOWN • FLAPS—LAND • AIRSPEED—VREF

2. DECISION POINT: “GO-AROUND”; SIMULTANEOUSLY APPLY TAKEOFF POWER, ROTATE 8° NOSE UP ATTITUDE, (GA MODE ON FLIGHT DIRECTOR FOR REFERENCE) AND CHECK/SET FLAPS TO TO/APR.

18 MANEUVERS AND PROCEDURES

3. RAISE THE GEAR WHEN A POSITIVE RATE OF CLIMB IS ESTABLISHED. AT A COMFORTABLE ALTITUDE AND A MINIMUM AIRSPEED OF VREF + 10 KT, RETRACT THE FLAPS, ACCELERATE TO NORMAL CLIMB SPEED, AND COMPLETE THE AFTER TAKEOFF-CLIMB CHECKLIST.

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Figure 18-14.   Missed Approach - Single Engine (Precision/Nonprecision)

1. FINAL APPROACH: • GEAR—DOWN • FLAPS—TO/APR • AIRSPEED—VAPP

2. DECISION POINT: “GO-AROUND”; SIMULTANEOUSLY APPLY TAKEOFF POWER, ROTATE 8° NOSE UP ATTITUDE, (GA MODE ON FLIGHT DIRECTOR FOR REFERENCE) AND CHECK / SET FLAPS TO TAKEOFF AND APPROACH.

3. GEAR UP WHEN POSITIVE RATE OF CLIMB IS ESTABLISHED. MAINTAIN A MINIMUM CLIMB SPEED OF VAPP UNTIL 400' AGL OR CLEAR OF OBSTACLES, WHICHEVER IS HIGHER; ACCELERATE TO VAPP +10, RETRACT FLAPS, AND ACCELERATE TO VENR.

4. SET MAXIMUM CONTINUOUS CLIMB POWER, AND COMPLETE THE SINGLE-ENGINE GO-AROUND CHECKLIST AND THE AFTER TAKEOFF-CLIMB CHECKLIST.

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THRESHOLD: • AIRSPEED—VREF • YAW DAMP—OFF • SPEEDBRAKES—RETRACTED BY 50 FEET

DOWNWIND (1,500FT AGL): • AIRSPEED—VREF +30 • FLAPS—TO/APR • GEAR—DOWN (ABEAM THRESHOLD)

FINAL: • FLAPS—LAND • AIRPSEED—VREF TO VREF +10

18 MANEUVERS AND PROCEDURES BASE: • AIRSPEED—NO SLOWER THAN MINIMUM MANEUVERING* STRAIGHT-IN (4–5 MILES OUT): • GEAR DOWN • BEFORE LANDING CHECK COMPLETE • AIRSPEED VREF

NOTE: * MINIMUM MANEUVERING SPEEDS ARE BASED ON FLAP LAND VREF FOR THE AIRCRAFT CURRENT WEIGHT PLUS ADDITIVES AS SHOWN BELOW. THE OPEN GREEN CIRCLE ON THE AIRSPEED TAPE REPRESENTS 1.3 VS1 FOR THE CURRENT FLAP SETTING. • IF FLAPS LAND—VREF +10 • IF FLAPS TO/APR—VREF +20 • IF FLAPS UP—VREF +30

Figure 18-15.   Visual Approach

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18 MANEUVERS AND PROCEDURES

PROCEDURE • AIRSPEED—200 KTS • THROTTLES—APPROXIMATELY 72% N1 • MAINTAIN ALTITUDE—TRIM AS REQUIRED • MAINTAIN AIRSPEED • THROTTLES—INCREASE SLIGHTLY (2%) AS AIRCRAFT ROLLS THROUGH 30° OF BANK • INITIATE ROLLOUT APPROXIMATELY 20° PRIOR TO ENTRY HEADING

Figure 18-16.   Steep Turns

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CHAPTER 19 WEIGHT AND BALANCE CONTENTS INTRODUCTION................................................................................................................ 19-1 GENERAL............................................................................................................................ 19-2 Weight............................................................................................................................ 19-2 Balance.......................................................................................................................... 19-2 Basic Formula................................................................................................................ 19-2 Weight Shift Formula.................................................................................................... 19-2 Weight Addition or Removal......................................................................................... 19-2 Definitions..................................................................................................................... 19-2 Maximum Weight Limits............................................................................................... 19-3 Maximum Design Center-of-Gravity Limits ................................................................ 19-3 Balance Limits for Normal Ground Operations............................................................ 19-4 FORMS................................................................................................................................. 19-4 Weight-and-Balance Record.......................................................................................... 19-4 Weight-and-Balance Computation................................................................................ 19-5 Standard Seating Configuration.................................................................................... 19-5 Baggage/Cabinet Weight-and-Moment Table................................................................ 19-5 Fuel Loading Weight-and-Moment Table...................................................................... 19-5 Center-of-Gravity Limits Envelope Graph.................................................................... 19-5 EXAMPLES...................................................................................................................... 19-24 General....................................................................................................................... 19-24 Sample Loading Problem (Using Aircraft Model 510-0193 and ON)....................... 19-24 LIMITATIONS................................................................................................................... 19-25 EMERGENCY/ABNORMAL........................................................................................... 19-25

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Aircraft Weighing Form................................................................................................ 19-4

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ILLUSTRATIONS Figure Title Page 19-1.

Baggage Door Opening Dimensions..................................................................... 19-4

19-2.

Baggage Compartment Limits............................................................................... 19-5

19-3.

Aircraft Weighing Form - U.S. Units..................................................................... 19-6

19-4.

Weight-and-Balance Record - U.S. Units.............................................................. 19-7

19-5.

Weight-and-Balance Computation Form 2245 - U.S. Units.................................. 19-8

19-6.

Weight-and-Balance Computation Form 2286 - U.S. Units.................................. 19-9

19-7.

Standard Seating Configuration - U.S. Units...................................................... 19-10

19-8.

Baggage and Cabinet Compartments Standard Weight-and-Moment Tables - U.S. Units (form 2239)......................... 19-11

19-9.

Baggage and Cabinet Compartments Standard Weight-and-Moment Tables - U.S. Units (Form 2288).......................... 19-12

19-10. Fuel Loading Table - U.S. Units.......................................................................... 19-13 19-11. Center-of-Gravity Limits Envelope Graph Form 2235 - U.S. Units................... 19-14

19-13. Weight-and-Balance Record Form 2248 - Metric Units..................................... 19-16 19-14. Weight-and-Balance Computation Form 2246 - Metric Units........................... 19-17 19-15. Weight-and-Balance Computation Form 2287 - Metric Units........................... 19-18 19-16. Standard Seating Configuration - Metric Units.................................................. 19-19 19-17. Baggage and Cabinet Compartments Standard Weight-and-Moment Tables - Metric Units (form 2240)....................... 19-20 19-18. Baggage and Cabinet Compartments Standard Weight-and-Moment Tables - Metric Units (Form 2289)...................... 19-21 19-19. Fuel Loading Table - Metric Units...................................................................... 19-22 19-20. Center-of-Gravity Limits Envelope Graph Form 2236 - Metric Units............... 19-23

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19-12. Aircraft Weighing Form 2234 - Metric Units..................................................... 19-15

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19-21. Example Aircraft Weighing Form - U.S. Units................................................... 19-26 19-22. Example Weight-and-Balance Computation Form 2245 - U.S. Units................ 19-27 19-23. Example Weight-and-Balance Computation Form 2286 - U.S. Units................ 19-28 19-24. Example Center-of-Gravity Limits Envelope Graph - U.S. Units...................... 19-29 19-25. Example Aircraft Weighting Form - Metric Units.............................................. 19-30 19-26. Example Weight-and-Balance Computation Form 2246 - Metric Units............ 19-31 19-27. Example Weight-and-Balance Computation Form 2287 - Metric Units............ 19-32 19-28. Example Center-of-Gravity Limits Envelope Graph - Metric Units.................. 19-33

19 WEIGHT AND BALANCE

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INTRODUCTION This chapter provides procedures for establishing the basic empty weight and moment of the Mustang aircraft. It also provides procedures for determining the weight and balance for flight. This section also describes items on the Weight and Balance Data Sheet, which is provided with the aircraft as delivered from Cessna Aircraft Company.

WARNING It is the responsibility of the pilot to make sure the aircraft is loaded properly. The aircraft must be loaded so as to remain within the weight and balance limits prescribed in the Aircraft Flight Manual (AFM) throughout the flight from takeoff to landing.

CAUTION This manual presents data in both U.S. and metric units. Make sure that you use the appropriate data in the weight-and-balance computations for your airplane.

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CHAPTER 19 WEIGHT AND BALANCE

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GENERAL

This formula can be utilized to shift weight if the CG is found to be out of limits. Use of this formula avoids working the entire problem over again.

WEIGHT Aircraft maximum weights are predicated on structural strength and performance requirements. It is necessary to ensure that the aircraft is loaded within the various weight restrictions to maintain structural integrity and to ensure that performance is acceptable.

BALANCE Balance, or the location of the center of gravity (CG), deals with aircraft stability. The horizontal stabilizer must be capable of providing an equalizing moment to that which is produced by the wing and the aircraft overall. Since the amount of force produced by the horizontal stabilizer through elevator movement is limited, the range of movement of the CG is restricted so proper aircraft stability and control is maintained. Stability increases as the CG moves forward. However, if the CG is located out of limits too far forward, the aircraft may become so nose heavy that it cannot be rotated at the proper speed or flared for landing.

19 WEIGHT AND BALANCE

Locating the CG aft of limits is considerably worse because the stability decreases. Eventually the airplane becomes unstable.

BASIC FORMULA Weight x Arm = Moment This is the basic formula upon which all weight and balance calculations are based. Remember that the arm or CG location can be found by adapting the formula as follows: Arm = Moment Weight

WEIGHT SHIFT FORMULA Weight Shifted Distance CG is shifted (x) = Total Weight Distance weight is shifted

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WEIGHT ADDITION OR REMOVAL If weight is to be added or removed after a weight and balance has been computed, a simple formula can be used to determine the shift in the center of gravity. Weight added (or removed) Distance CG is shifted New total weight = Distance between the weight arm and the old CG arm If it is desired to find the weight change needed to accomplish a particular CG change, the formula can be adapted as follows: Weight to be added (or removed) Distance CG is shifted Old total weight = Distance between the weight arm and the new CG arm

DEFINITIONS General Basic Empty Weight—Standard empty weight plus installed optional equipment. This is the weight reflected on the Weight and Balance Data Form supplied with the airplane. MAC—Mean Aerodynamic Chord is an engineering term that represents an airfoil’s chord in aircraft design. As such, it is a constant length, which is also used in the calculation of center-of-gravity location in terms of percent MAC. Operational Landing Weight—This is the weight at the start of touchdown. It is subject to airport, operational, and related restrictions. It must not exceed maximum landing weight. Operational Takeoff Weight—This is the weight at the start at the takeoff run. It is subject to airport, operational, and related restrictions. It must not exceed maximum takeoff weight.

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Ramp Weight—Operational takeoff weight plus fuel necessary for start, taxi, and run-up. It must not exceed maximum ramp weight. Standard Empty Weight—Weight of standard airplane including standard items. Standard Items—Equipment and fluids not an integral part of a particular airplane and not a variation for the same type of airplane. These items may include, but are not limited to the following: • Unusable fuel • Trapped fuel • Engine oil • All hydraulic fluid • Serviced fire extinguisher Trapped Fuel—Fuel remaining when airplane is defueled by normal means using the procedures and attitudes specified for draining the tanks. Unusable Fuel—Fuel remaining after a fuel runout test has been completed in accordance with governmental regulations. Usable Fuel—Fuel available for flight planning. Useful Load—Difference between takeoff weight, or ramp weight if applicable, and basic empty weight. It includes payload, usable fuel, and other usable fluids not included as standard items. Zero Fuel Weight—Basic empty weight plus payload. It must not exceed maximum zero fuel weight.

Weight Limitations Maximum Landing Weight—Maximum weight approved for the landing touchdown. Maximum Ramp Weight—Maximum weight for ground maneuvers as limited by airplane strength and airworthiness requirements. It includes weight of start, taxi, and run-up fuel.

Maximum Takeoff Weight—Maximum weight for takeoff as limited by airplane strength and airworthiness requirements. This is the maximum weight approved for the start of the takeoff run. Maximum Zero Fuel Weight—Maximum weight allowed exclusive of usable fuel.

MAXIMUM WEIGHT LIMITS Maximum ramp weight .................. 8,730 pounds (3,960 kilograms) Maximum takeoff weight*.............. 8,645 pounds (3,921 kilograms) Maximum landing weight*............ 8,000 pounds (3,629 kilograms) Maximum zero fuel weight ............ 6,750 pounds (3,062 kilograms) Maximum nose baggage weight ....... 320 pounds (145 kilograms) Maximum tailcone baggage weight .. 300 pounds (136 kilograms) * Refer to the FAA-approved Aircraft Flight Manual for additional restrictions that may apply due to runway, pressure, altitude, and temperature.

MAXIMUM DESIGN CENTER-OF-GRAVITY LIMITS (U.S. Units) Forward limit 285.59 inches aft of the datum • (19.00% MAC) at 5,550 pounds to 6,927 pounds and/or 287.04 inches aft of the datum (21.32% MAC) at 8,730 pounds with straight line variation between these points Aft limit 292.46 inches aft of the datum • (30.00% MAC) at 5,314 pounds to 8,730 pounds

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Payload—Zero fuel weight minus basic empty weight. This is the weight of crew, passengers, baggage, and cargo.

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(Metric Units) Forward limit 7,254 millimeters aft of the datum • (19.00% MAC) at 2,517 kilograms to 3,142 kilograms and/or 7,291 millimeters aft of the datum (21.32% MAC) at 3,960 kilograms with straight line variation between these points Aft limit 7,428 millimeters aft of the datum • (30.00% MAC) at 2,410 kilograms to 3,960 kilograms

BALANCE LIMITS FOR NORMAL GROUND OPERATIONS Removal of any large components forward of the wing may require temporary ballast in the nose. There are no balance limits for fuel loading operations. For example, one wing fuel tank can be completely filled prior to adding fuel to the opposite wing fuel tank.

19 WEIGHT AND BALANCE

As with any aircraft, use caution when loading. Always load aircraft from the nose to the aft. Low fuel, unlevel terrain, wind, and snow aft of the main landing gear can magnify the potential for tipping the aircraft on its tail.

Zero Reference Datum Line The zero reference datum line of the Model 510 is located 143.70 inches (3,650 millimeters) in front of the jig point (nose jack pad location). Horizontal distance from zero datum line to the leading edge

of the MAC is 273.71 inches (6952 millimeters). The addition of any weight in any location therefore, results in a positive moment change.

FORMS The Weight-and-Balance forms are discussed below, and examples of the forms are included in Figures 19-3 through 19-20. If the aircraft has a different seating configuration from the one depicted in the example, the form appropriate to that configuration is in the AFM.

AIRCRAFT WEIGHING FORM The aircraft weight, CG arm, and moment (divided by 100) are all listed at the bottom of this form as the aircraft is delivered from the factory (Figures 19-3 and 19-12). Ensure that the basic empty weight figures listed are current and have not been amended.

WEIGHT-ANDBALANCE RECORD The Weight-and-Balance Record amends the Aircraft Weighing Form (Figures 19-4 and 19-13). After delivery, if a service bulletin is applied to the aircraft or if equipment is removed or added that would affect the CG or basic empty weight, it must be recorded on this form in the AFM. The crew must always have access to the current aircraft basic weight and moment in order to be able to perform weight and balance computations. The tables already have computed moments/ 100 for weights in various seating locations in the aircraft (Figures 19-5, 19-6, 19-14, and 19-15).

Figure 19-1.   Baggage Door Opening Dimensions

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WEIGHT-AND-BALANCE COMPUTATION

flooring in that area can support. This same point applies to the aft extension and cabinet contents, as well.

A step-by-step process is outlined for determining weight and CG limits. The payload computations are made in the left column, while the rest of the computations are shown in the right column.

FUEL LOADING WEIGHT-AND-MOMENT TABLE

STANDARD SEATING CONFIGURATION

All of the tables have arms listed for the various locations except the fuel table (Figures 19-10 and 19-19). Notice that the arm varies depending on the quantity of usable fuel.

The tables already have computed moments/100 for weights in various seating locations in the aircraft (Figures 19-7 and 19-16).

BAGGAGE/CABINET WEIGHT-AND-MOMENT TABLE Notice in the baggage compartment contents table the last weight that a moment/100 is listed for under the last weight for which a moment/100 listed under the column corresponds to the placard limit in that compartment (Figures 19-8, 19-9, 19-17, and 19-18). Remember that this limit is structural in nature. It is based on the maximum weight the

CENTER-OF-GRAVITY LIMITS ENVELOPE GRAPH After summing all the weights and moments, it is necessary to determine whether the CG is within allowable limits. This graph represents the allowable CG envelope (Figures 19-11 and 19-20). The way to plot the location of the CG on the graph is to determine the CG location in inches aft of datum, then plot it against the weight. To determine the CG arm, the total moment (moment x 100) is divided by the total aircraft weight.

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Figure 19-2.   Baggage Compartment Limits

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Figure 19-3.   Aircraft Weighing Form - U.S. Units

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Figure 19-4.   Weight-and-Balance Record - U.S. Units

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Figure 19-5.   Weight-and-Balance Computation Form 2245 - U.S. Units

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Figure 19-6.   Weight-and-Balance Computation Form 2286 - U.S. Units

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Figure 19-7.   Standard Seating Configuration - U.S. Units

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Figure 19-8.   Baggage and Cabinet Compartments Standard Weight-and-Moment Tables - U.S. Units (form 2239) FOR TRAINING PURPOSES ONLY

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Figure 19-9.   Baggage and Cabinet Compartments Standard Weight-and-Moment Tables - U.S. Units (Form 2288)

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Figure 19-10.   Fuel Loading Table - U.S. Units

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Figure 19-11.   Center-of-Gravity Limits Envelope Graph Form 2235 - U.S. Units

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Figure 19-12.   Aircraft Weighing Form 2234 - Metric Units

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Figure 19-13.   Weight-and-Balance Record Form 2248 - Metric Units

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Figure 19-14.   Weight-and-Balance Computation Form 2246 - Metric Units

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Figure 19-15.   Weight-and-Balance Computation Form 2287 - Metric Units

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Figure 19-16.   Standard Seating Configuration - Metric Units

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Figure 19-17.   Baggage and Cabinet Compartments Standard Weight-and-Moment Tables - Metric Units (form 2240)

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Figure 19-18.   Baggage and Cabinet Compartments Standard Weight-and-Moment Tables - Metric Units (Form 2289) FOR TRAINING PURPOSES ONLY

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Figure 19-19.   Fuel Loading Table - Metric Units

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Figure 19-20.   Center-of-Gravity Limits Envelope Graph Form 2236 - Metric Units

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EXAMPLES GENERAL The contribution that any loading item makes to a shift in the aircraft center of gravity depends upon its distance from the aircraft Basic Empty Weight center of gravity. Any weight placed in the aft baggage compartment will shift the center of gravity aft since it is aft of the typical Basic Empty Weight center of gravity. Adding fuel or passengers will shift the center of gravity forward since they are forward of the typical Basic Empty Weight center of gravity. The magnitude of the shift for any given weight is proportional to the length of the moment arm from the center of gravity.

SAMPLE LOADING PROBLEM (USING AIRCRAFT MODEL 510-0193 AND ON) The following step-by-step procedure illustrates a logical manner in which to determine the takeoff weight and center of gravity. Loading tables can be found in this manual and in the Weight and Balance Data Sheets. 19 WEIGHT AND BALANCE

The Example Center-of-Gravity Limits Envelope Graph (Figures 19-24 and 19-28) is an example plot of the fuel burn calculated from the Example Weight and Balance Computation Form (Figures 19-23 and 19-27).

NOTE For the purposes of this sample problem, weights are rounded to the nearest whole pound and moment index to two (2) decimal places for entry on the Example Weight and Balance Computation Form.

1. Enter the Basic Empty Weight, moment index (moment/100 for U.S. Units or moment/1,000 for Metric Units) and center of gravity as weighed from the Airplane Weighing Form (Figures 19-21 and 19-25) on line 1 BASIC EMPTY WEIGHT of the Example Weight and Balance Computation Form (Figures 19-23 and 19-27). 2. Determine the moment index for each passenger using the Crew and Passenger Weight and Moment Table (see Figures 19-7 and 19-16). Enter the weight and moment index for each passenger on the PAYLOAD COMPUTATIONS side of the Weight and Balance Computation Form (Figures 19-23 and 19-27). 3. Determine the moment index for any cabinet contents using the Baggage and Cabinet Compartments Standard Weight and Moment Table (see Figures 19-9 and 19-18). Enter the weight and moment index on the PAYLOAD COMPUTATIONS side of the Weight and Balance Computation Form. 4. Determine the moment index for baggage loading in the nose, toilet seat, and tail cone compartments using the Baggage and Cabinet Compartments Standard Weight and Moment Table (see Figures 19-9 and 19-18). Enter the weight and moment index on the PAYLOAD COMPUTATIONS side of the Weight and Balance Computation Form. 5. Total the weights and moment indices of the payload items at the bottom of the PAYLOAD COMPUTATIONS side of the Weight and Balance Computation Form and enter these values on line 2 PAYLOAD of the Weight and Balance Computation Form. 6. Enter the sums of the values on lines 1 and 2 onto line 3 ZERO FUEL WEIGHT of the Weight and Balance Computation Form. 7. Divide the zero fuel moment index by the zero fuel weight and multiply by 100 (1,000) to determine the zero fuel weight center of gravity. 8. Check the zero fuel weight and center of gravity by locating the weight and center of gravity on the Example Center-of Gravity Limits Envelope Graph (Figures 19-24 and 19-28).

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NOTE

NOTE

Approved points are located below the maximum zero fuel weight line.

Approved points are located inside the envelope below the maximum takeoff weight line.

10. Enter the sums of the values on lines 3 and 4 onto line 5 RAMP WEIGHT of the Weight and Balance Computation Form. 11. Divide the ramp moment index by the ramp weight and multiply by 100 (1,000) to determine the ramp weight center of gravity. 12. Check the ramp weight by locating the weight and center of gravity on the Example Center of Gravity Limits Envelope Graph (Figures 19-24 and 19-28).

NOTE Approved points are located below the maximum ramp weight line. 13. Enter the taxi fuel weight and moment index on line 6 LESS FUEL FOR TAXIING of the Weight and Balance Computation Form. A standard 85-pound (38.56-kilogram) burn-off is assumed. The moment index for taxi is the difference between the moment index of the fuel loaded and the moment index of the fuel remaining on-board after taxi. 14. Enter the differences between the values on lines 5 and 6 onto line 7 TAKEOFF WEIGHT of the Weight and Balance Computation Form. 15. Divide the takeoff moment index by the Takeoff Weight and multiply by 100 (1,000) to determine the takeoff center of gravity. Enter the takeoff center of gravity on line 7 of the Weight and Balance Computation Form. 16. Check the takeoff weight by locating the weight and center of gravity on the Example Centerof-Gravity Limits Envelope Graph (Figures 19-24 and 19-28).

17. Determine the estimated weight of the fuel to be used in flight. The moment index of the fuel used in flight is the difference between the moment index of the fuel remaining after taxi and the moment index of the fuel remaining after reaching the destination. 18. Enter the weight and moment index of the fuel used in flight on line 8 LESS FUEL TO DESTINATION of the Weight and Balance Computation Form. 19. Enter the differences between the values on lines 7 and 8 onto line 9 LANDING WEIGHT of the Weight and Balance Computation Form. 20. Divide the landing moment index by the landing weight and multiply by 100 (1,000) to determine the landing center-of-gravity. Enter the landing center-of-gravity on line 9 of the Weight and Balance Computation Form. 21. Check the landing weight by locating the weight and center of gravity on the Example Center of Gravity Limits Envelope Graph (Figures 19-24 and 19-28).

NOTE Approved points are located inside the envelope below the Maximum Landing Weight line.

LIMITATIONS For specific limitations, refer to the FAA-approved AFM.

EMERGENCY/ ABNORMAL For specific information on emergency/abnormal procedures, refer to the appropriate checklist or FAA-approved AFM.

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19 WEIGHT AND BALANCE

9. Look up the weight of fuel to be used for the flight and the corresponding moment index in the Fuel Loading Table (see Figures 19-10 and 19-19). Enter the weight and moment index on line 4 FUEL LOADING of the Weight and Balance Computation Form.

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Figure 19-21.   Example Aircraft Weighing Form - U.S. Units

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Figure 19-22.   Example Weight-and-Balance Computation Form 2245 - U.S. Units

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Figure 19-23.   Example Weight-and-Balance Computation Form 2286 - U.S. Units

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Figure 19-24.   Example Center-of-Gravity Limits Envelope Graph - U.S. Units

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Figure 19-25.   Example Aircraft Weighting Form - Metric Units

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Figure 19-26.   Example Weight-and-Balance Computation Form 2246 - Metric Units

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Figure 19-27.   Example Weight-and-Balance Computation Form 2287 - Metric Units

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Figure 19-28.   Example Center-of-Gravity Limits Envelope Graph - Metric Units

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CHAPTER 20 FLIGHT PLANNING AND PERFORMANCE CONTENTS INTRODUCTION................................................................................................................ 20-1 FORMULAS......................................................................................................................... 20-2

ILLUSTRATIONS Figure Title Page 20-1.   Calculation of Takeoff Performance......................................................................... 20-3

20 FLIGHT PLANNING AND PERFORMANCE

20-2.   Calculation of Landing Performance........................................................................ 20-4

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CHAPTER 20 FLIGHT PLANNING AND PERFORMANCE

INTRODUCTION

20 FLIGHT PLANNING AND PERFORMANCE

Performance is calculated using a combination of charts and tables in the Aircraft Flight Manual and the Aircraft Performance Manual. The takeoff and landing performance data is found in Section IV—“Performance” and Section VII “Advisory” of the AFM. The climb, cruise, and descent performance data is found in the Performance Manual.

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GENERAL

Climb rate (feet per minute) =

This aircraft is certified to Part 25 standards. Keep in mind that the FAR Part 25 performance requirements do not meet the minimum requirements (3.3% or 200 ft/nm) of the FAA “IFR Takeoff Flight Path”.

Groundspeed x Gradient A simplified block diagram of the calculation of takeoff performance is illustrated in Figure 20-1. A simplified block diagram of the calculation of landing performance is illustrated in Figure 20-2.

The maximum takeoff weight–pounds permitted by climb requirements chart only guarantees second segment climb performance, not any of the other segments. The following are the minimum climb gradients as specified by FAR Part 25: • 1st segment ................................ 0% gross • 2nd segment ............................... 1.6% net • 3rd segment ....................................... N/A • Final segment ......................... 1.2% gross

NOTE The gross climb gradient reduced by a required factor and used for calculation of take-off flight path.

FORMULAS Runway Slope = Change in Elevation Between Ends of the Runway (Rise) Runway Length

X 100

Gradient (in %) = Feet per NM 20 FLIGHT PLANNING AND PERFORMANCE

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CALCULATE TAKEOFF PERFORMANCE · Determine gross weight of aircraft for type of loading desired · Obtain airport information (i.e. active runway, available runway length, temperature, pressure altitude, wind, runway conditions and runway gradient (if applicable) and obstacles in the takeoff ) · Determine that the temperature is within the ambient temperature limits · Determine crosswind/parallel wind component for active runway

Recalculate performance at a lower aircraft weight

YES

Does calculated T/O weight exceed the max T/O permitted by climb requirements? Using the calculated T/O gross weight, determine TOFL and VSPEEDS for dry conditions Correct for Runway Gradient

YES

Contaminated runway?

AFM Section VII: Calculate the corrected TOFL for Adverse Runway Conditions

Recalculate performance at a lower aircraft weight

NO

YES

Available runway LESS than TOFL? NO Determine level-off altitude

YES

Minimum climb requirements? NO

Climb requirements met?

NO

Recalculate performance at a lower aircraft weight

20 FLIGHT PLANNING AND PERFORMANCE

AFM Section IV: Calculate SECOND SEGMENT TAKEOFF NET CLIMB GRADIENT – PERCENT

YES Complete

Figure 20-1.   Calculation of Takeoff Performance

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CALCULATE LANDING PERFORMANCE · Determine gross weight of aircraft at the time of arrival at the destination airport. · Obtain airport information; i.e., active runway, available runway length, temperature, pressure altitude, wind, runway conditions and runway gradient if applicable. Determine that the temperature is within the ambient temperature limits. · Determine crosswind/parallel wind component for active runway. · Check the maximum landing weight permitted by approach requirements and the brake energy limits.

YES

Must burn off fuel prior to landing

YES

AFM Section VII: Calculate the corrected

Landing Weight Restricted? NO

Contaminated runway? NO

adverse runway conditions

YES

Must reduce the airplane landing weight

YES

Divide the landing distance by 0.6

Avail. Runway less than required? NO

FAR 135 Operations? NO

20 FLIGHT PLANNING AND PERFORMANCE

Determine the takeoff/go-around thrust setting using the approach climb and landing climb gradient tables in the event that a go-around is necessary

Complete

Figure 20-2.   Calculation of Landing Performance

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CHAPTER 21 CREW RESOURCE MANAGEMENT CONTENTS Page WHAT IS CREW RESOURCE MANAGEMENT?............................................................. 21-1 SITUATIONAL AWARENESS............................................................................................ 21-2 COMMAND AND LEADERSHIP...................................................................................... 21-3 COMMUNICATION PROCESS.......................................................................................... 21-4 Communication Techniques: Inquiry, Advocacy, and Assertion................................... 21-5 DECISION-MAKING PROCESS........................................................................................ 21-6

ILLUSTRATIONS 21-1.   Situational Awareness in the Cockpit....................................................................... 21-2 21-2.   Command and Leadership........................................................................................ 21-3 21-3.   Communication Process........................................................................................... 21-4 21-4.   Decision Making Process......................................................................................... 21-6

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CHAPTER 21 CREW RESOURCE MANAGEMENT

WHAT IS CREW RESOURCE MANAGEMENT? According to the Federal Aviation Administration, Crew Resource Management (CRM) is described as “the effective use of all resources to achieve safe and efficient flight operations.” In practice, CRM is a set of competencies designed to enhance safety and reduce human error. Resources can include, but are not limited to, additional crewmembers, maintenance technicians, flight attendants, air traffic controllers, dispatchers and schedulers, and line service personnel. CRM was not designed to usurp the authority of the pilot in command; rather, it was developed as a means to assist with situational awareness and decision making to increase safety margins and achieve accident- and incident-free flight ­operations. Most experts agree that a highly coordinated crew using a standardized set of procedures is more likely to avoid and identify errors. Effective communication and the use of briefing and debriefing are tools that can be used to build the “team concept” and maintain situational awareness. Utiliz-

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ing a standard set of callouts provides a means to incorporate CRM. Standardization keeps all crewmembers “in the loop” and provides an opportunity to detect an error early on, before it has an opportunity to build into an accident chain. Proficiency in CRM requires all crewmembers to have a working knowledge of how to maintain situational awareness, techniques for o­ ptimum decision making, desirable leadership and followership characteristics, cross-checking and monitoring techniques, means of fatigue and stress management, and ­communication. CRM training is an important part of your FlightSafety training experience. Throughout your training event, your instructor will p ­ rovide general CRM guidance as well as ­identify CRM issues, philosophies, and techniques that are specific to the aircraft you fly. To a­ ssist with this, the FlightSafety CRM model has been incorporated into this training guide. The model can be used as a guide or a

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refresher on how to incorporate CRM principles into your day-to-day line operations. This model is not intended to replace a formalized course of CRM instruction, and attendance at a CRM-specific course is highly recommended.

SITUATIONAL AWARENESS Situational awareness is a fundamental CRM concept. Often described as “knowing what’s going on around you,” the loss of situational awareness is often identified as a causal factor in an incident or accident. Collective s­ ituational awareness is a measurement of the total situational awareness among all m ­ embers involved in the operation.

To maintain a high level of collective situational ­awareness open, timely, and accurate communication is ­required. In the situational awareness model two-way arrows represent the two-way ­communication that must occur between the pilot flying and the pilot monitoring. Each pilot contributes to collective situational awareness. Circumstances will sometimes present clues that situational awareness is becoming ­impaired. These “behavioral markers” are listed under clues to identifying loss of situational awareness. As the number of these clues increases, the chance of losing situational awareness increases as well. Maintaining situational awareness requires a constant state of vigilance. Complacency has often been the precursor to a loss of situational awareness (Figure 21-1).

Figure 21-1.   Situational Awareness in the Cockpit

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COMMAND AND LEADERSHIP Command and leadership are not synonymous. The status “pilot in command” is designated by an organization. Command responsibility can’t be shared with other crewmembers. Leadership, on the other hand, is a role that can be shared. Effective leadership should focus on “what’s right,” not on “who’s right.”

There is no “ideal” or “best” leadership style. An immediate crisis might require fairly strict leadership, to ensure stability and to reassure other crewmembers, while other situations might be handled more effectively by encouraging crew participation in the ­decision-making process.

Leadership styles range from “autocratic” to “laissez-faire.” An autocratic leadership style exercises a high degree of control and allows a low degree of participation from team ­members in reaching decisions. A laissez-faire leadership style exercises a low degree of control and allows a high degree of participation from team members. Effective leaders tend to be less extreme, relying on either authoritarian or democratic leadership styles (Figure 21-2).

Figure 21-2.   Command and Leadership

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COMMUNICATION PROCESS

• An event occurs, creating a need to communicate. The event may be a change in the status of some operational goal, such as rate of descent.

Communication is the most important tool for maintaining situational awareness. Effective communication requires the ability to provide appropriate information, at the appropriate time, to the appropriate person (Figure 21-3). Communication may be verbal (aural) or written. Written communications in the cockpit include symbolic messages and indications that are electronically transmitted and displayed.

• A sender observes the event.

As illustrated on the CRM Blue Card, some e­ lements are common to most cockpit ­communications:

• The sender transmits a message to a receiver, conveying occurrence of the event. • The receiver transmits feedback to the sender, acknowledging the message. • The receiver’s feedback may include an additional message, confirming the intended corrective action, or instructing the sender to continue monitoring the ­operational goal.

Figure 21-3.   Communication Process

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Revision 1.1

Barriers to communication limit our ability to maintain situational awareness. As illustrated on the Blue Card, internal (or personal) communication barriers can diminish our perception of the need to communicate. An observer who is distracted, for example, may fail to detect a change in the status of an operational goal. Internal barriers can also inhibit a sender’s willingness to communicate, or affect a receiver’s acceptance and interpretation of a transmitted message. External communication barriers, such as overcrowded radio frequencies, can interfere with the sender’s ability to transmit a message, or with the receiver’s ability to transmit feedback. Differences in language or dialect can also become external barriers to communication. CRM provides three techniques for overcoming communication barriers: • Inquiry—A technique for increasing your own situational awareness

COMMUNICATION TECHNIQUES: INQUIRY, ADVOCACY, AND ASSERTION Inquiry, advocacy, and assertion can be effectively used in the aviation environment to help solve communication problems. Each item is a step in the process. The steps provide a metaphor that emphasizes the principle of escalation. In other words, a person must first practice inquiry, then advocacy, then assertion. A person practicing assertiveness is not trying to be insubordinate or disrespectful; rather, assertion is an expression of the fact that a level of discomfort exists with a particular situation. Assertion is an attempt to seek resolution. The goal of inquiry is to increase individual situational awareness, the goal of advocacy is to increase collective situational awareness, and the goal of assertion is to reach a ­conclusion.

• Advocacy—A technique for increasing someone else’s awareness • Assertion—A technique for getting your point across When conflict on the flight deck interferes with communication, it usually originates from one pilot’s tendency to make “solo” decisions. Avoid this kind of conflict by focusing your questions and comments on WHAT is right, rather than on WHO is right.

Revision 1.1

FOR TRAINING PURPOSES ONLY

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21 CREW RESOURCE MANAGEMENT

CITATION MUSTANG PILOT TRAINING MANUAL

21 CREW RESOURCE MANAGEMENT

CITATION MUSTANG PILOT TRAINING MANUAL

DECISION-MAKING PROCESS

1. Recognize the need for a decision.

Aeronautical decision making (ADM) provides a systematic approach to risk assessment. It is a tool you can use to select the best response for a given set of circumstances. FlightSafety recommends the decision-makin­g process illustrated on the second page of the Blue Card (Figure 21-4). This continuous-loop process includes eight steps:

3. Collect facts.



8. Evaluate the effects of your response.

2. Identify the problem and define it in terms of time and risk. 4. Identify alternative responses to the need. 5. Weigh the impact of each alternative ­response. 6. Select a response. 7. Implement that response.

Figure 21-4.   Decision Making Process

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WALKAROUND

CITATION MUSTANG PILOT TRAINING MANUAL

WALKAROUND The following section is a pictorial walkaround. Each item listed in the exterior power-off preflight inspection is displayed. The general photographs contain circled numbers that correspond to specific steps displayed on the subsequent pages.

FOR TRAINING PURPOSES ONLY

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PRELIMINARY EXTERIOR INSPECTION

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5 4

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3

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1.   BATTERY - CONNECTED

2.  ENGINE COVERS (4) - REMOVED

3.   PITOT COVERS (2) - REMOVED

4.  STATIC WICKS COVERS - REMOVED

5.   GROUND POWER UNIT - NOT CONNECTED

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COCKPIT/CABIN INSPECTION

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3

24 16

23

5 14 9

7 22 21 19 20 18 8 17

10

25 11

6 13 15

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2

1.  DOCUMENTS, MANUALS AND CHARTS - CHECK ABOARD a. TO BE DISPLAYED IN THE AIRPLANE AT ALL TIMES: 1) AIRWORTHINESS AND REGISTRATION CERTIFICATES 2) TRANSMITTER LICENSE(S) (AS REQUIRED).

1.   DOCUMENTS, MANUALS AND CHARTS - CHECK ABOARD b. TO BE CARRIED IN THE AIRPLANE AT ALL TIMES: 1) FAA APPROVED AIRPLANE FLIGHT MANUAL 2) GARMIN G1000 AVIONICS COCKPIT REFERENCE GUIDE 3) OTHER APPLICABLE PILOT’S MANUALS AS REQUIRED IN SECTION III, OPERATING LIMITATIONS OR APPLICABLE AFM SUPPLEMENT.

2.   REQUIRED EQUIPMENT - ONBOARD AND SERVICED

3.  CABIN - CHECK

3.a.   EMERGENCY EXIT - SECURE/CLEAR/ LOCK PIN REMOVED/COVER IN PLACE

3.b.   PASSENGER SEATS - UPRIGHT/CONDITION

FOR TRAINING PURPOSES ONLY

WA-5

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CITATION MUSTANG PILOT TRAINING MANUAL

WALKAROUND 3.c.   EXIT PLACARDS - SECURE

3.d.  DOOR ENTRY LIGHTS - OFF

4.  PORTABLE FIRE EXTINGUISHER - SERVICED AND SECURE

5.  GUST LOCK - REMOVE AND STOW

6.  CIRCUIT BREAKERS - IN

7.  LANDING GEAR HANDLE - DOWN

WA-6

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8.   ANTISKID SWITCH - ON

9.   ALL OTHER SWITCHES - OFF OR NORM

10.   ELEVATOR TRIM - CHECK/SET (TRIM INDICATOR WITHIN TO RANGE)

11.   THROTTLES - CUTOFF

12.   EMERGENCY GEAR RELEASE HANDLE - STOWED AND COVER INSTALLED

13.   BATTERY DISCONNECT SWITCH - DISCONNECT

FOR TRAINING PURPOSES ONLY

WA-7

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WALKAROUND 14.   BATT SWITCH - BATT (ALL DISPLAYS OFF)

15.   BATTERY DISCONNECT SWITCH - NORMAL/COVER DOWN (PFD 1/2 AND MFD POWERED)

16.  GROUND POWER UNIT (IF DESIRED) - CONNECTED

17.   PARKING BRAKE - SET

18.   PITOT-STATIC SWITCH - PITOT STATIC (30 SECONDS); OFF

19.   LANDING LIGHT SWITCH - ON (CHECK ILLUMINATION; OFF, IF SEEN FROM COCKPIT)

WA-8

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20.   OTHER EXTERNAL LIGHTING SWITCHES - ON (CHECK ILLUMINATION; OFF, IF SEEN FROM COCKPIT)

21.   PAX SAFETY SWITCH - PAX SAFETY (CHECK ILLUMINATION); OFF

22. LANDING GEAR POSITION LIGHTS - THREE GREEN LIGHTS/ NO RED LIGHT

23.   DATABASE/CHART CURRENCY - CHECK

24.  FUEL QUANTITY AND BALANCE - CHECK

25.   FLAP HANDLE - AGREES WITH FLAP POSITION INDICATOR

FOR TRAINING PURPOSES ONLY

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WALKAROUND 26.   AILERON AND RUDDER TRIM - CHECK/SET

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HOT ITEMS/LIGHTS

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WALKAROUND 1.  HOT ITEMS/LIGHTS - CHECK

a.   LEFT AND RIGHT STATIC PORTS (4) - CLEAR AND WARM

b.  LEFT AND RIGHT PITOT TUBES (2) - CLEAR AND HOT

c.  STALL WARNING VANE - CONDITION AND HOT

d.  LANDING LIGHTS - ON (IF NOT OBSERVED FROM COCKPIT)

e.  BEACON LIGHT - ON AND FLASHING (IF NOT OBSERVED FROM COCKPIT)

WA-12

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f.  RIGHT NAV AND ANTI-COLLISION LIGHTS - ON (IF NOT OBSERVED FROM COCKPIT)

g.   LEFT WING INSP, NAV AND ANTI-COLLISION LIGHTS - ON (IF NOT OBSERVED FROM COCKPIT)

g.   LEFT WING INSP, NAV AND ANTI-COLLISION LIGHTS - ON (IF NOT OBSERVED FROM COCKPIT)

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EXTERNAL LIGHTING SWITCHES

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2.  EXTERNAL LIGHTING SWITCHES - OFF

WA-14

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BATT SWITCH

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3.  BATT SWITCH - OFF

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LEFT NOSE

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5

WA-16

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4.   LEFT NOSE - CHECK

a.   STATIC PORTS (2) AND SURROUNDING FUSELAGE SKIN - CLEAR AND NO DAMAGE

b.  OAT PROBE INLET AND SENSORS (2) - CLEAR AND NO DAMAGE

c.   ACCUMULATOR BLEED VALVE - OPEN; BLEED DOWN; CLOSED

c.   ACCUMULATOR BLEED VALVE - OPEN; BLEED DOWN; CLOSED

c.   ACCUMULATOR BLEED VALVE - OPEN; BLEED DOWN; CLOSED

FOR TRAINING PURPOSES ONLY

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WALKAROUND d.   HYDRAULIC ACCUMULATOR PRECHARGE PRESSURE GAUGE - CHECK (PER PLACARD)

e.  HYDRAULIC RESERVOIR - CHECK FLUID LEVEL

f.  BAGGAGE DOOR - SECURE AND LOCKED

g.  NOSE GEAR, DOORS, WHEEL, TIRE AND STRUT - CONDITION

g.  NOSE GEAR, DOORS, WHEEL, TIRE AND STRUT - CONDITION

WA-18

h.  OVERBOARD VENT LINE - CLEAR

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RIGHT NOSE AND FUSELAGE RIGHT SIDE

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WALKAROUND 5.  RIGHT NOSE AND FUSELAGE RIGHT SIDE - CHECK

a.  AUX BRAKE PNEUMATIC PRESSURE GAUGE - CHECK (PER PLACARD)

b.   AUX GEAR PNEUMATIC PRESSURE GAUGE - CHECK (PER PLACARD)

c.  NOSE COMPARTMENT LIGHT - OFF

d.  BAGGAGE DOOR - SECURE AND LOCKED

e.   OXYGEN BLOWOUT DISC - GREEN

WA-20

FOR TRAINING PURPOSES ONLY

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f.  FRESH AIR INLET - CLEAR

g.   OVERBOARD VENT AND DRAIN LINES - CLEAR

h.   STALL WARNING VANE - ROTATES FREELY

I.  STATIC PORTS (2) AND SURROUNDING FUSELAGE SKIN - CLEAN AND NO DAMAGE

j.  LANDING LIGHT - CONDITION

k.   TOP AND BOTTOM ANTENNAS - CONDITION

FOR TRAINING PURPOSES ONLY

WA-21

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RIGHT WING

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8 5 1 2

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WA-22

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6.   RIGHT WING - CHECK

a.  FUEL QUICK DRAINS (4) - DRAIN AND CHECK FOR CONTAMINATION

b.  MAIN GEAR DOOR, WHEEL, TIRE AND STRUT - CONDITION

c.   EMERGENCY EXIT - SECURE

d.   WING DEICE BOOT - CONDITION

e.   STALL STRIP - CONDITION

FOR TRAINING PURPOSES ONLY

WA-23

CITATION MUSTANG PILOT TRAINING MANUAL

WALKAROUND f.   VORTEX GENERATORS (8 PER WING) - CONDITION

g.   FUEL FILLER CAP - SECURE

h.  FUEL TANK VENT - CLEAR

i.   STATIC WICKS - CHECK (3 INSTALLED. 1 MAY BE MISSING. NO MORE THAN 2 TOTAL MISSING ON ENTIRE PLANE)

j.  AILERON, FLAP, AND SPEED BRAKES - CONDITION (MAKE SURE FLAP POSITION MATCHES INDICATOR)

WA-24

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RIGHT ENGINE/NACELLE 3

1

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CITATION MUSTANG PILOT TRAINING MANUAL

WALKAROUND 7.   RIGHT ENGINE/NACELLE - CHECK

a.  ENGINE AIR INLET - CLEAR

b.  ENGINE FAN DUCT AND FAN - CHECK (FOR BENT BLADES, NICKS AND BLOCKAGE OF FAN STATORS)

c.   ENGINE T2 PROBE - CONDITION

d.   PYLON PRECOOLER INLET - CLEAR

e.   GENERATOR COOLING AIR INLET - CLEAR

WA-26

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f.   ENGINE ANTI-ICE EXHAUST - CLEAR

g.  GENERATOR COOLING AIR EXHAUST - CLEAR

h.  ENGINE FLUID DRANS - CLEAR

i.   OIL FILTER DIFFERENTIAL PRESSURE INDICATOR - NOT EXTENDED

j.  OIL LEVEL - CHECK

k.   FILLER CAP AND ACCESS DOOR - SECURE

FOR TRAINING PURPOSES ONLY

WA-27

CITATION MUSTANG PILOT TRAINING MANUAL

WALKAROUND k.   FILLER CAP AND ACCESS DOOR - SECURE

l.  ENGINE EXHAUST AND BYPASS DUCT - CONDITION AND CLEAR

m.  PYLON PRECOOLER EXHAUST - CLEAR

WA-28

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EMPENNAGE/AFT FUSELAGE 10 13

8 12 WALKAROUND

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WA-29

CITATION MUSTANG PILOT TRAINING MANUAL

WALKAROUND 8.  EMPENNAGE/AFT FUSELAGE - CHECK

a.   GROUND POWER SERVICE DOOR - SECURE

b.   AIR CONDITIONING INLET AND EXHAUST - CLEAR

c.   FAIRING VENT (BOTTOM OF AFT FUSELAGE ON RIGHT SIDE) - CLEAR

d.  OVERBOARD DRAINS/VENTS - CLEAR

e.   FADEC STATIC PORTS (L AND R) - CLEAR

WA-30

FOR TRAINING PURPOSES ONLY

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CITATION MUSTANG PILOT TRAINING MANUAL

f.  TAIL STRAKES (L AND R) - CONDITION

g.  TAILCONE AIR INLETS - CLEAR

h.   RIGHT HORIZONTAL STABILIZER DEICE BOOT - CONDITION

i.   VERTICAL STABILIZER DEICE BOOT - CONDITION

j.  RIGHT HORIZONTAL STABILIZER, ELEVATOR, AND TRIM TAB - CONDITION (MAKE SURE TRIM TAB POSITION MATCHES INDICATOR

k.   RUDDER AND TRIM TAB - SECURE (MAKE SURE TRIM TAB POSITION MATCHES INDICATOR)

FOR TRAINING PURPOSES ONLY

WA-31

CITATION MUSTANG PILOT TRAINING MANUAL

WALKAROUND

l.  STATIC WICKS (RUDDER, BOTH ELEVATORS, AND TAILCONE - CHECK (10 INSTALLED. 1 MAY BE MISSING FROM EITHER ELEVATOR AND 1 MAY BE MISSING FROM RUDDER OR TAILCONE. NO MORE THAN 2 TOTAL MISSING ON ENTIRE AIRPLANE.

n.   LEFT HORIZONTAL STABILIZER DEICE BOOT - CONDITION

WA-32

m.   LEFT HORIZONTAL STABILIZER, ELEVATOR, AND TRIM TAB - CONDITION (MAKE SURE TRIM TAB POSITION MATCHES INDICATOR)

o.   RUDDER GUST LOCK - DISENGAGE

FOR TRAINING PURPOSES ONLY

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LEFT ENGINE/NACELLE

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WALKAROUND 9.  LEFT ENGINE/NACELLE - CHECK

a.   PYLON PRECOOLER EXHAUST - CLEAR

b.   ENGINE EXHAUST AND BYPASS DUCT - CONDITION AND CLEAR

c.   OIL LEVEL - CHECK

d.  FILLER CAP AND ACCESS DOOR - SECURE

e.   ENGINE FLUID DRAINS - CLEAR

WA-34

FOR TRAINING PURPOSES ONLY

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f.  GENERATOR COOLING AIR EXHAUST - CLEAR

g.  ENGINE ANTI-ICE EXHAUST - CLEAR

h.   GENERATOR COOLING AIR INLET - CLEAR

i.   ENGINE T2 PROBE - CONDITION

j.  ENGINE AIR INLET - CLEAR

k.   ENGINE FAN DUCT AND FAN - CHECK (FOR BENT BLADES, NICKS AND BLOCKAGE OF FAN STATORS)

FOR TRAINING PURPOSES ONLY

WA-35

CITATION MUSTANG PILOT TRAINING MANUAL

WALKAROUND l.  PYLON PRECOOLER INET - CLEAR

WA-36

m.   OIL FILTER DIFFERENTIAL PRESSURE INDICATOR - NOT EXTENDED

FOR TRAINING PURPOSES ONLY

CITATION MUSTANG PILOT TRAINING MANUAL

3

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AFT COMPARTMENT

1

2

5 4

6

10.  AFT COMPARTMENT - CHECK

a.   FIRE BOTTLE PRESSURE GAUGE - CHECK PER PLACARD

FOR TRAINING PURPOSES ONLY

WA-37

CITATION MUSTANG PILOT TRAINING MANUAL

WALKAROUND b.  JUNCTION BOX CIRCUIT BREAKERS - IN

c.   EQUIPMENT AND JUNCTION BOX ACCESS DOORS - SECURE

d.   AFT COMPARTMENT BAGGAGE - SECURE

e.   AFT COMPARTMENT LIGHT - OFF

f.  AFT COMPARTMENT ACCESS DOOR - SECURE AND LOCKED

WA-38

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LEFT WING

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6

5

1

2

3

9 8

FOR TRAINING PURPOSES ONLY

WA-39

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4

CITATION MUSTANG PILOT TRAINING MANUAL

WALKAROUND 11.  LEFT WING - CHECK

a.   FLAP, SPEED BRAKES, AILERON, AND TRIM TAB - CONDITION (MAKE SURE FLAP AND TRIM TAB POSITIONS MATCH INDICATORS)

b.   STATIC WICKS - CHECK (3 INSTALLED. 1 MAY BE MISSING. NO MORE THAN 2 TOTAL MISSING ON ENTIRE AIRPLANE)

c.   FUEL TANK VENT - CLEAR

d.  FUEL FILLER CAP - SECURE

e.   WING DEICE BOOT - CONDITION

WA-40

FOR TRAINING PURPOSES ONLY

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f.  VORTEX GENERATORS (8 PER WING) - CONDITION

g.  STALL STRIP - CONDITION

h.   MAIN GEAR DOOR, WHEEL, TIRE AND STRUT - CONDITION

i.  FUEL QUICK DRAINS (4) - DRAIN AND CHECK FOR CONTAMINATION

FOR TRAINING PURPOSES ONLY

WA-41

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WALKAROUND

FUSELAGE LEFT SIDE

1

3

2

WA-42

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12.  FUSELAGE LEFT SIDE - CHECK

a.   WING INSPECTION LIGHT - CONDITION

b.   LANDING LIGHT - CONDITION

c.   CABIN DOOR SEAL - CHECK FOR RIPS AND TEARS

FOR TRAINING PURPOSES ONLY

WA-43

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WALKAROUND

INTENTIONALLY LEFT BLANK

WA-44

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CITATION MUSTANG PILOT TRAINING MANUAL

APPENDIX A ANSWERS TO QUESTIONS

CHAPTER 2 1. A 2. B 3. A 4. A 5. B 6. C 7. A 8. C 9. C 10. C 11. B 12. B CHAPTER 3 1. C 2. A 3. D 4. B CHAPTER 4 1. D 2. D 3. C 4. D 5. A 6. A 7. B 8. C

CHAPTER 5 1. A 2. D 3. B 4. A 5. B 6. C 7. B 8. A 9. B 10. B 11. B 12. A 13. C CHAPTER 7 1. B 2. A 3. B 4. B 5. A 6. A 7. A 8. D 9. B 10. C 11. A 12. D 13. D 14. B 15. B 16. A 17. A CHAPTER 8 1. A 2. A 3. B 4. A 5. D 6. B CHAPTER 9 1. B 2. C 3. B

Revision 1.1

CHAPTER 10 1. A 2. B 3. A 4. A 5. C 6. A 7. B 8. A 9. B 10. B 11. C 12. A 13. C 14. B 15. C CHAPTER 11 1. A 2. A 3. B 4. B 5. B 6. A 7. B 8. D 9. B 10. A 11. D CHAPTER 12 1. A 2. A 3. C 4. D 5. A 6. B 7. D 8. C 9. B 10. C

FOR TRAINING PURPOSES ONLY

CHAPTER 13 1. A 2. B 3. B 4. A 5. C 6. A 7. D 8. C CHAPTER 14 1. B 2. A 3. A 4. A 5. D 6. A 7. D 8. B 9. A 10. B 11. C 12. A CHAPTER 15 1. A 2. C 3. B 4. B 5. C 6. B 7. C 8. A 9. A 10. D 11. C

APPA-1

APPENDIX A

CHAPTER 1 1. A 2. B 3. C 4. B 5. B 6. D 7. D

CITATION MUSTANG PILOT TRAINING MANUAL

CHAPTER 16 1. C 2. A 3. B 4. B 5. D 6. D 7. C 8. C 9. B 10. A 11. D

APPENDIX A

CHAPTER 17 1. A 2. A 3. A 4. A 5. A 6. C 7. A 8. A 9. C 10. D 11. B 12. B

APPA-2

FOR TRAINING PURPOSES ONLY

Revision 1.1