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AVIONICS DIVISION PHOEIWX, ARIZONA

IMPORTANT

NOTICE

The avionicsbusiness unitsformerlyowned by UNISYS Corp. and identified with the Sperry name or logo have been acquired by products,and components marked Honeywell Inc. Publications, or identified hereinwiththe Sperry name or logo are publications, products,and components of Honeywell Inc. Allreferencesto the Sperry name or logo should be taken as referring to Honeywell Inc.

SHZ-412 Integrated Flight Control System Bell 412

Maintenance Manual 22-15-00 PRINTED INU.S.A.

WB.

NO 09.1 I69-10

1

MAY 198!

MAINTENANCE MANUAL

+5TAERW

FLIGHT SYSTEMS

AVIONICS

DIVISION

BELL

412

RECORD OF REVISIONS Retain this record in front of manual. On receipt of revision, insert revised pages in the manual, and enter revision number, date inserted and initial. Revision Number

Revision Date

Insertion Date

By

Revision Number

Revision Date

Insertion Date

By

22-15-00 Page RR-1 ?&y 1/81

+>TdERw

MAINTENANCE MANUAL

FLIGHT SYSTEMS

AVIONICS

DIVISION

BELL

412

LIST OF EFFECTIVE PAGES - Original....O....May 1/81 SUBJECT ——

PAGE

Title

SUBJECT

PAGE 6 Blank 7 8 Blank

Record of Revisions List of Effective Pages : 3 4 Blank Contents 1 : 4 Blank List of 11Iustrations ; List of Tables ; Blank Introduction ; : Systm Description 1 : 4 Blank 5

Component Description 101 102 103 104 105 106 107 108 109 110 111 112 113 114 115 116 117 118 119 120 121 122 123 124 125 126 127 128 129 130 131 132 133 134 135

Blank B1ank

B1ank Blank

This manual was released for distribution May 26/81.

22-15-00 List of Effective Pages (Page 1) hy 1/81

+sT-ER~ FLIGHT SYSTEMS AVIONICS

SUBJECT —.

PAGE *

Component Description (cent) 136 137 138 139 ;$; Blank 142 143 144 145 146 147 148 149 150 151 152

Blank Blank Blank

Blank

System Operation 201 202 203 204 205 206 207 208 209 210 211 212 Blank 213 214 BIank 215 216 BIank 217 218 Blank 219 220 Blank 221 222 B1ank 223 224 B1ank 225 226 Blank

DIVISION

MAINTENANCE MANUAL BELL

412

—SUBJECT

.PAGE 227 228 229 230 231 232 233 234 235 236 237 238 239 240 241 242 243 X: 246 247 248 249 250 251 252 253 254 255 256 257 258 259 260

B1ank

B1ank Blank Blank B1ank B1ank Blank Blank

B1ank BIank Blank Blank

Ground Operational Tests (Ground Chec;) 302 B1ank 303

306 307 308 309 310 311

22-15-00 List of Effective Pages (Page 2) hj’ 1/81

MAINTENANCE MANUAL AVIONICS

SUBJECT ——

PAGE

REV. NO.

Ground Operational Tests (Ground Ch::~) (cent) 313 314 315 316 Blank Fault Isolation 401 402 403 404 405 406 407 408 409 410 411 412 413 414 Blank Interconnects 501 502 503 504 505 506 507 508 509 510 511 512 Schematics 601 602 603 604 605 606 607 608 609 610

: 0 0 0

0 0 0 0 0 0 0 0 0 : : 0 0

Blank

DIVISION

; o 0

Blank : Blank 1?

BELL

412

—SUBJECT

—PAGE 611 612 613 614 615 616 617 618 619 620 621 622 623 624 625 626

Blank Blank Blank Blank Blank Blank Blank Blank

Removal/Reinstallation and Adjustment 701 702 703 704 705 706 707 708 709 710 Shipping, Handling, and Storage 801 80261 ank

Blank : Blank

0 Blank : Blank : Blank : Blank Blank

; o

22-15-00 List of Effeetive Pages (Page 3/4) My 1/81

MAINTENANCE MANUAL

~”E%?~ FLIGHT SYSTEMS AVIONICS

BELL

412

DIVISION

TABLE OF CONTENTS

Section 1

Paraqraph

1 1

General System Description

Component Description

101

1. 2.

General Sensors

101 102

A. TARSYN Three-Axis Reference B. FX-220 Flux Valve C. CS-412 Remote Compensator D. RG-203 Rate Gyro E. RT-220 Radio Altimeter Receiver Transmitter F. AC-702 Altitude Sensor G. AS-702 Airspeed Sensor

102 110 112 114 116 120 122

Conputers

124 ‘

A. B. C.

124 126 128

3.

4.

5.

6.

FZ-702 Flight Director Conputer SP-711 Helipilot Computer TZ-701 Trim”Coqwter-

134

Indicators A.

3

Page 1

System Description 1. 2.

2

Subject

RA-335 Radio Altimter

Indicator

134

Selectors and Controllers

138

A. B.

138 142

MS-702 Mode Selector PC-702 Helipilot Controller

Actuators

150

A.

150

Control Rod Assemblies

System Operation

201

1. 2. 3.

201 202 204

General System Performance/Operating Limits H%lipilot Functional”Description A. B.

Helipilot System Operation Helipilot Computer Functional Description

204 206

22-15-00 Contents (Page 1) May 1/81

MAINTENANCE MANUAL

++T=EFw FLIGHT SYSTEMS A\/[~NIcs

BELL

412

DIVISION

TABLE OF CONTENTS (cent)

.— Section 3

Paragraph

Flight Director Functional Description

229

A. B.

229 231

Flight Director Mode Conditions or Functions Flight Director Mode Flow Description

TARSYN Conpass System Operation Power Distribution

251 255

Ground Operational Tests (Ground Check)

301

1. 2. 3.

301 301 301

~. .

5

Paqe

System Operation (cent) 4.

4

W&Q

General Equipment and Materials Procedure

Fault Isolation

401

1. 2.

401

General Procedure

401

6

Interconnects

501

7

Schematics

601

8

Removal/Reinstallation and Adjustment

701

1. 2. 3.

701 701 701

General Equipnmt and Materials Procedure for Sensors A. B. i E. F. G. H.

4.

TARSYN Three-Axis Reference FX-220 Flux Valve Conpass Swing and Adjustnmt CS-412 Remote Compensator RG-203 Rate Gyro RT-220 Radio Altimeter Receiver Transmitter AC-702 Altitude Sensor AS-702 Airspeed Sensor

701 702 703 704 705 706 706 707

Procedure for Computers

707

A. B.

707 708

Flight Director and Helipilot Computers TZ-701 Trim Computer

22-15-00 Contents (Page 2) May 1/81

MAINTENANCE MANUAL

~>~”ER?~ FLIGHT SYSTEMS AVIONICS

BELL

412

DIVISION

TABLE OF CONTENTS (cent)

Section 8

Paragraph

Paqe

Removal/Reinstallation and Adjustment (cent) 5.

6.

7. 9

&!!&l

Procedure for Indicators

708

A.

RA-335 Radio Altimeter Indicator

708

Procedure for Selectors and Controllers

708

A. B.

MS-702 Mode Selector PC-702 Helipilot Controller

708 709

Procedure for Control Rod Assemblies

710

Shipping, Handling, and Storaqe

801

22-15-00 Contents (Page 3/4) May 1/81

MAINTENANCE MANUAL

+SPER?f

FLIGHT SYSTEMS

AVIONICS

DIVISION

BELL

412

LIST OF ILLUSTRATIONS Page

Name

M!E 1

System Flow Diagram

3

2

Component Locations

7

101

TARSYN Three-Axis Reference

102

102

TARSYN Block Diagram

105

i03

FX-220 Flux Valve

110

104

FX-220 Flux Valve Schematic

111

105

CS-412 Remote Compensator

112

106

CS-412 Remote Compensator Block Diagran

113

107

RG-203 Rate Gyro

114

108

RG-203 Rate Gyro Block Diagran

115

109

RT-220 Radio Altimeter Receiver Transmitter

116

110

RT-220 Radio Altimeter Receiver Transmitter Block Diagran

118

111

AC-702 Altitude Sensor

120

112

AC-702 Altitude Sensor Block Diagran

121

113

AS-702 Airspeed Sensor

122

114

AS-702 Airspeed Sensor Block Diagran

123

115

FZ-702 Flight Director Computer

124

116

SP-711 Helipilot Computer

126

117

-IZ-701Trim Computer

128

118

TZ-701 Trim Computer Schematic

129

119

RA-335 Radio Altimeter Indicator

134

120

RA-335 Radio Altimeter Indicator Block Diagram

136

121

MS-702 Mode Selector

138

22”15-00 List of 11lustrations (P;g;,~j

MAINTENANCE MANUAL

+S=ERW

FLIGHT SYSTEMS

AVIONICS

DIVISION

BELL

412

LIST OF ILLUSTRATIONS (cent) Page

Name

EQ!!!z 122

MS-702 Mode Selector Schematic

139

123

PC-702 Helipilot Controller

142

124

PC-702 Helipilot Controller Schematic

145

125

Typical Control Rod Assembly

150

126

Linear Actuator Block Diagram

151

201

Helipilot No. 1 Engage Logic

213

202

Helipilot No. 2 Engage Logic

215

203

llelipilotRoll Axis Signal Flow Diagram

217

204

Helipilot Pitch Axis Signal Flow Diagram

219

205

Helipilot Yaw Axis Signal Flw

221

206

Roll Axis Automatic Trim Signal Flow Diagrm

223

207

Pitch Axis Autanatic Trim Signal Flow Diagran

225

208

Cyclic Control (Pitch or Rol1) Diagram

227

209

Wde

239

210

Heading Select (HDG) and Go-Around (GA) Mode Flow Diagram

241

211

VOR/LOC Navi ation (NAV), VOR Approach (VOR APR), and Back Course 7BC) Mode Flow Diagram

243

212

Glide Slope and Auto-Level (ILS) Mode Flm

245

213

Altitude Hold (ALT), Airspeed Hold (IAS), Vertical Speed Hold (VS), and Go-Around (GA) Mode Flow Diagran

247

214

System Failure Warning/Valid Logic Diagran

249

215

Canpass System Signal Flow Diagran

253

216

DC Pov@r Distribution Diagram

257

217

AC Powr

259

501

System Interconnect Diagram

601

System Schematics

Diagra

Selection Logic Diagram

Distribution Diagran

Diagran

502

22-15430 List of Illustrations (page 2) my 1/81

‘

MAINTENANCE MANUAL

~s~”E%6~ FLIGHT SYSTEMS AVIONICS

DIVISION

BELL

412

LIST OF TABLES Name

Table

Page

1

System Components

2

2

Required Equipment not Supplied by Sperry

5

101

TARSYN Three-Axis Reference Leading Particulars

103

102

FX-220 Flux Valve Leading Particulars

110

103

CS-412 Remote Compensator Leading Particulars

112

104

RG-203 Rate Gyro Leading Particulars

114

105

RT-220 Radio Altimeter Receiver Transmitter Leading Particulars

116

106

AC-702 Altitude Sensor Leading Particulars

120

107

AS-702 Airspeed Sensor Leading Particulars

122

108

FZ-702 Flight Director Computer Leading Particulars

125

109

SP-711 Helipilot Computer Leading Particulars

127

110

lz-7ol Trim Computer Leading Particulars

128

111

RA-335 Radio Altimeter Indicator Leading Particulars

134

112

MS-702 Mode Selector Leading Particulars

138

113

PC-702 Helipilot Controller Leading Particulars

142

114

Linear Actuator Leading Particulars

150

201

system Performance/Operating Limits

202

301

Ground Check

303

401

Fault Isolation Procedures

403

22-15-00 List of Tables (Page 1/2) May 1/81

+-W

FLIGHT SYSTEMS AVIONICS

MAINTENANCE MANUAL BELL

412

DIVISION

INTRODUCTION

This manual provides general system maintenance instructions and theory of operation for the SHZ-412 Integrated Flight Control System for the Bell 412 Helicopter. The System components described in this manual are manufactured by Sperry Flight Systems Avionics Division, Phoenix, AZ. Comnon System maintenance procedures are not presented in this manual. established shop and flight line practices should be used.

The best

Sperry has an Airworthiness Analysis procedure performed for all its airborne products to ensure that equipment designed by Sperry will not create a hazardous in-flight condition. As a result of the Analysis, certain installations have been designated INSTALLATION CRITICAL, and 100 percent compliance with those installations is required. INSTALLATION CRITICAL is defined as: Specific methods of installation are required to ensure that either the failure of the assembly or part is extremely improbable or that its failure cannot create a hazardous condition. Abbreviations used in this manual are defined belW. Equivalent

Abbreviation ACTR

Actuator

ADI

Attitude Director Indicator

AFCS

Automatic Flight Control System

ALT

Altitude

ALTM

Altimeter

ALTE

Altitude Error

ANN

Annunciator

Aoss

After Over Station Sensor

APR

Approach

AS, A/S

Airspeed

22-15-00 Introduction (Page 1) May 1/81

+L>TAERW FLIGHT SYSTEMS AVIONICS

DIVISION

MAINTENANCE MANUAL BELL

412

Abbreviation

Equivalent

ATT

Attitude

BC

Back Course

CAP

Capture

CAS

Command Augmentation System

CBB

Collective Bar Bias

CE

Course Error

CLIV

Collective

cm

Command

CONT

Controller

CPL

Couple

CRS

Course

CUR

(Xrrent

Cx

Control Transform

CY

Cyc1e

DCPL

Decouple

DEC, DECR

Decrease

DEG

Degree

DEV

Deviation

DG

Directional Gyro

DTY

Duty

ERR

Error

EXC

Excitation

FD, F/D

Flight Director

FDBK

Feedback

FDC

Flight Director Computer

FTR

Force Trim Release

GA, GIA

Go Around

GEN

Generator

GS, G/S

G1ide S1ope

HELI, HP

Helipilot Computer

HDG

Heading

HpR

Progmmd

IAS

Indicated Airspeed

IASE

Indicated Airspeed Error

-

Altitude

22-15-00 Introduction (Page 2) my 1/81

MAINTENANCE MANUAL

~s~”ERT-ERT-ERW

FLIGHT SYSTEMS

AVIONICS

BELL

412

DIVISION

Dimensions (maximum): Length ................................*...*...* .. 16.08 in. Width ...****..*......**..*..****..**........*....*. 6.53 in. [::::: :] Height ............................................. 7.38 in. (187.4 mm) ●

Weight (nominal)

●

.**.*..**.*.*....*****. F*.*.*.*... ..*.*** 15 lb (6.8 kg) ●

●

Power Requirements: Starting ****.**.*...*.***.....***............... 115 V, 400 Hz, 120 VA Operating ....****.*. ..**.....*........... ....... 115 V, 400 Hz, 70 VA ●

●

●

Vertical Gyro Characteristics: Gyro Rotor Speed ............................................ 22,000 rpm Gyro Erection Time ..*..***..* *O.*.***** .9...*.*.. 0............ 3 min Verticality Error .*.*****..*.*.****...* **...*...*............ 0.25 deg Fast Erection Rate (rein imum) ...*..*...*..00.0............... 20 deg/min S1ow Erection Rate (nominal) ........... .........*......... 2.5 deg/min VG Signal Output Data (pitch and rol1) .....**.** Three-wire synchro output of 200 mV ac/deg, Two-wire transformer output of 200 mV ac/deg, Two-wire transformer output of 50 mV ac/deg ●

●

●

●

●

●

Directional Gyro Character stits: Gyro Rotor Speed ............................................ 23,000 rpm Slav~ng Accuracy ................................................ *2 d~g Slavlng Rate (nomial) ..**.*....* O.....*.** ...****. 2.5 to 5.0 deg/mln Gyro Free Drift Rate (exclusive of earth rate) *99..****.. * 524 deg/hr Slew Rate (automatic fast or manual) *.*9.***..* .00.0.000.0 30 de;/~e: Gyro Gimbal Freedom (Azimuth) ***O***.*...***..**** ........... Gyro Gimbal Freedom (Pitch and Rol1) ******.**.* ............*. ~85 deg ●

●

●

●

●

●

●

●

●

●

Mating Connectors: 1J2 ....*...*..9....**..**..*.****-**.................8.. MS3126F22-55SY 1J3 .*.*....*...*...*..***..*..***.*...*.****... ..*****. MS3126F18-32SY ●

Mounting ..................................... Tray, Sperry Part No. 2594816 TARSYN Three-Axis Reference Leading Particulars Table 101

22-15-00 Page 103 May 1/81

+-TSERW

FLIGHT SYSTEMS

AVIONICS

DIVISION

MAINTENANCE MANUAL BELL

412

The TARSYN Three-Axis Reference provides three-wire synchro output signals that are electrical analogs of the aircraft heading and pitch and roll attitudes. The three-wire output signals are used to drive flight instruments. In addition, two-wire transformer isolated pitch and roll output signals can be used to supply attitude reference information for Flight Director, Helipilot, and radar stabilization use. The vertical gyro roll gimbal freedom is unlimited (i360 degrees) and the pitch gimbal freedom is *8O de9rees= Gwo verticality 1S maintained by gravity-sensing liquid switches and toque motors. Initial gyro erection is accomplished by sequencing voltages tothe torque motors and gyro rotor until the rotor achieves 85 percent of its speed and the spin axis is essentially vertical. At this time, the valid flag circuit is activated. An internal erection cutoff circuit turns off roll erection when bank angles exceed 6 degrees. This circuit prevents the gyro from driving to a false vertical that is sensed by the gravity-sensitive liquid switches while in a turn. If the gyro goes off-level after initial start-up by more than 6 degrees for any reason, a remote “VG Fast Erect” switch can be used to reerect the gyro. The directional gyro provides three-wire synchro heading information to various flight instruments, the Helipilots, and navigation receivers. A renotely mounted MAG/DG switch controls canpass slaved or nonslaved modes of operation. A remotely mounted slew switch (+, “) is USed to align gyro heading in case of loss of the slaving amplifier.

22-15-00 Page 104 May 1/81

RUE!*NCE

>1-kl=

PITCH FO OUTPUT ,2OO mvACDEG,

>BB~ ! >cc>~

I (c)

1]

($4)

1

(cl

1>

}

k ~

}

ROLL FO OUTPUT (200 mVAC/OEG)

y+

! I --—

DO>~ 1; I

(H) 2K

~

I

$j3Jj=}HR~TpuT !:: (c)

tt ROLL 2-WIRE SIGNAL TRANSFORMER

- T

10

PITCH 2-wIRE SIGNAL TRANSFORMER

ii

Rol !. TOREIJER

—--

~

t_

i[

.-—

I

I

I

L—-—-—-

!!

II

I

w’~”nz

I

..-

1 I I

I I

1

~-

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fi~,d,t i I--f

●

i.

PITCti

TORCJUER DRIvE TP

I

i 1’ ,

+EiH+F’~::::w’TcH

‘ITCH :RECT IMPL ROLL ERECTION CUT-OFF CIRCUIT

ROLL ERECT AMPL

1 TO 1J3.G

-u

--D&l-+

#-MONITOR 1) FAST ERECT .-J 2) uNOER VOLT 3) WHEEL SPEED

Q

+

I ERECTION CUT-OFF OISABLE

I

~

-p

.27 VOC

1, L~ I i I I I +----

I

ROLL TOROUER ORIVER TP

1’

ExT lTECO INTLK

!> K6

H+-

RECO CONTROL (NORMALLV JUMPERED)

I ++

l)

1

TO EXT RECO SWITCH

l!

A

-,

I

1>n~

e

l-’+

1>

}

v

3 N ~~

EXT FAST ERECT

%

‘uTOp’LOT ‘NTERLOcK ~ I

1,2

‘“puT ‘OwER +-(8(! ‘c) (H)

115V 403Hz

.__

pi

:1

i 1

REFERENCE

+---m

I

~~

401 ATTITuDE

rn~,~,;,~-—-—-—-—-—-—

I

{

cHAss’sGNo ‘i%

T

+

I t u-POWER SUPPLY

VERTICAL GYRO SECTION

TARSYN B1ock Diagra Fiqure 102 (Sheet 1 of 2)

22-15-00 Page 105/106 May 1/81

p-j--

26 v

4W

Hz OUTPUT

5 v 400 HZ OUTPUT

I

‘>’+-

2;v DC OUIE’U1 AD t6W @

+“SPERjI *P

I

k

,x,

; > R;

;

(Ii)

! >H>

;

(z)

p

I

SLAVED/FREE SWITCH

1

FLUX VALVE CT INPUT

+’+

i_k&-_-J

‘“”d”’”L.

r Q4j;

~ I

I

Pt ––-.

SYNCHRONIZER _— -

___

___

-—

--

-—

--—

HEADING SY;IU:O

L–-

-—-

I —-.

—

_—

__

_

I

——.

——— —,—— ___

22-15-00 Page 107/108 May 1/81

___,

J

{4x

~

(+Z

+)

~>

——— ..—

26 v

ROTOR EXCITATION

HEADING

SYNCHRODUTP”T

‘}

1

) 2’

s>+. (y)

__

>

~“~

04 -_—-

L-

)

- Ic),>y> ;

PART OF BASE ASSEMBLY

TARSYN Block Diagrm Figure 102 (Sheet 2)

SWITCH

l+

v ‘OTOR

HEADING

‘xc’TAT’ON

SYNCHROOUTPUT

I)b ;~

——— 4

1 AD. 1659 &

MAINTENANCE MANUAL

~j~”ER~”ER?f FLIGHT SYSTEMS AVIONICS

2.

c.

t3ELL 412

DIVISION

CS-412 Remote Compensator (See figures 105 and 106, and table 103.)

7 Ao-ssl

CS-412 Remote Compensator Figure 105

llimensions(maximum): .....0 ..*.000.. ........*... *... .. .. 5.62 in. (142.7 mm Length Width .......... ...**.... *...*.... ..*............ 2.56 in. (65.0 mm 1 2.99 in. (75.9 mm) Height ...... 0...... *...*. ..... ....... ..*.**0 ●

●

● ●

● ●

●

●

Weight (maximum) Powr

●

●

●

●

●

●

●

●

●

●

●

●

●

●

●

00**9*.*** *.**....*. .*.*..****...*..*** 1.0 lb (0.45 kg) ●

●

Requiranents (From TARSYN DG) ....**..***.*.*...***..***.

26 V, 400 Hz

Mating Connectors: ..0000.-..0..*...*..*..........0.......*..*....0....00 MS3126F14-19SX ....**.*.**......**-*-..-...0.0.0...0.....-- *.*.**.** MS3126F14-19SY ●

Mounting ................................................. 8ase Flange Mount CS-412 Remote Compensator Leading Particulars Table 103

22-15-00 Page 112 May 1/81

MAINTENANCE MANUAL

~~>~-ER~~ FLIGHT SYSTEMS AVIONICS

BELL

412

DIVISION

The CS-412 Remote Compensator compensates the flux valves by inserting small dc currents to cancel the errors caused by aircraft magnetic disturbances.

SCREWDRIVER ADJUSTMENTS

r ——— I J1

(c)

\ E

(:’

I I

I

mm

I REGULATED DC POWER SUPPLY

I

‘T-, I

FLUX VALVE NO. 1

1

+6.8VOC

(H)

D 26V.400HZ INPUTPOWER

.

1

J L NOTE:

.

—

.—

I I

L CURRENT LIMITING RESISTORS

—----J ——— ——— ——— ——— THEDUAI.REMOTE COMPENS’TORCONTAINS TWOIDENTICAL COMPENSATION CIRCUIT6.CIRCUITRY ANDCONNE~lONS SHOWNAREFOR THENUMBERONE SY=EMTHROUGH CDNNECTORJ1.CIRCUITRY FORTHENUMBER TWOSWEM IS IDENTICAL TO SYSTEM NUMBER ONE. EXCEPT CONNE~lONS ARE MADE THROUGH CONNECTORJ2.

2S?66R4

CS-412 Remote Compensator B1ock Diagran Figure 106

22-15-00 Page 113 May 1/81

MAINTENANCE MANUAL

~_~”ER?~ FLIGHT SYSTEMS AVIONICS

2.

D.

BELL

DIVISION

412

RG-203 Rate Gyro (See figures 107 and 108, and table 104.) The Rate Gyro provides a rate-of-turn signal to the ADI proportional to the rate of angular displacement about an axis perpendicular to the mounting surface.

RG-203 Rate Gyro Figure 107

Dimensions (maximum): Length *...... ........................... .... ..*.. 3.32 in. 84.3 mm) Width ...****.* *.*..*....*...*.*.*....m***.*. .*-.* 2.04 in. [51.8 mm) Height =.*.*-*.* *..*.*.*..................*.*...... 1.82 in. (46.2 nun) ●

●

●

●

●

●

●

●

Weight ..*...**.*..*....*..*...*...**..* ..*....*.. ...**.. ●

●

1.0 lb (0.45 kg)

Motor Excitation (400 Hz) ==*.*.***.0.00.000-.0 ..*-..*.*..*......*.* Powr (runnirtg).**...**..* *...*....* *.****.*.. *****...*. *.*. Runup time (maximun) 0000000..0 .*..*....*...................... Phase ........................................................... ●

●

●

●

●

●

●

●

Sense, polarity (input axis down, cw rate) (rate viewed from top) ...*.**.**................... Mating Connector .............................................

Positive

Ted Mfg B1700

Mounting ...*....*.*.*..*-...*...............0................... RG-203 Rate Gyro Leading Particulars Table 104

26 V ac 6.5 VA 90 sec Single

Base Mount

22-15-00 Page 114 May 1/81

+5T 6}

GYRO

VALID

OU7PW (NOTE 11

NOTE Wno GYRO

v&uo = 4.s INVAUO

● 0.5 Voc = -

Altitude Trip Switches ........

60 mA current sink provided at and below trip points indicated below:

●

Trip Point

Accuracy *4ft *10 ft Ho ft

50 ft 250 ft 1200 ft Mating Connectors:

TRANSMIT .............c................*....*............ GRFF 4007-0002 (GRFF Connectors, GRFF Division, Solitron Devices, Inc) RECEIVE ................*...*............................ GRFF 4007-0002 MS3126F16-26S J1 * . *. ...---0 . * . *..*- *.....** 9***. .0 ●

●

●

●

● ● ●

●

●

●

●

● ●

Mounting ...................*...... .. ●

●

● ●

●

●

●

●

....**......0 0 ●

● ● ●

● ●

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Base Flange Mount

RT-220 Radio Altimeter Receiver Transmitter Leading Particulars Table 105 (cent) The Radio Altimeter Receiver Transmitter provides a dc output voltage which is proportional to the aircraft absolute altitude above terrain. In addition, it provides radio altitude trip points, an indicator warning flag output, and an auxiliary radio altitude output. The recision output is used to drive the Radio Altimeter Indicator and supp1’ ies absolute altitude information to the flight director system. The auxiliary output drives the ADI rising runway display.

22-15-00 Page 117 May 1/81

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Indicators A.

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Radio Altimeter Indicator Leading Particulars Table 111

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The RA-335 Radio Altimeter Indicator displays absolute altitude frcxnO to 1500 feet. The Indicator has the following features: (1) Decision Height Annunciator - The DH annunciator lights when the helicopter descends to or below the altitude indicated by the position of the DH cursor. (2)

Mask - The mask provides a means of keeping the pointer out-of-view for altitudes above 1500 feet.

(3)

Failure Warning Flag - The OFF flag canes into view when the altitude signal from the Receiver Transmitter becomes invalid.

(4)

Decision Height Set Knob and Cursor - The orange DH cursor is positioned on the altitude dial by the DH SET knob to select a decision height altitude.

(5)

Pointer and Altitude Scale - The pointer displays altitude above the terrain.

(6)

Self-Test Button- When pressed, the TEST switch causes the pointer to indicate 100 x 20 feet, the DH annunciator to illuminate if set above 100 feet, and the OFF flag to drop into view.

22-15-00 Page 135 May 1/81

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22-15-00 Page 147/148 May 1/81

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Actuators

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22-15-00 Page 150 May 1/81

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MAINTENANCE MANUAL

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AVIONICS

DIVISION

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412

The Control Rod Assembly Actuator is a motor-driven jackscrew, extendable piston type.device. The actuator is installed in the aircraft control linkages in a series arrangement. Electrical canmands from the Helipilot Compfiers drive the motor of the actuator to extend or retract the control linkages which position the aircraft controls to command the desired maneuver. Position feedback to the Helipilot servo loop is provided by a 1inear variable differential transformer (LVDT). An electromechanical brake releases the motor for operation when power is applied at HP engagement.

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22-15-00 Page 215/216 Revised Mar 20/86

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S-EIEr~~ FLIGHT SYSTEMS AVIONICS

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(6)

DIVISION

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412

Flare Command Sensor When making an approach in the ILS mode, the auto-level mode will engage at an altitude of approximately 75 feet above the terrain. This w“ll cause the aircraft to auto-level and level off at a radio altitude of 50 feet. The auto-level mode is indurative when the radio altimeter system is inoperative.

(7)

Flight Director Valid The flight director valid signal is used to drive the AOI flight director warning flag and internal logic circuitry in the Flight Director Computer. ‘me valid signal is generated by a circuit that monitors the input voltage to the FDC and the *1O volt power supply. ~~ A 28-volt dc output is supplied whenever the power supply monitors indicate valid.

(8)

ADI Roll or Pitch Command Bar When an error canmand signal is applied to the bar input, the bar moves left or right (roll) or up or down (pitch) to indicate a canmand to the pilot to maneuver the aircraft in the indicated direction to reduce the generated error signal to zero. Dual ADIs with flight director bars are installed in-this aircraft. Both sets of bars are driven by the one Flight Director Computer.

(9)

Rol1 Bar Bias (RBB) The RBB signal drives the ADI roll canmand bar out of view whenever the RBB logic is tresent. RBB louic will be Dresent anvtim the flight dir~ctor valid logic is in~alid, the DG or VG is-invalid, or the standby mode is selected. RBB will also result if the NAV valid signal is not present when the NAV mode is engaged.

(lo)

Pitch Bar Bias (PBB) The PBB signal drives the ADI pitch cumnand bar out of view whenever the PBB logic is present. PBB logic W-ll be present anytime the flight director logic is invalid, the VG is invalid, or the standby mode is selected. If the GS mode is engaged, the absence of the GS valid signal will also cause a PBB condition.

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Flight Director Mode Flow Description ~~p~~:ow”ng . .

modes of operation are used tith the Flight Director

Standby (SBY) Lateral (Roll) Axis Modes Heading Select (HDG) VOR/Localizer (NAV) VOR Approach (VOR APR) Back Course (BC) Go-Around (GA) Longitudinal (Pitch) Axis Modes G1ide S1ope and Auto-Level (ILS) Altitude Hold (ALT) Airspeed HoId (IAS) Vertical Speed Hold (VS) Go-Around (GA) (1)

Coupled Operation (See figures 210 thru 213.) Whenever the System is coupled, the same inf nnation driving the AD I ccmmand bars is routed via smnation point 2 to the Helipilot 6 Computers for automatic flight path control. These pitch and roll signal paths contain integral control circuitry 3 which reinforce Q the small residual errors into signals large enoug to drive the Helipilots to correct any system standoff errors. When the system hot coUPl~, these integrators are set to zero by the PCPL or RCPL signals fran the Helipilot Computers. the Helipilot System operates the autotrim actuators, the pitch and roll position potentiometers detect stick movement and feed signals to the Helipilot Computers so that the Helipilot can balance the control input into the swashplate, thus ending up with centered actuators and no path disturbance. When

(2)

Standby (SBY) Mode (See figure 209.) When no mode has been selected or SBY is pressed on the Mode Selector or the cyclic control, the ADI canmand bars are retracted from view. In this mode, SBY is annunciated on the Mode Selector and the Flight Director System is ready for operation if the FD flag on the ADI is out-of-view. Pressing SBY provides a tamp test mode for the Mode Selector and Helipilot Controller and cancels any other flight director modes.

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412

(3) Lateral (Roll) Axis Modes (a)

Heading Select (HDG) Mode (See figure 210.) The heading select mode is engaged by pressing the HDG switch on the Mode Selector. The annunciator will show ON. The canputer then generates the necessary roll ccmnands to bank the helicopter to intercept and maintain the heading selected on the HSI. A heading error signal, composed of the difference between aircraft heading and the selected heading, is routed through a demodulator and bank angle limiter to sumnation point @. Roll attitude 4 fran the TARSYN-Three-Axis Reference 1s amplified, demo9 ulated and also routed to sunmation point @ . The roll attitude and command signals are of opposite polarlty and thus try to cancel each other. Any difference (error signal) is then amplified and routed to the ADI(s) roll canmand bar. The bank angle limiter limits the commanded bank angle to t20 degrees. ADI roll bar displacement gains are set to provide easy pilot interpretation. Roll bar bias (RBB) is a fixed signal also summed at point @ and is switched on if the VG, DG, or FDC develop an invalid condition. This signal drives the roll command bars fully out of view and prevents the pilot fran following erroneous comnands.

(b)

VOR Mode (See figure 211.) The VOR mode provides intercept, capture, and tracking of a selected VOR radial. To use the mode, the NAV radio is tuned to the desired VOR facility frequency, and the desired course to the station is set on the HSI ~. The HSI then displays, pictorially, the relative position of the aircraft to the station and desired ground track. The aircraft must be flown initially on a heading that wil1 intercept the selected course. The heading select mode 5 is normally used to establish the intercept. Pressing the 9 V R mode switch causes the system to ‘ARM” if outside the normal capture range [approximately 5 degrees (one dot) depending on flight condition]. When in “ARM”, the system continues to flY the heading select mode until the capture point is reached. At this point, the lateral bean sensor @ trips, the heading mode light goes out, the VOR mode annunciator shows “CAP”, and the roll command bar displays a turn command toward the station. If coupled, the Helipilot follows the command and steers the aircraft to zero the ccmnand bar. The stable banked condition is a 20-degree angle maintained until nearing the selected

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course, at which time the com uter calls for a rollout onto course. If a long way frcm t Ee station, the initial maneuver is to establish a 30-degree cut @ at the selected course until near bem center, then a second turn is commanded to the selected course and beam center. align wi-th When the aircraft is aligned w“th the course within 15 degrees and the roll angle is less than 6 degrees (NOC), the systm bank limit is reduced to 10 degrees and crosswind correction is enabled. This correction is accomplished by washing out course error through a 200-second filter @ until beam center is reached. The aircraft will then track this course, constantly correcting for disturbances. As the station is approached, the radio bea becomes less stable until the beam is not usable for guidance. This beam deviation is sensed by the over-station sensor @ and, when the deviation rate reaches 8 mV/see, the radio signal is sm”tched off. The system then follows the selected course signal. The beam rate sensor will continue to monitor and, 40 seconds after the beam again becomes usable, will switch on again and the system will fly the radio bean, outbound fran the station. This mode is the primary cross-country navigation method. 4.

B.

(3)

(c)

VOR Approach (VOR APR) Mode (See figure 211.) This mode is most useful in flying close to VOR facilities and using them as ground fixes to let down to an airfield which may or may not have 11S facilities. The system and procedures are identical to those of the VOR mode except that the gains are set to accommodate the proximity to the station with strong but erratic bean characteristics. The VOR APR mode sm”tch is used to select the mode.

(d)

ILS Localizer Mode (See figure 211.) The.ILS localizer mode is used for guidance to an airport equipped with standard localizer transmitters. The procedure to intercept and track the beam is similar to the VOR mode. The navigation radio is tuned to the desired station’s frequency and the inbound runway heading is set on the HS I course select knob. Maneuver the aircraft by means of the heading mode to intercept the localizer beam on the front course 10 to 12 miles out. Arm the 11S mode by pressing the ILS switch on the Mode Selector. The NAV “ARM” and the ILS “ARM” annunciators W-ll illuminate and the heading mode continues to operate. The lateral beam sensor 6 monitors 9 to 2.0 bean deviation and W-ll trip at approximately 1. degrees (1.5 to 2 dots) depending on flight conditions and close the NAV switch, thereby activating the beam tracking function. The NAV annunciator switches to “CAP” and the heading mode turns off. The canputer commands a 20-degree (maximum) bank and turn towad the runway. When the aircraft

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aligns m-thin 15 degrees of the inbound course and rolls out to within 6 degrees of level, logic switches to “On Course” which reduces bank limits and starts crossw”nd correction 8 . As miles the aircraft nears the outer-marker, typically about out, the center of the glide slope beam is reached and a letdown is initiated. When the aircraft reaches 1200 feet as sensed by the radio altimeter, localizer gain programming is initiated in the Ft)Cwhich progressively reduces radio gains to compensate for bean convergence. If the radio altimeter is invalid, programming starts at glide slope capture and is controlled by the vertical speed signal.

Q

Normal system operation continues until runway threshold or decision height is reached. Roll attitude, m“th a 33-second washout, and course error are added at various points in the signal path to provide damping and assure precise path tracking. 4.

B.

(3)

(e) Back Course (BC) Mode (See figure 211.) The back course (or reverse localizer) mode is used in making approaches to facilities with limited runways or ILS equipment. The mode is flown in exactly the same manner as a normal localizer approach except that the intercept is made from the opposite direction. Front (inbound) course must be set on the HS1. Reversing switches in the Flight Director Computer changes the polarity of roll attitude, course error, and radio deviation. Proper HSI display and steering to bem center is accomplished w“th no additional mental effort required.

(4)

Longitudinal (Pitch) Axis Modes (a)

Glide Slope (ILS) Mode (See figure 212.) In ILS mode, the FDC computes guidance to follow the glide slope radio bean and displays the information on the ADI pitch command bar. The pilot (or Helipilot), by foll~”ng the commands, will guide the aircraft along the desired path. Initial conditions for using the mode are: (1) NAV radio tuned to 11S frequency, (2) Mode Selector engaged in 11S, and (3) localizer captured. With the aircraft stabilized on the localizer bean, flying level and approaching the glide S1ope bean with ILS “ARM”, the system is ready for glide slope capture. A vertical bean sensor @ monitors bean deviation and will trip at 20 to 30 mV Men the sensor operates, logic circuits (1/3 dot on HSI). close the GS switch pmnitting radio data to enter the FDC canputation. St ultaneously, a nose-down signal is applied momentarily at 65 which causes the aircraft to pitch over about 2 degrees to approximately align w“th bean center. The

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control signal is routed to ~ and to the AOI pitch ccmmand bar. As the pilot follows the pitch command, pitch attitude (of opposite polarity) suns at 1 and, when the signals balance, the camnand bar shows cQ ter and the aircraft now flies the desired attitude down the glide path. Any residual glide slope error signal canmands an attitude change until the beam is centered.

When the pitch-over occurs at bean capture, the helicopter tends to gain speed. The collective control is reset downward by the pilot to maintain airspeed. The collective position potentiometer senses the co trol position and feeds a signal through a washout circuit b6 permitting a short perio of and 61 with signal effectivity. The signal is summed at @ radio beam and attitude. The result is an attitude with the nose slightly higher to keep the speed from increasing Mile still tracking the radio bean. As the helicopter gets nearer the transmitter, the radio beam converges which requires system gain changes to maintain optimum tracking accuracy. Again programmer @) adjusts the radio signal proportional to radio altitude which is also proportional to beam width. The programming is started at an altitude of 1200 feet and reduces the signal to 25 percent of full value as the altitude changes from 1200 to 200 feet. If the radio altimeter becanes invalid, the altitude sensor drives the programmer and starts working at glide slope capture. Note that the baranetric altitude signal is mixed with an attitude term at @ to form an altitude rate or vertical speed signal. In the event of a gyro or FDC failure, a itch bar bias (PBB) fixed signal is supplied to @ then to t Ie ADI bar which is driven out of view, thus preventing the pilot from attempting to follow an erroneous signal. An invalid glide slope w“ll prevent (3Scapture or cause bar bias after capture and the Helipilot to revert to attitude retention. 4.

B.

(4)

(b) Auto-Level Mode (See figure 212.) Auto-level mode comes into play below an altitude of 150 feet after an 11S descent if the radio altimeter is functional. The flare sensor monitors radio altitude and barometric altitude rate. It trips at approximately 90 feet and causes a 2-degree, nose-up, momentary signal to be directed to the system at @ which gives an ADI nose-up command. As the ship changes attitude, the collective and attitude signals are washed out and the system follows the radio altitude signal to fly level at 50 feet.

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(c) Altitude Hold (ALT) Mode (See figure 213.) The altitude hold mode measures altitude static pressure and controls the aircraft to maintain the altitude at the moment of engagement. ~ An altitude signal from the altitude sensor is routed to @ , summed with an attitude signal, and differentiated to form a pseudo-altitude rate or instantaneous vertical velocity @. At altitude engagement, the signal level is synchronized to zero by a digital counter circuit and serves as a reference for altitude hold. Any variation in altitude then generates an error signal which is sunmd with the IVY signal and routed thru summation point @ to summation point @ . At this point, attitude signal is added via a 12-second washout circuit, providing short-te?m stabilization while the altitude error is being corrected. The signal then passes to the AOI to drive the pitch command bar. If the Helipilot is coupled, the signal also is routed to the Helipilot Computer for automatic path following. (d)

Vertical Speed Hold (VS) Mode (See figure 213.) The vertical speed hold mode maintains a selected altitude rate or vertical speed. As in the altitude hold mode, an altitude sensor provides a signal which is processed into an instantaneous vertical velocity signal at @ . From there, it passes to a synchronizer circuit at @ . When the VS mode is selected, the signal level is locked and becomes the operating voltage. Any variation in vertical speed then generates an error signal which is routed through sumnation point @ to attitude summation point @ . At sunmation point @ information is added via a 12-second washout circuit, to provide short tetm stabilization while the vertical speed error is being corrected. The signal then passes to the ADI to drive the pitch cunmand bar. Subsequent canmands at the AOI and Helipilot return the vertical speed to the reference level. To improve altitude and vertical sped performance, a collective position signal is fed into the circuits at @ to help the pitch axis anticipate the changing effects due to the pover change. If a change in the reference level is desired, operating the beep switch auses a 200-fpm/sec reference voltage to be applied at 68 . The switch is held until the desired vertical speed is reached.

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412

DIVISION

(e) Airspeed Mode (See figure 213.) In the airspeed mode, the system responds to a signal frcm the airspeed sensor to control the aircraft to fly the s ed at the keeps moment of engagement. The synchronizer circuit at 6 its output at zero until mode engagement. Then, any speed variation results in an error signal at @ and subsequent camnands at the AOI and Helipilot to return the airspeed to the reference level. If a new, slightly different speed is desired, the beep switch is operated to change the reference level at a rate of 3 knots per second. {f)

Go-Around Mode (See figure 213.) The go-around mode is used to transition from a descent into a climb-out situation, when a missed approach has occurred. The pilot selects the mode by pressing the go-around button on the collective control. When go-around has been selected, the GA annunciator on the Mode Selector will illuminate and the pilot must then apply collective power to perform the go-around maneuver. If the aircraft airspeed is less then 55 knots, the flight director will camnand a wings level and zero (0) fpm climb rate. At airspeeds between 55 and 65 knots, the flight director will canmand a linear rate of climb between O and 750 fpn. Airspeeds above 65 knots prod ce a climb command of 750 fpm. The fixed signal applied at d9 causes the itch attitude to change until the vertical speed signal fran d6 is equal. Thus, the go-around mode is really a vertical speed tracking function w“th a fixed reference input signal.

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AvIL)NICS DIVISION I MODE

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—

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A

5.

—mmy

;;:;:WE

bHl SYSTEMS . AVIONICS DIVISION

BELL412

TAI RSYM Compass System Operation (See figure 215.) Basic power for the system is the 115-volt, 400-tiz in ut to the TARSYN directional gyro. This provides internal power for tRe gyro and a 26-volt ac power output for the Flux Valve and CS-412 Remote Compensator. The F1ux Valve senses the horizontal ccanponentof the earth’s magnetic field. Using the 26-volt ac reference signal, it provides an output signal that represents aircraft heading in the earth’s magnetic field. This signal will provide a command to keep the rotor spin axis of the directional gyro aligned to magnetic north in the slaved mode. The CS-412 Remote Compensator compensates the Flux Valve for single-cycle error (hard-iron effects) by inducing small dc currents into the Flux Valve coils to correct for errors caused by ferrous metal in the aircraft. The procedure for adjusting this level of canpensation is discussed in the REMOVAL/REINSTALLATION AND ADJUSlllENTprocedures. In the slaved mode of operation, the TARSYN directional gyro is slaved to a position relative to the magnetic heading reference as supplied by the Flux Valve and the canpensator. Slaving of the directional gyro is accomplished by supplying current flow through precession coils affecting the sensitive axis of the gyro. The MANUAL SYNCHRONIZATION switch is used to engage fast slaving of the directional gyro. When fast slaving Is engaged, the slaving rate is increased from approximately 3.5 degrees per minute to approximately 40 degrees per minute. Once engaged, fast slaving continues until the compass card of the HSI indicates within 4 degrees of actual heading, at which time the normal slaving rate is reestablished. In the free mode of operation, magnetic information frcinthe Flux Valve and compensator is disabled in the gyro and no slaving is performed. The TARSYN directional gyro provides cunpass information as a product of the position of the aircraft with reference to the position of the unslaved gyro. As no slaving is performed, the displayed heading information is subject to error as the result of free gyro drift. During operation of the directional gyro, any of the following conditions will cause a loss of the heading valid signal supplied to the HSI and cause the HDGflag to cane into view: (a)

Low voltage to the TARSYN directional gyro p-r

(b)

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(c)

Fast sync (manual synchronization)

supply

The HSI receives three-wire canpass information fran heading synchro transmitter No. 1 (B3) of the TARSYN directional gyro and uses it to move the HSI compass card to the proper heading.

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Synchronization between the Flux Valve and the actual headin of the aircraft is indicated by the compass synchronization annunciator of t !e gyro control panel. When the+ is in view on the annunciator, the compass card is rotating in the counterclockwise direction (actual heading greater than When the ● is-in view, the compass card is rotating in the indicated). clockwise direction (actual heading less than indicated). When synchronized, the annunciator slowly oscillates between the+ and the ● .

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DIVISION

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Power Distribution Figures 216 and 217 contain power distribution information for the Integrated Flight Control System. Figure 216 contains dc power distribution and figure 217 contains 26- and l-15-voltac power distribution.

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OIVISION

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MAINTENANCE MANUAL

~>~”Em~~ FLIGHT SYSTEMS AvIONICS

DIVISION

BELL

412

SECTION 4 GROUND OPERATIONAL TESTS (GROUND CHECK~

1.

General This section describes procedures for checking the SHZ-412 Integrated Flight Control System for correct installation and proper operation of all canponents in the system. NOTICE Procedures in table 301 are based on Sperry Engineering Bulletin 521OH-16O6, revision A. Should any failure arise while the follow”m wound refer to FAULT ISOLATION as required. - -

2.

check is beinq -. wrformed.

Equipment and Materials The test equipment required to test the system is listed in table 301.

3.

Procedure Instructions for performing the test are listed in table 301.

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TITLE:

1.

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SHZ-412 INTEGRATED FLIGHT CONTROL SYSTEM GROUND TEST PROCEDURES FOR THE BELL 412 HELICOPTER

SCOPE This document contains procedures for checking and aligning the Bel? 412 Integrated Flight Control System for correct installation and proper operation of all components. All tests described herein shall be performed on each aircraft prior to initial flight test or whenever a malfunction is suspected.

2. 2.1

2.2

APPLICABLE DOCUMENTS Bell Installation Drawings 412-075-037

Wiring Diagran, AFCS

412-075-038

Wiring Diagran & Cable Assembly - FLT Director.

Sperry Documents EB7001446

3.

Installation Bulletin for the Bell 412 AFCS/NAVC System.

SYSTEM COMPONENTS The canponents covered in this procedure are 1istealin tables 1 and 2 of this manual. Al1 other equipment instal1ed as part of the IFR package shal1 be covered by basic aircraft test procedures; i.e., navigation and conrnunicationradios.

4. 4.1

REQUIREMENTS System Wiring Prior to initial installation of system components, all aircraft wiring shall be checked for proper continuity in accordance with the applicable dr~” ngs referenced in paragraph 2.

4.2

Power Application The Bell 412 Integrated Flight Control System requires 115 volts 400 Hz, 26 volts 400 Hz, and 28 volts dc power. With the inverters installed and connected, apply external dc power to the helicopter. Engage all circuit breakers individually ensuring that each breaker applies pover to the appropriate system cartponent. .

4.3

Prior to checkout of the AFCS/NAVC verify that the attitude system, canpass system, and navigation radios are fully operational.

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DIVISION

412

AFCS CHECKOUT PROCEDURE Prior to AFCS checkout, ver~fy that the pitch control link contains two 4012373-903 actuators, roll contains two 4012373-913 actuators, and yaw contains one 4012373-905 actuator. Connect external hydraulic and electrical power to the aircraft. Initial Switch Positions: Inverter 1 Inverter 2 HP1 HP2 Force Trim

5.1

-

ON ON OFF OFF OFF

Lamp Test

5.1.1

Press and hold the AFCS lamp test pushbutton (SBY on the Mode Selector if a flight director is installed). All lamps on the Helipilot Controller (HP1 - ON, HP2 - ON, CPL - ON, SAS, and ATT) shall be on.

5.1.2

Turn the PED LT switch on. dimly illuminated.

All lamps on the Helipilot Controller shall be

Release the lamp test pushbutton. Turn the PED LT switch off. 5.2

Force Trim Check

5.2.1

Move the cyclic stick and pedals. minimun friction on the controls.

5.2.2

Turn the Force Trim switch on. Both the cyclic stick and pedals shal1 have spring force when moved fran their detent position. “

5.3

There shall be no spring force and

Helipilot Engage Tests

5.3.1

Momentarily press the HPl pushbutton on the Helipilot Controller. The HP1 ON and ATT legends will illuminate and remain engaged if the gyro ATTflag in the pilot’s ADI has retracted.

5.3.2

Press the SAS/ATT pushbutton. legend shall illuminate.

The ATT legend W-ll extinguish and the SAS

5.3.3

Press the HP2 pushbutton. l%e HP2 ON and ATT legend W-ll illuminate if the ATT flag in the copilot’s ADI has retracted. The SAS legend shall extinguish.-

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5.3.4

Pull the PILOT ATTD 115 V ac circuit breaker. The HP1 ON legend shall extinguish; the HP2 ON and ATT legends shall renain on. The AFCS caution light shall illuminate.

5.3.5

Reset the PILOT ATTO 115 V ac breaker. Press the SAS/ATT switch to engage the SAS mode. Reengage the HP1. HP1 ON and ATT legends shall illuminate while the SAS legend extinguishes.

5.3.6

Pull the CPLTATTD 115 V ac circuit breaker. The HP2 ON legend shall extinguish; the HP1 ON and ATT legends shall remain on. The AFCS caution light shall illuminate. Reset the CPLTAITD 115 V ac breaker. ,. ATTGain/Gain Doubling Tests

5.4

5.4.1

Tilt the pilot’s VG to simulate a 5-degree pitch up attitude. API shall indicate maximum downward deflection.

5.4.2

Press the HP2 pushbutton. The HP2 legend shall illuminate and the pitch API shall reduce to a 50-percent downward deflection. Level the gyro.

5.4.3

Disengage simulate

HP2.

!bmentarily

5-degree left roll maximun right deflection. a

press the attitude.

The pitch

Tilt the pilotgs VGto The roll API shall indicate

FTR.

5.4*4

Enga e HP2. The roll API shalt reduce to a 50-percent right deflection. Leve? the gyro.

5.4*5

left Disengage HP1. Tilt the copilot’s VG to simulate a 5-degree attitude. The roll API shall indicate maximum right deflection.

5.4.6

Hold the SYS 2 switch and engage HPl.- The SYS 2 roll API shall reduce to a 50-percent right deflection. Release the SYS 2 sm”tch.

5.4.7

Tilt the copilot’s VGto simulate a 5-degree pitch up attitude. Press and hold the SYS 2 sw”tch. The SYS 2 pitch API shall indicate 50-percent downward deflection. Disengage HP1. The SYS 2 API shall indicate maximum downward deflection. Release the SYS 2 sm”tch and level the gyro. Manentarily press FTR.

5.5

All

Beep/Trim

Computer

rol1

Tests (Auto Trim)

5.5.1

Hold the pilot’s four-way beep switch in the forward position for for approximately 5 seconds and note approximately 5 seconds. Beep left that the pitch and roll API remain centered.

5.5.2

Disengage HP2 and engage HPI.

Repeat step 5.5.1.

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5.5.3

Position the cyclic stick to the full aft position. Enga e HP2. Beep forward on the cyclic beep switch to obtain one-half need7ewidth downward deflection on the pitch-API. Observe that the pitch API deflection moves downward. Men approximately a 25-percent deflection is reached, the cyclic stick shall move forward and the API shall return toward center.

5.5.4

Move the cyclic stick 5 cm aft (out of detent); the auto trimming shall stop.

5*5.5

Allow the stick to return to the detent position. that the trimming stops.

5.5.6

Reengage HP1 and momentarily press FTR. Position the cyclic stick to the full aft position. Beep forward on the cyclic beep switch for 3 seconds. The time for the cyclic to travel fran the full aft position to the full forward position shall be 60 A 15 seconds.

5.5.7

Momentarily press

Disengage HP1 and note

FTR.

Beep the API slowly

cyclic aft approximately one-half needlewidth up. Observe the drifts further up and that, at approximately a 25-~rcent deflection, the cyclic drives aft and the API returns toward center. Disengage tJP2and note that the cyclic stops trimming.

5.5.8

Engage HP2. forward

Momentarily press FTR and pl ace the cyclic

to

the

ful 1

position.

Beep aft on the cyclic beep switch for3 seconds. The time for the c clic to travel frun the full forward position to the full aft position shaz 1 be 60 ~ 15 seconds. 5.5.9

Momentarily press the FTR button. Beep the cyclic right to obtain one-half needlewidth deflection on the’roll /@I. Observe that the API slowlv drifts further riaht. Men a 25-r)ercentdeflection is obtained. the C-X1ic shall drive r~ght and the API’shall return toward center.

5.5.10

Move the cyclic stick left 5 cm (out of detent); the trimming motion shall stop;

5.5.11

Allow the stick to return to the detent position. Disengage HP1 and note that the cyclic stops trimming.

5.5.12

Reengage HPI. left position.

Momentarily

press

~R

and

position

the cyclic to the full

The time for the c lic for 3 seconds. beep sw”tch right from the full left position to the full right position shaY 1 be 60 ~ 15 seconds.

Beep the to travel

cyclic

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5.5.13

Momentarilypress

DIVISION

BELL

412

FTR.

Beep the cyclic left to obtain a one-half needlewidth deflection. Observe that the API slowly drifts further left. When a 25-percent deflection is obtained, the cyclic shall drive left and the API shall return to near center. 5.5.14

Disengage HP2 and note that the cyclic stops trimming.

5.5.15

Reengage HP2. Momentarily -. press FTR and position the cyclic to the full right @sition. Beep the cyclic beep switch left for 3 seconds. The time for the cyclic to travel from the full right position to the full left position shall be 60 ~ 15 seconds.

5.6

SAS Mode Tests Initial Switch HP1

Positions:

~:/ATT

- ON - OFF - SAS

Force

-

Trim

ON

5.6.1

Momentarily press the FTR switch. Tilt the pilot’s VGto simulate a 5-degree pitch up attitude. The pitch API shall initially deflect downward and then slowly return to center in approximately 30 seconds. Level the gyro.

5.6.2

Momentarily press

FTR.

Tilt the pilot’s VGto API shall deflect 30 seconds.

5.6.3

Return Move the

5.6.4

the

pilot’s

left

simulate a 5-degree roll right attitude. The roll slowly return to center in approximately then

and

VG to

level

cyclic 1 inch right.

The

roll

Fbmentarily API

shall

press

deflect

FTR. maximum

right.

Momentarily press FTR. Pull the cyclic 1 inch aft.

5.6.5

position.

The pitch API shall deflect fully upward.

Disengage HP1 and engage HP2 in the SAS mode. Tilt the copilot’s VG to simulate a 5-degree nose-up attitude. The pitch API shall deflect downwati and then slowly return to center in approximately30 seconds.

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5.6.6

Momentarily press Tilt roll

the API

DIVISION

BELL

412

“

FTR.

copilot’s VGto simulate a 5-degree shall deflect left and then slowly

roll return

right attitude. in to center

The

approximately 30 seconds. 5.6.7

5.6.8

Return the copilot’s VG to level position.

Fbmentarily press FTR.

Move the cyclic 1 inch right.

shall

Momentarily

The roll

deflect

maximum

right.

press FTR.

Push the cyclic 1 inch forward. downward. 5.7

API

The pitch API shal1 deflect fully

Yaw Axis Tests

5.7.1

Engage HP1. Rotate the rate gyro to simulate a right yaw condition. While the adjustment is made, yaw API shall deflect left.

5.7.2

Connect an air data test set to the copilot’s pitot. Simulate a 100-knot airspeed. Tilt the pilot’s VG to simulate a 20-degree rol1 right attitude. The yaw API shall initialIy deflect right and then S1owly return to center.

5.7.3

Apply 1 inch right pedal and observe that the yaw API is not affected by this motion.

5.7.4

Reduce the simulated airspeed to40 knots. Apply 1 inch right pedal. The yaw API shall deflect maximun right. Return the pedals to neutral. The API shall deflect to slightly left of center and then slowly return to neutral.

5.7.5

Tilt the pilot’s VG to simulate a 5-degree rol1 right and rol1 left attitude and observe that this action has no affect upon the yaw API.

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412

FLIGHT DIRECTOR (NAVC) CHECKOUT PROCEDURE

Before testing the Flight Director, ensure that the navigation radios .and “ gyros are operational. Apply po~r in accordance w“th paragraph 4.2. The SBY annunciator on the Mode Selector shall light upon power application. 6.1

Logic Tests

6.1.1

Prior to receiving gyro valids, engage HDG and VS modes on the Mode Selector. The HDG and VS annunciator shall light and the ADI pitch and roll command bars shall be out of view.

6.1.2

After 3 minutes following povm indicate the following:

6.1.3

application, the ADI/HSI flags shall

HSI

-

HDG flag out of view

ADr

-

ATT and FD flags out of view and pitch and rol1 cunmand bars shall come into view.

Press and hold Mode Selector SBY pushbutton. All lights on the Mode Selector and Helipilot Controller shal1 1ight. The FD flag on the ADI shall also be in view. Release the Mode Selector SBY pushbutton.

6.1.4

Turn the pilot’s Instrument Light M“tch on. Press and hold the remote S8Y pushbutton on the pilot’s cyclic stick. All lights on the Mode Selector shall be lit dimly.

6.1.5

Turn the Dilot’s Instrument Liqht sw”tch off. Press and hold the remote SBY switch on the copilot’s cfilic. All lights on the Mode Selector and Helipilot Controller shal1 be lit.

6.2

Heading Mode and ATT Feedback Tests

6.2.1

Set the heading bug on the HSI to the aircraft heading. Engage the HDG mode on the Mode Selector. The HDG annunciatoron the Mode Selector shall light and the roll camnand bar on the ADI shall center.

6.2.2

Set the heading bug on the HSI to 10 degrees right of the aircraft heading. Tilt the pilot’s VGto simulate a 10 f l-degree right roll attitude to center the ADI roll bar.

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6.2.3

Set the heading bug on the HSI to 45 de rees right of the aircraft heading and tilt the fJllOt S VG tO simulate a 28 ~ 2-de pee right roll attitude. The ADI roll ccmmand bar shall be centered. Se? the heading bug on the HSI to 45 degrees left of aircraft heading and tilt the pilot’s VGto simulate a 20 ~ 2-degree left ‘oil attitude= The MI roll c~mand bar shall be centered. Return the VG to level position.

6.2.4

Engage the VS mode on the Mode Selector. The pitch bar on the shall ADI center. Tilt the pilot’s VG to simulate a 5-degree pitch down attitude. The pitch bar shall move up and slowly return to center in 10 seconds. Return the pilot’s VGto

6.3

level position.

VOR Mode Tests

6.3.1

Set the course pointer on the HSI to aircraft heading. Usin a VOR/ILS test set, transmit a VOR radial positioning the aircraft 10 i egrees right of the selected course. Turn on No. 1 and No. 2 navigation radios and tune to the transmitted signal. The HSI and ~S deviation bars shall displace full-scale left and the TO-FROl pointers shall display TO for each radio. In addition, check the following: HSI - NAV flag, Glide Slope pointer and flag are retracted

6.3.2

Engage the NAV mode on the Mode Selector. The NAVARM and HDG annunciators shall light and the ADI roll ccmnand bar shall still respond to the heading bug.

6.3.3

Rotate the cwrse pointer on the HSI 10 degrees left. The course deviation bar shal1 be centered, the Mode Selactor HOG and NAV ARM annunciator shall be out, and the NAV CAP legend shall be lit. The ADI roll command bar shall be displaced left.

6.3.4

After-approximately 3 minutes, the roll canmand bar shall slowly return toward center. Press SBY and align the course pointer to the aircraft heading.

6.4

0SS, VOR Mode Tests

6.4.1

~:nfl~~

a VOR radial corresponding to the aircraft heading. Engage the NAV CAP shall be lit and the roll camnand bar shall be centered:

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6.4.2

Simultaneously start timer and rapidly change (step function if possible) the transmitted VOR signal positioning the aircraft 10 degrees right of the selected course.. After approximately40 seconds, the roll bar shall deflect left.

6.4.3

Press SOY. Repeat test 6.4.1. Rapidly change the transmitted VOR signal positioning the aircraft 10 degrees right of the selected course. Rotate the course pointer 5 degrees clockm”se, the roll bar shall move to the right. Realign the course pointer to the aircraft heading.

6.5

VOR APR Mode Tests

6.5.1

. Press SBY. Engage the VOR APR mode on the Mode Selector. The VOR APR ARM and HDG annunciators shall light and the ADI roll ccmmand bar shall still respond to the heading bug.

6.5.2

Rotate the course pointer on the HSI 10 degrees ccw. The course deviation bar shall be centered, the Mode Selector VOR APR ARM and HOG annunciator shall be out, and the VOR APR CAP legend shall be lit.

6.5.3

The roll cunmand bar shall slowly return toward center in approximately20 seconds.

6.6

0SS, VOR APR Mode Tests Press SBY.

6.6.1

Realign the course pointer to the aircraft heading. Transmit a VOR radial corresponding to the aircraft heading. Engage the VOR APR mode. VOR APR CAP shall be lit and the roll cunmand bar shall be centered.

6.6.2

Rapidly change the transmitted VOR signal positioning the aircraft 6 degrees right of the selected course. The roll ccmand bar shall initially deflect left, return to center for approximately 4 seconds, then deflect 1eft.

6.7

ILS Mode Tests

6.7.1

Engage the SBY mode on the Mode Selector. Set the VOR/ILS test set to a localizer frequency and adjust deviation for full-scale left (greater than two dots) deflection (aircraft right of desired course). Pull the 28 VDC RADIO ALT circuit breaker (if applicable). radios to localizer frequency.

Tune NAV

Engage the ILSmode on the Mode Selector. The HOG, NAV ARM, and ILS ARM annunciators shall be illuminated. Observe that the Glide Slope flag and pointer are in view and theADI roll bar responds to the heading bug.

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6.7.2

Reduce deviation on VOR/ILS test set to sli tly less than 2 de rees two s rall be out and the %1 dots). The Mode Selector HDG annunciator roll camnand bar shall be displaced left.

6.7.3

Rotate the HSI course knob cw 30 degrees. The AOI roll command bar shall move towar% center. Engage SBY mode on the Mode Selector.

6.7.4

Rotate the HSI course knob until course pointer indicates 180 degrees fron aircraft heading. Engage the BCmode on the Mode Selector. The ADI roll canmand bar shall displace right.

6.7.5

Rotate the HSI course knob ccw 30 degrees. move toward center.

6.7.6

Vary deviation on VOR/ILS test set. in the same direction..

6.7.7

Rotate the HSI course knob to align the course pointer m”th the aircraft heading.

. 6.7.8

The ADI roll canmand bar shall

The ADI and tlSIdeviation shall track

Set the VOR/ILStest set localizer deviation to zero and glide slope deviation to maximun up (aircraft belw beam). The glide S1ope pointer on the HSI shal1 indicate full-seale up and the GS flag on the HSI shal1 be out of view. Engage

the

ILS and ALT modes.

The NAV CAP, 11S ARM, and ALT annunciators shal1 light and the AD I pitch and roll canmand bars shall be centered. 6.7.9

Reduce the lide slope deviation on VOR/ILS test set to zero. The Mode Selector ALf and ILS ARM shal1 go out and the ILS CAP annunciator shal1 light. The ADI pitch cumnand bar shall move down slightly at GS capture and slowly return to center.

6.7.10

Set the VOR/ILS test set glide slope deviation to maximun down (aircraft above beam). The ADI pitch camnand bar shall move down.

6.7.11

LOC Progranner/Flare Command (only applicable if a Radio Altimeter is installed). Set the VOR/ILS test set glide slope deviation for zero deviation.

6.7.12

Simultaneously engage the 28 VDC RADIO ALT circuit breaker and start the timer, After 1 minute. the Radio Altimeter Indicator shall read zero and the Indicator OFF flag-shall be retracted. The pitch bar shall remain centered.

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DIVISION

6.7.13

After approximately 2 climb cormnand.

6.7.14

Pull the 28 VDC RADIO ALT circuit breaker; the pitch bar shall bias from view.

6.8

minutes,

the

pitch

bar

shall

move

up indicating a

Air Data Mode Tests

6.8.1

Engage GA on the Mode Selector. extinguish

and

the

GA annunciator

The NAV CAP and ILS GS annunciators shall The pitch bar on the shall illuminate.

ADI shall remain centered. 6.8.2

Connect an air data test set to theAFCS/NAVC pitot static system and simulate a 750-fpm rate of climb at an airspeed of 100 knots. The pitch bar shall center.

6.8.3

Engage the VS mode on the Mode Selector maintaining climb. The pitch bar shall remain centered.

6.8.4

SirrulateO fpm, the pitch bar shall move up.

6.8.5

Press and hold the pilot’s four-way beep switch to the forward position. The pitch bar shall move toward center.

6.8.6

Press and hold the pilot’s four-way beep switch to the aft position. The pitch bar shall move up.

6.8.7

Repeat the above tests (6.8.5 and 6.8.6) using the copilot’s beep switch.

6.8.8

Engage the ALT mode on the Mode Selector. With O fpm input to the pitot-static system, the pitch bar shall be centered.

6.8.9

Decrease the simulated altitude by 50 feet. The pitch bar shall move up.

6.8.10

Sinulate a 100-knot airspeed and engage AS on the Mode Selector. The AS annunciator shall light and the pitch bar on the ADI shall be centered.

6.8.11

Decrease the sinulated airspeed to 90 knots, the pitch bar shall move down.

6.8.12

Press and hold the pilot’s four-way beep switch in the aft position. The pitch bar shall move towards center.

6.8.13

Hold the pilot’s beep switch forward. The pitch bar shall move down.

a 750-fpm

rate of

Repeat the above tests (6.8.2 and 6.8.13) using the copilot’s beep switch. Remove the air data test set.

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COUPLING TESTS Initial Switch Positions: Inverter 1 Inverter 2 Force Trim

-

-

ON ON ON

7.1

Engage both HP1 and HP2 in the ATT mode. Engage HDG on the Mode Selector. Observe that the CPL light on the Helipilot Controller illuminates.

7.2

Engage the VS mode on the Mode Selector. Turn the heading select knob on the HSI to 10 degrees right of aircraft heading. 8eep the cyclic stick trim button aft. The pitch API shall deflect up and the roll API shall deflect right. Press the SYS 2 switch. The SYS 2 API’s shal1 track the SYS 1 API’s.

7.3

Press the CPL pushbutton on the Helipilot Controller. The CPL legend shall extinguish and the pitch and roll API shall move to center.

7.4

Press the CPL pushbutton again. and roll API’s shall deflect.

7.5

Press the HPl pushbutton on the Helipilot Controller. The HP1 legend shall extinguish, the CPL legend shall extinguish, the DCPL legend on the instrument panel shall light and the pitch and roll API’s shall move to center. Press the SYS 2 sm”tch. SYS 2 API’s shall be centered.

7.6

Press the HPl pushbutton again. The HPl and CPL legends shal1 light. The DCPL legend shall extinguish. The pitch and roll API’s shall deflect.

7.7

Press the HP2 pushbutton on the Helipilot Controller. The HP2 legend shall extinguish, the CPL 1egend shall extinguish, and the DCPL legend shal1 light. The pitch and roll API’s shall nmve to center. Press the SYS 2 SWi tch. SYS 2 API’s shall be centered.

7.8

Press the HP2 pushbutton again. The HP2 and CPL le end shall li ht. The DCPL legend shall extinguish. The pitch and roll #1’s shal1 de% ect.

7.9

Align the heading select bug with the aircraft heading. En age both the HOG and the VS modes. Move the heading bug 10-degrees right. ?he roll API shall deflect right.

7.10

Tilt the Dilot’s VG to simulate a 10-de9ree rol1 right attitude. The rot1 API shall-center. Return the VG and t% heading bug to zero. (It may be necessary to initially move the VG more than 10 degrees in order to canpensate for integral control.)

The CPL legend shall light and the pitch

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7.11

Hold the cyclic beep switch aft for approximately 2 seconds.

7.12

Quickly tilt the pilot’s VG to simulate a 5-degrees pitch up attitude. API shall initially center, but then slowly deflect to 60 percent pitch down.

7.13

Engage the ALT mode. Quicklytiltthe ild‘s VG to siml ate a 5-degree pitch down attitude. The pitch API shaf1 initially deflect up, but then slowly return toward center.

7.14

Press SBY. Take care to align the heading bug exactly to aircraft heading and the pilot’s VG to zero roll attitude.

7.15

Engage the HDG mode and observe the rol1 API. Rotate the heading select bug until the roll API deflects just S1ightly right. The rol1 API shal1 The SYS continue to drift right slowly. Quickly press the SYS 2 switch. API shall track the SYS 1 API.

lhe

2

7.16

Disengage HDG and engage the VS mode. Beep the cyclic trim switch aft until the pitch API deflects slightly upward. The API shall continue to move up slowly. Press the SYS 2 switch. The SYS 2 API shal1 track the SYS 1 API.

7.17

Disengage the VS mode. Align the heading bug to the aircraft’s heading. Engage HDG and VS on the mode selector.

7.18

Move the cyclic stick forward and right. The roll API shal1 deflect left and the pitch API shall deflect down. Press the SYS 2 switch. SYS 2 API’s shall track SYS 1 API’s.

7.19

Press S8Y.

Press FTR.

Disengage HP1 and HP2.

Ground Check Table 301 (cent)

22-15-00 Page 315/316 May 1/81

MAINTENANCE MANUAL -