Avionic Gulfstream IV-1

Honeywell

CommercialFlightSystemsGroupBusinessand Commuter AviationSystemsDivisionHoneywellInc.BOX 29000Phoenix,Arizona85038

SPZ-8000 Digital Automatic FlightControl System

Gulfstream IV

SystemMaintenance Manual

Volume I — System and Component Description,and System Operation

22-14-00TITLE PAGE T-1

PRINTED IN U.S./L PUB. NO, Al 5-1146-38REVISED 15 APRIL 1993

1 JUNE 1987

PROPRIETARY NOTICE

This document and the information disclosed herein are proprietary data of Honeywell Inc. Neitherthis document nor the information contained herein shall be used, reproduced, or disclosedto others without the written authorization of Honeywell Inc., except to the extent required forinstallation or maintenance of recipient’s equipment.

NOTICE - FREEDOM OF INFORMATION ACT (5 USC 552) ANDDISCLOSURE OF CONFIDENTIAL INFORMATION GENERALLY (18 USC 1905)

This document is being furnished in confidence by Honeywell Inc. The information disclosed hereinfalls within exemption (b) (4) of 5 USC 552 and the prohibitions of 18 USC 1905.

S93

LASEREF and PR!MIJS are registered trademarks of Honeywe/f Inc.COLORCAL, COLORARAR, and LASERTRAK are additional trademarks of Honeywell Inc.

CopyrigtM 1993 Honeywell Inc.All Rights Reserved

REVISED 15 APRIL 19931 JUNE 1987

Honeywell’s Continuous Quality

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RECORD OF REVISIONS - VOLUME I

For each revision, put the revised pages in your manual and discard thesuperseded pages. Write the revision number and date, date put in manual, andthe incorporator’s initials in the applicable columns on the Record of Revisions.The initials HI show Honeywell Inc. is the incorporator.

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LIST OF EFFECTIVE PAGES

Original . . 0 . . Jun 1/87Revision . . 1 . . Feb 1/88Revision . . 2 . . Mar 1/89Revision . . 3 . . Ott 1/89

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IntroductionINTRO-1INTRO-2INTRO-3INTRO-4INTRO-5INTRO-6INTRO-7INTRO-8INTRO-9INTRO-10INTRo-11/INTRo-12

System Description

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System Description (cent)18.318.418.5/18.6192021222324252627/28

F 29/30

Component Description101102103104105106107108109110111/112

F 113/114115116117118119/120

F 121/122123124125126127128128.1128.2128.3128.4128.5128.6128.7128.8128.9128.10128.11

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128.12128.13/128.14129130131132133134135/136

F 137/138139140141142143144145146147148149150151/152153/154155156157158159160161162163164165166166/1/166.2

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Component Description (cent)183184185186187188189190191192193194195196196.1196.2196.3196.4196.5/196.6197198198.1198.2198.3/198.4198.4.1/198.4.2

F 198.5/198.6F 198.7/198.8198.9198.10198.11198.12198.13198.14198.15198.16198.17/198.18

F 198.19/198.20198.21198.22198.23198.24198.25198.26198.27198.28198.29198.30198.31/198.32

F 198.33/198.34198.35

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198.36198.37198.38198.39198.40198.41198.42198.43198.44198.45198.46198.47198.48198.49198.50198.51198.52198.53/198.54

F 198.55/198.56198.57198.58198.59/198.60

F 198.61/198.62198.63198.64198.65/198.66

F 198.67/198.68F 198.69/198.70198.71198.72198.73198.74

F 198.75/198.76198.77198.78198.79198.80

F 198.81/198.82198.83198.84198.85198.86198.87198.88198.89198.90198.91198.92198.93198.94198.95

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Component Description (cent)198.96 -198.97/198.98

F 198.99/198.100198.101198.102198.103198.104198.105198.106198.107198.108198.109198.110198.111198.112198.113198.114198.115198.116198.117198.118198.119198.120198.121198.122

F 198.123/198.124F 198.125/198.126198.127198.128198.129198.130198.131198.132198.133/198.134

F 198.135/198.136198.137198.138198.139198.140198.141198.142198.143198.144198.145198.146198.147198.148

F 198.149/198.150198.151198.152

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198.153198.154198.155198.156198.157198.158198.159198.160198.161198.162198.163198.164198.165198.166198.167198.168198.169198.170198.171198.172198.173198.174198.175198.176198.177198.178198.179198.180198.181198.182198.183198.184198.185198.186198.187198.188198.189198.190198.191198.192198.193198.194198.195198.196198.197198.198198.199198.200198.201198.202198.203

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Component Description (cent)198.204198.205198.206198.207198.208198.209/198.210

System Operation201202203204205206207208209210211212213214215216217218219220221222223224225226227228229230231232233234235236237238239240241242

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F 267/268269270271272273274275276277278279280281282283284285286287288289290291292293294

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Honeywell

System Operation (cent)295296297298298.1298.2298.3298.4298.5298.6298.7298.8298.9298.10298.11298.12298.13298.14298.15298.16298.17298.18298.19298.20298.21298.22298.23298.24298.25298.26298.27298.28298.29298.30298.31298.32298.33298.34298.35298.36298.36.1298.36.2298.36.3298.36.4298.36.5298.36.6298.36.7298.36.8298.36.0298.36.10

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298.36.11 ■

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298.36.16298.36.17/298.36.18 :298.37298.38298.39298.40298.41298.42298.43298.44298.45298.46298.47298.48298.49/298.50

F 298.51/298.52F 298.53/298.54F 298.55/298.56F 298.57/298.58298.59298.60298.61298.62298.63298.64298.65298.66298.67298.68298.69298.70298.71298.72298.73298.74298.75/298.76

F 298.77/298.78F 298.79/2-98.80298.81298.82298.83298.84298.85298.86298.87298.88

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System Operation (cent)298.89298.90 ■

298.91 ■

298.92298.93298.94298.95298.96298.97298.98298.99298.100

F 298.101/298.102F 298.103/298.104 =F 298.105/298.106 m

F 298.107/298.108 =298.109298.110 ■

298.111298.112 9298.112.1/298.112.2 ■

298.113298.114298.115 m

298.116 a298.117 ■

298.118298.119298.120298.121298.122298.123298.124 ■

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298.167298.168298.169298.170298,171298.172298.173298.174298.175 ■

298.176298.176.1/298.176.2 =298.177298.178298.179 ■

298.180298.181298.182 ■

298.183298.184298.185298.186298.187/298.188

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F 298.193/298.194 =298.195298.196298.197/298.198 ■

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298.201298.202 ■

298.203/298.204F 298.205/298.206F 298.207/298.208 ■

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System Operation (cent)298.214298.214.1298.214.2298.214.3298.214.4298.214.5/298.214.6298.215298.216298.217298.218298.219298.220298.221298.222298.223/198.224

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TABLE OF CONTENTS

Section

1

Subheading ~

VOLUME I

Svstem Descri~tion

General;: System Description

A. LASEREF@ II Inertial Reference System (IRS)B. ADZ-81O Air Data SystemC. AA-300 Radio Altimeter SystemD. EDZ-884 Electronic Display System (EDS)E. DFZ-820 Dual Flight Guidance System

PRIMUS@ 800 Weather Radar System;: PRIMUS@ 870 Weather Radar SystemH. FMZ-800 Flight Management SystemI. Engine Pressure Ratio SystemJ. VLF/Omega System (Optional)[. LSZ-850 Lightning Sensor System (Optional). TCZ-91O Traffic Alert and Collision Avoidance

System (Optional)~. MLZ-850 Microwave Landing System (Optional). Global Positioning System (Optional)

1

10121213141515

1:

18!!18.4

18.518.5

3. Avionics Standard Communications Bus (ASCB) Description 19

Coml)onentDescription

General;: LASEREF@ II Inertial Reference System (IRS)

A. Inertial Reference UnitB. Mode Select Unit

Inertial System Display Unit (ISDU);: Optional LASERTRAKW Navigation Display Unit (NDU)

3. ADZ-81O Air Data System4. AA-300 Radio Altimeter System

A. RT-300 Radio Altimeter Receiver TransmitterB. AT-222 Radio Altimeter Antenna

Paragraph 5 is not applicable to this System.:: EDZ-884 Electronic Display System

A. DU-880 Display UnitB. SG-884 Symbol GeneratorC. DC-884 Display ControllerD. DA-884 Data Acquisition UnitE. DP-884 Dimmer Panel

101

101102

102116124

128.4

130140

140144

145146

146156172

198.10198.22

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Section

2

TABLE OF CONTENTS (cent)

Subheading

VOLUME I

Comoonent Description (cent)

6. F. FC-880 Fault Warning ComputerG. MD-880 Checklist Module

7. DFZ-820 Dual Flight Guidance System

A. FZ-820 Flight Guidance ComputerB. GP-820 Flight Guidance Controller

PC-880 Turn Pitch Controller:: SM-600 Dual Servo/SB-600 Bracket and

TM-260 Dual Trim Servo/TB-261 Bracket

8. PRIMUS@ 800 Weather Radar System

A. WR-800 Receiver TransmitterB. WC-81O Weather Radar Controllerc. WA-800 Antenna Pedestal and FP-900 24-Inch

Flat-Plate Radiator

8.1 PRIMUS@ 870 Weather Radar System

A. WU-870 Antenna and Receiver Transmitter UnitB. WC-874 Weather Radar Controller

9. FMZ-800 Flight Management System (FMS)

A. NZ-920 Navigation ComputerB. CD-81O Control Display UnitC. DL-800 or DL-900 Data LoaderD. PZ-800 Performance ComputerE. SM-81O Servo-Autothrottle

10. Engine Pressure Ratio (EPR) System

11. Optional VLF/Omega System

A. OZ-800 Receiver Processor Unit (RPU)B. AT-800 Antenna Coupler Unit (ACU) -

Teardrop H-FieldC. AT-801 Antenna Coupler Unit (ACU) -

Brick H-FieldD. AT-803 Antenna Coupler Unit (ACU) -

Blade E-Field

12. Optional LSZ-850 Lightning Sensor System

A. LP-850 Lightning Sensor ProcessorLU-860 Lightning Sensor Controller

:: AT-850 AntennaD. AT-855 Antenna

Paqe

198.24198.36

198.38

198.38198.50198.58

198.64

198.72

198.72198.78

198.84

198.88

198.88198.92

198.102

198.102198.128198.138198.142198.152

198.154

198.160

198.160198.166

198.168

198.170

198.172

198.172198.176198.180198.182

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Section

2

Paqe


Subheading

VOLUME I

ComDonent Descrit)tion(Cent)

13. Optional TCZ-91O Traffic Alert and Collision 198.184Avoidance System (TCAS)

A. RT-91O TCAS Computer 198.184B. AT-91O Directional Antenna 198.189c. Typical Bottom OmnidirectionalAntenna (Optional) 198.191

14. Optional MLZ-850 Microwave Landing System (MLS)

A. ML-850 Microwave Landing System ReceiverB. CM-850 Control/DisplayUnit

15. Optional Global Positioning System (GPS)

Global Positioning System Sensor Unit

System Operation

General;: EDZ-884 Electronic Display System

A.

::D.E.

::H.I.J.

K.

Display System FormatsEFIS/EICAS System ComponentsPrimary Flight Display (PFD)Navigation Display (ND) FormatsEngine (ENG) DisplayCrew Alerting System (CAS) DisplaySystem Page DisplaysCompacted EICAS DisplayEFIS/EICAS Reversionary ModesTraffic Alert and Collision Avoidance System(TCAS) DisplaysMicrowave Landing System (MLS)

3. DFZ-820 Dual Flight Guidance System

A. System Performance/OperatingLimits

198.192

198.192198.196

198.204

198.204

201

201202

202202206223261277292

298.22298.27

298.36.1

298.36.12

298.37

298.37B. Fiight Director/AutopilotF~nctional Description 298.43

4. FMZ-800 Flight Management System 298.149

A. General 298.149B. NZ-920 Navigation Computer 298.152c. PZ-800 Performance (Autothrottle)Computer 298.166D. Target Speeds 298.211

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,,,


Section Subheading

VOLUME II

4 Ground Check

1. GeneralEquipment and Materials

:: Procedure

VOLUME 111

5 Fault Isolation

General;: Procedure3. LASEREF@ II Inertial Reference System (IRS)

A. Self-TestB. System Navigation Performance Determination

and Removal CriteriaReject Criteria

k: Techniques to Improve Navigation Performance

4. AZ-81O Air Data System

A. DADC FunctionsB. Altitude Preselect OperationC. Angle of Attack (AOA) OperationD. DADC Red X Failures

5. AA-300 Radio Altimeter System

A. Preflight TestB. In-Flight Test

6. EDZ-884 Electronic Display System (EDS)

A. Trend and Limit MonitoringB. Troubleshooting Display Unit Red “X’’ing

7. DFZ-820 Flight Guidance System

A. List of Flow Chart FiguresB. List of Tables

8. PRIMUS@ 870 Weather Radar System9. FMZ-800 Flight Management System (FMS)

A. Airborne LogicB. Runway Alignment

Estimated Time Enroutek: Descent Time and Fuel Predictions

Paqe

301

301301301

401

401401402

402

402402403

413

413413422422

425

425425

426

426438

439

439439

460461

461461462463

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Section

5


Subheading

VOLUME III

Fault Isolation (cent)

9. E. Stored Flight Plan WaypointsTakeoff Vspeeds

:: V1 SelectionH. Speed/Altitude Entries

Wind/Temperature Model:: Temperature EnvelopeK. Autothrottle Disengages

Takeoff and Landing WeightFl: Level Off at 10,000 Feet for Airspeed Control

in G-IV Phase 11 AircraftN. CDU Blankingo. Fuel Used

Flight Plan Collapse~: EPR Bugs on Approach

Victor Airwayss: Data Loader Fault Codes

10. Engine Pressure Ratio Transmitter

Interconnects

System Schematics

Removal/Reinstallationand Ad.iustment

L3.4.

5.

6.

;:9.

GeneralEquipment and MaterialsProcedure for DU-880 Display UnitProcedure for WC-810/874 Weather Radar Controller,GP-820 Flight Guidance Controller, PC-880 Turn PitchController, DC-884 Display Controller, CD-81O ControlDisplay Unit, DL-800/900 Data Loader, or DP-884 DimmerPanelProcedure for AZ-81O Digital Air Data Computer, FZ-820Flight Guidance Computer, SG-884 Symbol Generator,FC-880 Fault Warning Computer, DA-884 Data AcquisitionUnit, or PZ-800 Performance ComputerProcedure for RT-300 Radio Altimeter ReceiverTransmitterProcedure for AT-222 Radio Altimeter AntennasProcedure for WR-800 Weather Radar Receiver TransmitterProcedure for WA-800 Weather Radar Antenna and FP-90024-Inch Radiator Plate

10. Procedure for WU-870 Antenna and Receiver TransmitterUnit

11. Procedure for SM-600 Dual Servo, TM-260 Dual TrimServo and Brackets, and SM-81O Servo

464464465465465467467467

468468469469469469470

471

501

601

701

701701701

702

703

703704704

706

710

713

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Section Subheading

VOLUME III

8 Removal/Reinstallation and Ad.iustment(cent)

12. Procedure for CM-850 MLS Control/Display Unit13. Procedure for Global Positioning System Sensor Unit

Procedure for AT-91O TCAS Directional Antennai:: Procedure for AT-800/AT-803 Antenna Coupler Unit16. Procedure for AT-850 Antenna17. Procedure for AT-855 Antenna and AT-801 Antenna

Coupler Unit18. Procedure for Updating the Navigation Database

9 Shiminq, Handlinq, and Storaqe

713713714715716716

716

801

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Em!!x!

1

2

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3

4

5

6

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2-2

2-3

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2-6

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6-1

6-2


List of Illustrations

VOLUME I

Title ~

SPZ-8000 System Flow Diagram 5

Component Locations for a Typical Gulfstream IV Installation 7

Lightning Symbols

Example System Using the ASCB

Illustration of a Typical User Subsystem

Example of Bus Activity (Frame O)

SPZ-8000 ASCB Configuration

Inertial Reference Unit

Inertial Reference Unit Block Diagram

Mode Select Unit

Mode Select Unit Schematic Diagram

Inertial System Display Unit

SYS DSPL Switch

ISDU Wiring Diagram

Navigation Display Unit

NDU Wiring Diagram

AZ-81O Digital Air Data Computer

AZ-81O Digital Air Data Computer Block Diagram

RT-300 Radio Altimeter Receiver Transmitter

RT-300 Radio Altimeter Receiver Transmitter Block Diagram

AT-222 Radio Altimeter Antenna

DU-880 Display Unit

Display Unit Cockpit Configuration

18.4

20

22

27

29

102

113

116

121

124

126

128.3

128.4

128.13

130

137

140

143

144

146

148

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Emre

6-3

6-4

6-5

6-6

6-7

6-8

6-9

6-10

6-11

6-12

6-13

6-14

6-15

6-16

6-17

6-18

6-19

6-20

6-21

6-22

6-22.1

6-22.2

6-22.3



VOLUME I

Title

Display Power Panel

EDZ-884 Electronic Display System Interface

DU-880 Display Unit Block Diagram

SG-884 Symbol Generator

SG-884 Symbol Generator Block Diagram

DC-884 Display Controller

Display Controller Declutter Mode

MAP Mode Menu

Comp Mode Menu

Plan Mode Menu

NAV Mode Menu

Preview Mode Submenu

SENSOR Mode Menu

FLT REF Mode Menu

Main TRS Mode Menu

TRS Mode Submenu

SYSTEM Mode Menu

CHECKLIST Mode Submenu

TEST Mode Menu

DISP Mode Menu

NAV Mode Menu with MLS Selected

Preview Mode Submenu with MLS Selected

MAP Mode Menu with TCAS Selected

Paqe

148

151

153

156

167

172

175

175

180

180

183

183

186

186

189

189

192

192

195

195

196.3

196.3

196.5

22-14-00Page TC-8Apr 15/93


-. -. MAINTENANCE

Em!!N

I 6-22.4

6-23

6-24

6-25

6-26

6-27

6-28

6-29

6-30

7-1

7-2

7-3

7-4

7-5

7-6

7-7

7-8

7-9

7-1o

7-11

8-1

8-2

8-3

SYSTEM

DC-884

DA-884

DA-884

DP-884

DP-884

FC-880

FC-880

MD-880

FZ-820

FZ-820

GP-820

GP-820

PC-880

PC-880

SM-600

TM-260

SM-600

TM-260

TB-261

WR-800

WR-800

WC-81O



VOLUME I

Title

Mode Menu with TCAS Selected

Display Controller Block Diagram

Data Acquisition Unit

Data Acquisition Unit Block Diagram

Dimmer Panel

Dimmer Panel Schematic

Fault Warning Computer

Fault Warning Computer

Checklist Module

Block Diagram

Flight Guidance Computer

Flight Guidance Computer Block Diagram

Flight Guidance Controller

Flight Guidance Controller Block Diagram

Turn

Turn

Dual

Dual

Dual

Dual

Pitch Controller

Pitch Controller Schematic

Servo and SB-600 Bracket

Trim Servo and TB-261 Bracket

Servo Schematic

Trim Servo Schematic

Bracket Schematic

Receiver Transmitter

Receiver Transmitter Block Diagram

Weather Radar Controller

I&g

196.5

198.5

198.10

198.19

198.22

198.23

198.24

198.33

198.36

198.38

198.48

198.50

198.55

198.58

198.61

198.64

198.64

198.67

198.69

198.71

198.72

198.75

198.78

22-14-00Page TC19Apr 15/93


Ei9u!2

8-4

8-5

8-6

8-7

8-8

8-9

8-9.1

8-10

8-11

9-1

9-2

9-3

9-4

9-5

9-6

9-7

9-8

9-9

9-1o

9-11

1o-1

10-2



VOLUME I

Title

WC-81O Weather Radar Controller Block Diagram

WA-800 Antenna Pedestal and FP-900 Flat-Plate Radiator

WA-800 Antenna Pedestal Block Diagram

PRIMUS@ 800 MPEL Boundary

WU-870 Antenna and Receiver Transmitter Unit

WU-870 Antenna and Receiver Transmitter Unit BlockDiagram

PRIMUS@ 870 MPEL Boundary

WC-874 Weather Radar Controller

WC-874 Weather Radar Controller Block Diagram

NZ-920 Navigation Computer

NZ-920 Navigation Computer Block Diagram

CD-81O Control Display Unit

CD-81O Control Display Unit Block Diagram

DL-800 Data Loader

DL-900 Data Loader

DL-800/900 Data Loader Block Diagram

PZ-800

PZ-800

SM-81O

SM-81O

Engine

Engine

Performance Computer

Performance Computer

Servo (Autothrottle)

Servo Schematic

Block Diagram

Pressure Ratio Transmitter

Pressure Ratio Transmitter Block Diagram

~

198.81

198.84

198.85

198.86

198.88

198.90

198.91

198.92

198.99

198.102

198.123

198.128

198.135

198.138

198.139

198.141

198.142

198.149

198.152

198.153

198.154

198.159

22-14-00Page TC-10Apr 15/93


Ei9.!Ke

11-1

11-2

11-3

11-4

11-5

12-1

12-2

12-3

12-4

12-5

12-6

13-1

13-2

13-3

13-4

14-1

14-2

14-3

14-4

15-1

15-2



VOLUME I

Title

OZ-800 Receiver Processor Unit

OZ-800 Receiver Processor Unit Block Diagram

AT-800 Antenna Coupler Unit



LP-850 Lightning Sensor Processor

LP-850 Lightning Sensor Processor Block Diagram

LU-860 Lightning Sensor Controller

LU-860 Lightning Sensor Controller Schematic

AT-850 Antenna

AT-855 Antenna

RT-91O TCAS Computer

RT-91O TCAS Computer Block Diagram

AT-91O Directional Antenna

Typical Bottom Omnidirectional Antenna

ML-850 MLS Receiver

ML-850 MLS Receiver Block Diagram

CM-850 MLS Control/Display Unit

CM-850 MLS Control/Display Unit Block Diagram

Global Positioning System Sensor Unit LeadingParticulars

Global Positioning System Sensor Unit Block Diagram

~

198.160

198.164

198.166

198.168

198.170

198.172

198.174

198.176

198.178

198.180

198.182

198.184

198.188

198.189

198.191

198.192

198.195

198.196

198.202

198.204

198.209

22-14-00Page TC-11Apr 15/93


Euu!E

201

202

203

204

205

206

207

208

209

210

211

212

213

213.1

214

215

216

217

218

219

220

221

222



VOLUME I

Title

Cockpit Layout of the EFIS/EICAS Display System

Display Power Panel

Primary Flight Display Format (SENSOR)

Primary Flight Display Format (DISP)

Primary Flight Display Format (FLT REF)

Primary Flight Display Format (AUTO VSPD - CONFIG Mismatch)

Primary Flight Display Format (NAV)

Primary Flight Display Format (VNAV)

PFD Failure Indications (IRS/DADC)

PFD Failure Indications (Mist)

IRS Test Mode Indications

Map Mode Format

Map Mode With Vertical Profile

Map Mode With Weather Radar Display

Vertical Profile Symbols

Map Caution/Warning Displays

IRS Test Mode Display

Compass Mode Display Format (COMP)

Navigation Preview Mode (NAV)

Compass Caution/Warning Displays (IRS)

Compass Caution/Warning Displays (Mist)

Plan Mode Display Format (PLAN)

Plan Mode Caution/Warning Displays

PacJ_e

203

205

210

211

214

215

224

225

229

230

232

236

239

240

243

244

245

248

251

253

254

257

260

22-14-00Page TC-12Apr 15/93


EGluE

223

224

225

226

227

228

229

230

231

232

233

234

235

236

237

238

239

240

241

242

243

244

245



VOLUME I

Title Paqe

Engine Instrument Display Format (SENSOR) 266

Engine Pressure Ratio Transmitter Interface Diagram 267

Engine Instrument Display Format (IRS) 269

Fuel Flow/Hydraulic Pressure Valve Symbology (TRS-MAN) 273

Engine Instrument Display Failure Indications

Crew Alerting System (CAS) Display Format

Master Warning/Caution Panel

CAS Display Failure Indications

Hydraulic System Page

Hydraulic System Page Failure Indications

Fuel System Page Display

Fuel System Page Failure Indications

APU/BLEED System Page Display

APU/BLEED System Page Failure Indications

Engine Start Page Display

Engine Start Page Failure Indications

Engine/APU Exceedances Page

No Exceedances Recorded Format

Exceedance Data Failure Indications

Checklist System Page Display

Checklist System Cursor Control

Waypoint List Display Page

System Page Declutter Mode

276

278

279

291

293

296

297

298.1

298.2

298.5

298.6

298.10

298.12

298.13

298.15

298.16

298.18

298.20

298.21

22-14-00Page TC-13Apr 15/93

Use or disclosure of information on this page is subject to the restrictions onthe title page of this document.

Em!r!2

246

247

248

249

250

251

252

253

253.1

253.2

253.3

253.4

253.5

253.6

253.7

253.8

253.9

254

255

256

257

258



VOLUME I

Title

Compacted EICAS Display Format

Compacted EICAS Failure Mode Indications

Display System Reversionary Panel

Symbol Generator Failure Mode Indication

Pilot’s PFD Reversionary Mode

Copilot’s PFD Reversionary Mode

Engine Display Failure Reversionary Mode

CAS Failure ReversionaryMode

TCAS System Page Display

TCAS Targets on the Navigation Display

TCAS Resolution Advisory on the Primary Flight Display

TCAS Test on the System Page Display

TCAS Resolution Advisory Test on the Primary FlightDisplay

TCAS Extended Test on System Page Display

MLS Displays on the Primary Flight Display

MLS Active Mode Displays on the Navigation Display

MLS Preview Mode Displays on the Navigations Displays

AP, YD, MACH TRIM, and PFD-CMD Select Diagram

~

298.24

298.26

298.28

298.29

298.30

298.31

298.35

298.36

298.36.6

298.36.7

298.36.8

298.36.9

298.36.10

298.36.11

298.36.14

298.36.16

298.36.17

298.51

Flight Director Mode Select Diagram 298.55

Autopilot Engage Logic Diagram 298.77

Power Interruptionof Both Channels 298.83

Power Interruptionof Engaged Channel Only 298.84

22-14-00Page TC-14Apr 15/93


I

I

Li9w!2

259

259.1

260

261

262

263

264

264.1

265

266

267

268

269

270

270.1

270.2

270.3

270.4

270.5

270.6

270.7



VOLUME I

Title ~

Flight Director/Autopilot Roll Channel Mode FlowDiagram 298.101

Dual Couple Approach 298.119

Flight Director/Autopilot Pitch Channel Mode FlowDiagram 298.127

F1ight

Flight

NZ-920

G1obal

Director/Autopilot Yaw Channel Mode Flow Diagram 298.145

Management System (FMS) Architecture 298.151

Navigation Computer Interface Diagram 298.153

DMU (AFIS) and DL-800/900 Data LoaderInterconnects - 298.163

Takeoff Mode Flow Chart 298.176.1

Basic Autothrottle Functions Over the Flight Profile 298.181

Autothrottle Switch Locations on Power Levers

Autothrottle Engage Logic Diagram

Autothrottle Arm, Takeoff, and Hold Mode Select

298.186

298.189

Diagram 298.193

Autothrottle Flight Level Change, Speed (IAS/MACH), andGo-Around Mode Select Diagram 298.199

Autothrottle Mode Flow Diagram 298.205

Climb Phase with No Descents 298.214

Climb Phase with Descent 298.214.1

Transition from Climb to Cruise Phase 298.214.1

Cruise - Climb and Cruise - Descent Subphase 298.214.1

Deswcents with 100 NM of TOD 298.214.2

Descent to Cruise Phase when more than 100 NM from TOD 298.214.3

Descent to Climb Phase 298.214.4

22-14-00Page TC-15Apr 15/93

Useor disclosure of information onthispage issubject to the restrictions on the title page of this document,

II

Im!E

270.8

271

272

273

274

401

402

403

404

405

406

407

408

409

410

411

412

413

414

415

416



VOLUME I

Title

Missed Approach Flight Phase

Example Vertical Flight Plan

Climb Phase

Cruise Phase

Descent Phase

VOLUME 111

IRU Performance Removal Criteria

Cabin Pressure Ratio Output

FAA VMO Function for the Gulfstream IV

CM VMO Function for the Gulfstream IV

SSEC (Low-Speed Range) for the Gulfstream IV

SSEC (High-Speed Range) for the Gulfstream IV

Altitude Alerting Sequence

DADC AOA Block Diagram

Exceedance System Page Format

Diagnosing Symptoms

Both FZ-820S Failing Power-Up(FGC 1 and 2 FAIL Messages on EICAS)

Single FZ-820 Failing Power-Up(FGC 1 or 2 FAIL Annunciated)

Unintended Priority Transfers

AP, YD, or Trim Engagement Inhibited

AP, YD, and Trim Disengagement (All Engaged Functions)

AP or Trim Disengagement (YD is Engageable)

PiJg

298.214.4

298.216

298.218

298.220

298.222

411

416

417

418

419

420

421

423

434

440

442

446

449

451

454

456

22-14-00Page TC-16Apr 15/93


EiwE!

417

418

701

702

703

704

705

706



VOLUME III

Title

Unintended Mode Disengagement

AP, YD, or Trim Control Problems(Oscillations, Kicks, Sluggishness, etc.)

Correct Orientation of AT-222 Antennas

Data to be Loaded Display

Transfer of NAV Database Display

Confirm Transfer of NAV Database Display

Percent Complete of Transfer Display

Completion of Transfer Display

Paqe

458

459

705

717

717

718

718

719

22-14-00Page TC-17Apr 15/93


I

I

Table

1

1.1

1.2

2

3

2-1

2-2

2-3

2-4

2-5

2-6

2-7

2-8

2-9

2-1o

2-11

2-12

2-13

2-14

2-15

2-16

2-17


List of Tables

VOLUME I

Title

System Components

Omega Stations

VLF Stations

ASCB Unit Addresses

ASCB Frame Structure Allowing 40, 20, 10, and 5 HzUpdate Rates

Inertial Reference Unit Leading Particulars

Inertial Reference Unit Input/Output Information

Inertial Reference Unit ASCB Transmitted Data

Inertial Reference Unit ARINC 429 Output Data

Mode Select Unit Leading Particulars

Test Mode Outputs

Inertial System Display Unit Leading Particulars

Inertial System Display Unit ARINC 429 Digital Output Data

Inertial System Display Unit ARINC 429 Digital Input Data

ARINC 429 IRU Discrete Word (Octal Label 270)

Paqe

1

18.2

18.2

21

24

103

104

107

111

116

119

124

128

128

128.1

ARINC 429 Time to NAV Ready Discrete Word (Octal Label 351) 128.2

Navigation Display Unit Leading Particulars 128.4

Navigation Display Unit ARINC 429 Digital Output Data 128.8

ARINC 429 Status Word Discrete (Octal Label 275) 128.9

Navigation Display Unit ARINC 429 Digital Input Data 128.10

ARINC 429 IRU Discrete Word (Octal Label 270) 128.11

ARINC 429 Time to NAV Ready Discrete Word 128.12(Octal Label 351)

22-14-00Page TC-18Apr 15/93


Table

3-1

3-2

3-3

4-1

4-2

6-1

6-2

6-3

I6-3.1

6-4

I 6-4.1

6-5

I6-5.1

6-6

6-7

6-8

6-9

6-10

7-1

7-2

7-3

7-4


List of Tables

VOLUME I

Title

AZ-81O Digital Air Data Computer Leading Particulars

AZ-81O Digital Air Data Computer ASCB Transmitted Data

AZ-81O Digital Air Data Computer ARINC 429 Outputs

RT-300 Radio Altimeter Receiver Transmitter LeadingParticulars

AT-222

DU-880

SG-884

SG-884

SG-884

Radio Altimeter Antenna Leading Particulars

Display Unit Leading Particulars

Symbol Generator Leading Particulars

Symbol Generator ASCB Transmitted Data

Svmbol Generator. Part No. 7008570-904.ASCB Tr~nsmitted Data Changes for TCAS/MLS Option

DC-884 Display Controller Leading Particulars

DC-884 Display Controller MLS Output Discrete Logic

DC-884 Display Controller ASCB Transmitted Data

DC-884 Display Controller, Part No. 7007540-941/942,ASCB Transmitted Data Changes for TCAS/MLS Option

DA-884 Data Acquisition Unit Leading Particulars

DA-884 Data Acquisition Unit ASCB Transmitted Data

DP-884 Dimmer Panel Leading Particulars

FC-880 Fault Warning Computer Leading Particulars

FC-880 Fault Warning Computer ASCB Transmitted Data

FZ-820 Flight Guidance Computer Leading Particulars

FZ-820 Flight Guidance Computer ASCB Transmitted Data

GP-820 Flight Guidance Controller Leading Particulars

PC-880 Turn Pitch Controller Leading Particulars

Pacle

131

132

135

140

144

147

156

161

166.1

172

196.1

197

198.4.1

198.10

198.12

198.22

198.25

198.27

198.39

198.41

198.50

198.58

44-14-00Page TC-19Apr 15/93


Table

7-5

8-1

8-2

8-3

8-4

8-5

8-6

9-1

9-2

9-3

9-4

9-5

9-6

9-7

9-8

9-9

1o-1

10-2

11-1

11-2

11-3

11-4

Honeywell ff~~~wc’


List of Tables

VOLUME I

Title

SM-600 and TM-260 Dual Servo Leading Particulars

WR-800 Receiver Transmitter Leading Particulars

WC-81O Weather Radar Controller Leading Particulars

WA-800 Antenna Pedestal and FP-900 Flat-Plate RadiatorLeading Particulars

WU-870 Antenna and Receiver Transmitter LeadingParticulars

WC-874 Weather Radar Controller Leading Particulars

Fault Display Format

NZ-920 Navigation Computer Leading Particulars

NZ-920 Navigation Computer

NZ-920 Navigation Computer ASCB TransmittedBackground Data

CD-81O

DL-800

DL-900

PZ-800

PZ-800

SM-81O

Engine

Control Display Unit Leading Particulars

Data Loader Leading Particulars

Data Loader Leading Particulars

Performance Computer Leading Particulars

Performance Computer ASCB Transmitted Data

Servo Leading Particulars

Pressure Ratio Transmitter Leading Part

Engine Pressure Ratio Transmitter Input/OutputInformation

OZ-800 Receiver Processor Unit Leading Particu”

Label 270 Discrete Word 1 Format

CU1ars

ars

Label 271 Discrete Word 2 or Label 272 DiscreteWord 3 Format

AT-800 Antenna Coupler Unit Leading Particularsnn

PaQe

198.65

198.73

198.78

198.84

198.89

198.92

198.97

198.103

198.106

198.110

198.129

198.138

198.139

198.143

198.144

198.153

198.154

198.157

198.161

198.162

198.163

198.167

U-14”00Page TC-20Apr 15/93


Table

I

11-5

11-6

12-1

12-2

12-3

12-4

12-5

13-1

13-2

13-3

13-4

13-5

14-1

14-2

14-3

14-4

15-1

15-2

15-3

15-4

15-5

201

202

203

AT-801

AT-803

LP-850

LU-860

LU-860

AT-850

AT-855

RT-91O

RT-91O

RT-91O

Mode S

AT-91O

ML-850

ML-850

CM-850

CM-850

G1obal


List of Tables

VOLUME I

Title

Antenna Coupler Unit Leading Particulars

Antenna Coupler Unit Leading Particulars

Lightning Sensor Processor Leading Particulars

Lightning Sensor Controller Leading Particulars

Switch Functions

Antenna Leading Particulars

Antenna Leading Particulars

TCAS Computer Leading Particulars

TCAS Computer ARINC 429 Output Data

TCAS Computer to Mode S Transponder Data

Transponder to TCAS Computer Data

Directional Antenna Leading Particulars

MLS

MLS

MLS

MLS

Receiver Leading Particulars

receiver ARINC 429 Outputs

Control/Display Unit Leading Particulars

Control/Display Unit Block Diagram

~

198.169

198.171

198.173

198.176

198.177

198.180

198.182

198.185

198.186

198.187

198.187

198.189

198.192

198.194

198.197

198.203

Positioning System Sensor Unit Leading Particulars 198.204

GPSSU Digital Accuracy and Resolution 198.206

GPSSU ARINC 429 Output Data (BNR Format) 198.207

GPSSU ARINC 429 Output Data (BCD Format) 198.208

GPSSUARINC 429 Output Data (DIS Format) 198.208

Flight Director/Autothrottle Mode Annunciations 212

CAS Red Warning Messages 281

CAS Amber Caution Messages 283

22-14-00Page TC-21Apr 15/93


Table

204

205

206

207

208

209

210

211

301

302

401

402

403

404

405

406

407

408

409

410

. 411


List of Tables

VOLUME I

Title

CAS Blue Advisory Messages

System Performance/OperatingLimits

ARINC 429 NAV Computer Output Data

Output Word Formats

Autothrottle Performance Limits

Advisory Messages (Blue)

Engine Synchronization

Climb and Descent Schedule

Ground

Ground

Paqe

285

298.37

298.158

298.159

298.177

298.178

298.202

298.216

VOLUME II

Maintenance Test Procedure 303

Check Procedure 398.285

Test Mode

Test Mode

Test Mode

VOLUME III

ARINC 429 Output Values 404

ASCB Output Values 407

outputs 409

ISDU Display of IRU Test Mode Outputs

Abbreviations for Test Modes

DADC Self-Test Analog Outputs

DADC Self-Test ARINC 429 Outputs

DADC ASCB Self-Test Outputs

Engine and Aircraft Trend and Limit Exceedance Parameters

APU Recording Parameters

Steady State Flight Condition Parameters

409

410

414

414

415

427

427

430

22-14-00Page TC-22Apr 15/93


Tab7e

412

413

414

415

416

417

418

419

420

421

422

501


List of Tables

VOLUME III

Title

Engine Trend Data Recording Parameters

APU Trend Recording Parameters

Parameters Monitored for Exceedance Event Recording

Engine Exceedance Recording Parameters

APU Exceedance Recording Parameters

Minimum Wiring/Power Requirement for FZ-820 to Run GMT

Minimum Servo Wiring Required for FZ-820 to SuccessfullyPower-Up

Normal Switch States

Display Format

Data Loader Fault Codes

Label 353 Fault Codes for

Interconnect Information

AC03 and BC03

~

430

431

432

435

435

444

445

453

460

470

472

502

22-14-00Page TC-23/TC-24

Apr 15/93Use or disclosure of information on this page is subject to the restrictions on the title page of this document,

INTRODUCTION

II

This manual provides general system maintenance instructions and theory ofoperation for the SPZ-8000 Digital Automatic Flight Control System for GulfstreamIV aircraft.

This manual provides block diagram information and interconnectdiagrams topermit a general understanding of System interface.

Common system maintenance procedures are not presented in this manual. The bestestablished shop and flight line practices should be used.

Refer to the Handling, Storage, and Shipping Procedures for Honeywell AvionicsEquipment Instruction Manual, Honeywell Pub. No. 09-1100-01, for information onshipping and storage of the System components.

Additional information on subsystems installed in the Gulfstream IV is availablein the following publications:

Table A is a loaic truth table for use as an aid in understanding the loqicfunctions on th; block diagrams in Sections 2 and 3.

Title

SPZ-8000 DAFCS (Phase II) Pilot’s Manual for theGulfstream IV

LASEREF@ II Installation Manual

LASERTRAKW II NDU Installation Manual

LASEREF@ II Pilot’s Manual

LASERTRAKW NDU Pilot’s Manual

LASERTRAK’”II NDU (with GPS readout) Pilot’s Manual

LASEREF@ II GPIRS Pilot’s Manual

GPIRS Installation Manual

Global Positioning System Sensor Unit (GPSSU)Installation Manual

Air Data ComDuter and Servoed Altimeter SystemTest and Inspection Technical

AA-300 Radio Altimeter Operat-

FMZ-600/800 Fliqht Manacjementand Maintenance-Manual

Use or disclosure of information

Newsletter “

on and InstallationManual

System (FMS) Installation

Publication No.

28-1146-64-00

95-8352

95-8308

95-8351

95-8440

95-8711

28-3341-001

15-3341-006

95-8698

23-1980-04

15-3321-06

A15-1147-15

22-14-00Page INTRO-1

Apr 15/93on this page issubject to the restrictions on the title page of this document.

Title

Gulfstream IV Performance Index

PRIMUS@ 800 COLORADARW System Description andInstallation Manual

PRIMUS@ 800 Pilot’s Handbook

PRIMUS@ 870 COLORADARW System Description andInstallation Manual

PRIMUS@ 870 Pilot’s Handbook

Radar Spoking

REACT Operation

Lightning Sensor System Description and Installation Manual

LSZ-850 Lightning Sensor System Pilot’s Handbook

FMZ-800 Flight Management System Pilot’s Operating Manual

TCAS II System Description and InstallationManual

TCAS II Pilot’s Operating Handbook

MLS System Description and Installation Manual

I

Abbreviations used in

Abbreviation

ACA/CACARSACCELADCAddrADFADIA/DAFCSAFGCSAFISAGCAGLAHRSAHRUA/IAILALRTALT

this manual are defined as follows:

EQuiva~ent

Alternating CurrentAircraft

Publication No.

28-1146-75-00

IB8023137

IB8023135-R1

A09-3946-01

28-1146-56

23-1988-19

23-1988-18

A09-3950-01

28-1146-54

28-1146-43

15-3840-001

28-1146-70-01

A15-3800-02

ARINC CommunicationsAddressing and Reporting SystemAccelerometer, AccelerationAir Data ComputerAddressAutomatic Direction FinderAttitude Director IndicatorAnalog to DigitalAutomatic Flight Control SystemAutomatic Flight Guidance Control SystemAirborne Flight Information SystemAutomatic Gain ControlAbove Ground LevelAttitude and Heading Reference SystemAttitude and Heading Reference UnitAnti-IceAileronAlertAltitude 22-14-00

Page INTRO-2Apr 15/93


I

Abbreviation

AMPSANN, ANNUNANTAOAAOSSAP, A/PAPEAPP, APRPP, APRAPSAPUARINCARMASASCBAT, A/TATCATRATTAUXAZ13/ABAROBATTBCBCDBITBITEBNRBODBOWBRGBRKBTMSCAACAPCAS

::WCDICDUCDSCDUCECHCHGCKSUMCKTCLBCLKCLRCMDCMPTR

Equivalent

AmperesAnnunciatorAntennaAngle of AttackAfter Over Station SensorAutopilotAutopilot EngageApproachAltitude PreselectAuxiliary Power UnitAeronautical Radio, Inc.ArmedAirspeedAvionics Standard Communications BusAutothrottleAir Traffic ControlAir Transport RequirementAttitudeAuxiliaryAzimuthBank AngleBarometricBatteryBack Course or Bus ControllerBinary-Coded-DecimalBuilt-In TestBuilt-In Test EquipmentBinaryBottom of DescentBasic Operating WeightBearingBrakeBrake Temperature Monitoring SystemCivil Aviation Authority (British)CaptureCrew Alerting System or Calibrated AirspeedCircuit BreakerCounterclockwiseCourse Deviation IndicatorControl Display UnitDifferential ResolverControl Display UnitCourse ErrorChannelChangeCheck SumCircuitClimbClockClearCommandComputer 22-14-00

Page INTRO~3Apr 15/93


Abbreviation

CNTLCOMCOMBCOMPCONFIGCONTCORRCosCPCPLCPUCRCCRSCRTCRZCsCTCTRLCwDAD/ADADCDAUDBDCDCTDDMDEFLDEGDEMODDESDETDEV, DEVNDGDHDIFFDISCDISENGDISPDISPLDISTDLDMADMEDNDRCDSRDTGDTRK

HPE

I

Equivalent

ControlCommonCombinedCompensation, Compass, or CompactedConfigurationControllerCorrectionCosineCross Pointers, CopilotCoupleCentral Processor UnitCyclic Redundancy CheckCourseCathode Ray TubeCruiseCross SideControl TransformerControlClockwiseDrift AngleDigital to AnalogDigital Air Data ComputerData Acquisition UnitDatabaseDisplay ControllerDirectDifference in Depth of ModulationDeflectionDegreeDemodulatorDescentDetector, DetentDeviationDirectional GyroDecision HeightDifferential, DifferenceDisconnectDisengageDisplayDisplacementDistanceData LoaderDirect Memory AccessDistance Measuring EquipmentDownDual Remote CompensatorDesiredDistance To GoDesired TrackDisplay UnitDu~licate


Apr 15/93Use or disclosure of informationon this page is subject to the restrictions onthe title page of this document,

Abbreviation

EEPROM, E2PROMEFISEGTEDSEICASEL, ELEVEMIENG

~OOFFEPREPROMESSETETAETEEVMEX LOCEXTFAAFARFD, F/DFDBK

IFFFFS

IFGCFLFLCHFLTFMSFPL, FPLNFPMFR. FRM

IFREQFTIU

F;CGA, G/AGCRGMAPGMTGPGPSGRDGS, G/SGSPD

NHBMHDGHDLCHFHORIZ

Hone~ell &$~~~~~~E

E~uivalent

Electronically Erasable ProgrammableElectronic Flight Instrument System

Read Only Memory

Exhaust Gas TemperatureElectronic Display SystemEngine Instrument and Crew Alerting SystemElevator, ElevationElectromagnetic InterferenceEngage, EngineEasy-OnEasy-OffEngine Pressure RatioErasable PROMEssentialElapsed TimeEstimated Time of ArrivalEstimated Time In RouteEngine Vibration MonitorExpanded LocalizerExtend, ExternalFederal Aviation AuthorityFederal Aviation RequirementFlight DirectorFeedbackFuel FlowFlight Fault SummaryFlight Guidance ComputerFlight LevelFlight Level ChangeF1ightFlight Management SystemFlight PlanFeet Per MinuteFromFrequencyFlight Test Interface UnitFlux ValveFault Warning ComputerGo-AroundGround Clutter ReductionGround MapGreenwich Mean TimeGuidance PanelGlobal Positioning SystemGroundGlideslope, GroundspeedGroundspeedGross WeightHeartbeatHeartbeat MonitorHeadingHigh Level Data Link ControlHigh FrequencyHorizontal

22-14-00Page INTRO~5


I

Abbreviation

HPHRHSIHYDH/WIASICAOID, IDENTIFIGNILSINC-DECINDINITINSINTFCINTGLINTLKINV1/0IRCIRSIRU1S0IVVKBPSkHzKN, KTLLATLBSL/CLHLIMLNAVLOCLONLORANLPLPVLRNLRULSBLTGLTSLVCLVDTMAGMAGVARMDSMFDMHz

Ecluivalent

High PressureHourHorizontal Situation IndicatorHydraulicHardwareIndicated AirspeedInternational Civil Aviation OrganizationIdentificationIntermediate FrequencyIgnitionInstrument Landing SystemIncrease-DecreaseIndicatorInitializationInertial Navigation SystemInterfaceIntegralInterlockInvertInput/OutputInstrument Remote ControllerInertial Reference SystemInertial Reference UnitIsolationInstantaneous Vertical VelocityKilo Bits Per SecondKilohertzKnotsLeftLatitudeLateral Beam SensorInductive/CapacitiveLeft HandLimitLateral Navigation/Lateral GuidanceLocalizerLongitudePosition Sensor TypeLow PressureLatched Power ValidLong Range NavigationLine Replaceable UnitLeast Significant BitLightingLong Term SensorLine Voltage CompensationLinear Variable Differential TransformerMagneticMagnetic VariationMinimum Discernible SignalMultifunction DisplayMegahertz 22-14-00



Abbreviation

MINMLSMM

MONMSBMSGMSLM/T, M/TRIMMUXN

I NAHPNAVNCNCD

INDNDBNMNONOCNORMNOTAMNRZ

I NVRAMNZ

I OATOBSo/cOMOsc0ssPPATTPBPcPERFPFDPISOPITPITCH SYNCPLAPLNPMSPosPPHPPOSPRESSPRI, PRIMPROCPROFPROGPROM

Equivalent

MinutesMicrowave Landing SystemMiddle MarkerMaximum Allowable Mach NumberMomentaryMonitorMost Significant BitMessageMean Sea LevelMach TrimMultiplexerNorthNot a Honeywell PartNavigationNo Connection, Normally Closed, or NAV ComputerNo Computer DataNavigation DisplayNondirectional Beacon, Navigation Data BaseNautical MileNormally OpenNAV on CourseNormalNotice To AirmanNonreturn To ZeroNon-Volatile RAMNavigation ComputerOutside Air TemperatureOmni Bearing SelectorOn CourseOuter MarkerOscillatorOver Station SensorPressurePitch AttitudePushbuttonPerformance ComputerPerformancePrimary Flight DisplayParallel In Serial OutPitchPitch SynchronizationPower Lever AngleP1anPerformance Management SystemPositionPounds Per HourPresent PositionPressurePrimaryProcessorProfileProgrammer, ProgrammingProgrammable Read Only Memory


Apr 15/93Use or disclosure of information on this page issubject to the restrictions on the title page of this document.

Abbreviation

P/sPSIPv

;!IMPWR

;EQTYRRA, R/A, RAD ALT

IRAMRCT, REACTRCVRREFRELRETRETRREVRGRHRMIRN, RNAVRNAPPRNGROLROMRPMRT, R/TRUDsSATSBY, STBYSCISCRScsSDISECSELSGSIDSIGSINSINGSIPO

ISPDSPSSRAMSRNSSECSSMSTAEL

Equivalent

Pitot SwitchPounds Per Square InchPower ValidPitch Wheel or Pulse WidthPulse Width ModulatedPowerPerformance ComputerQuality FactorQuantityRightRadio AltimeterRandom Access MemoryRain Echo Attenuation Compensation TechniqueReceiverReferenceReleaseReturnRetractReverse Source (Same as Back Course)Rate GyroRight HandRadio Magnetic IndicatorArea NAVRNAV ApproachRangeRol1Read Only MemoryRevolutions Per MinuteReceiver/Transmitter,Rate-of-TurnRudderSouthStatic Air TemperatureStandbySerial Control InterfaceSourceSingle Channel SelectSource/Destination IdentificationSeconds, SecondarySelectSymbol GeneratorStandard Instrument DepartureSignalSineSingleSerial In Parallel OutSpeedSamples Per SecondSystem Random Access MemoryShort Range NavigationStatic Source Error CorrectionSign Status MatrixSt~tion Elevation 22-14-00



Abbreviation

STARSTBYSTCSTCSSTPSTR, STRGSvoSW> s/wSYNCSYSTTASTAT

ITBDTCASTCSTEMPTGT

:[ETKODTLA

ITLETOTOCTODTOGATPTRKTRSTRUTSOTTGTTLUARTUTIL, UTYV1

IV2

VAVALVANGVAPP, VAPR

IVARVASL, VASELVBSVcoVEL

IVERTVFLC, VFLCH

E~uivalent

Standard Terminal Arrival RouteStandbySensitivity Time ControlSingle Trim Channel SelectSteepSteeringStart Valve OpenSoftwareSynchronizationSystemTemperatureTrue AirspeedTrue Air TemperatureTo Be DeterminedTraffic Alert and Collision Avoidance SystemTouch Control SteeringTemperatureTarget, Turbine Gas TemperatureTurn Knob, TrackTrack ErrorTurn Knob Out of DetentTorque Limit AileronTorque Limit ElevatorTake OffTop Of ClimbTop Of DescentTakeoff Go-AroundTest PointTrackThrust Reference SetTrueTechnical Standard OrderTime-To-GoTuned to LocalizerUniversal Asynchronous Receiver TransmitterUtilityTakeoff Decision SpeedTakeoff Safety Speed (Speed to be attained at35 feet AGL, assuming recognition of an enginefailure after Vl)Volt AmpereValidVertical AngleVOR ApproachVariableVNAV Altitude PreselectVertical Beam SensorVoltage Controlled OscillatorVelocityVerticalVNAV Flight Level Change


Apr 15/93Useor disclosure of information onthispage issubject to the restrictions on the title page of this document.

MAINTENANCEMANUALGULFSTREAM IV

Equivalent

I

Abbreviation

Vfs

VGVHFVLDVLFvV~AVVORVORTACVPTH, VPATHVrVRAMVrefVRT, VERTVs, v/sVse

VTAwWL, W/LWowowWPTWR, WXWSPXFERXMTRX-SIDEXTRKYD, Y/D

Final Takeoff Climb Speed (Airspeed for single-engineclimb in a clean confiwration below 1500 feet AGC)Vertical Gyro or Verti;al GeneratorVery High FrequencyValidVery Low FrequencyMaximum Allowable AirspeedVertical NavigationVHF Omni RangeCollocated VOR and Tacan StationsVertical PathTakeoff Rotation SpeedVideo Random Access MemoryReference Speed (Landing)VerticalVertical SpeedEnroute Climb Speed (Airspeed for single-engine climbin a clean configuration above 1500 feet AGC)Vertical Track AlertWestWings LevelWashed OutWeight-on-WheelsWaypointWeather RadarWord Sequence PositionTransferTransmitterCross-SideCross TRACKYaw Damper

Honeywell has an airworthiness analysis procedure performed for all its airborneproducts to ensure that equipment designed by Honeywell will not create ahazardous in-flight condition. As a result of the analysis, certain installa-tions have been designated INSTALLATION CRITICAL, and 100 percent compliance withthose installations is required.

INSTALLATION CRITICAL is defined as:

Specific methods of installation are required to ensure that either thefailure of the assembly or part is extremely improbable or that itsfailure could not create a hazardous condition.


Apr 15/93Useordisclosure ofinformation on this page issubject totheresttictions onthetitle page of this document.

. . m. MAINTENANCE

s

Q

T

R

CLR

ELECTRONIC SWITCH (FILP-FLOP) CHARACTERISTICS1. AN OUTPUT IS ALWAYS PRESENT AT EITHER Q OR 6 (BISTABLE).2. CLEAR (CLR) IS CONTINOUS INPUT AND WILL CAUSE OUTPUT TO SWITCH TO 6.3. RESET (R) IS PULSED INPUT AND ALWAYS CAUSES OUTPUT TO SWITCH TO 6.4. SET (S) IS PULSED INPUT AND ALWAYS CAUSES OUTPUT TO SWITCH TO Q.5. TOGGLE (T) IS MOMENTARY INPUT AND CAUSES OUTPUT TO SWITCH TO Q OR 6 SUCCESSIVELY

AC-3452E@

Logic Truth TableTable A 22-14-00

Page INTRO-n/INTRO-12Apr 15/93


SECTION 1SYSTEM DESCRIPTION

1. General

The SPZ-8000 Digital Automatic Flight Control System (Figure 1) providesflight director, autopilot, pitch trim, Mach trim,system operates in conjunction with the electronicconsists of the primary flight displays (PFD), theand the engine instrument and crew alerting system

Table 1 lists the components and part numbers thatFigure 2 shows the approximate component locationsinstallation.

and autothrottle. Thedisplay system (EDS) thatnavigation displays (ND),(EICAS) displays.

compose a System, andfor a Gulfstream IV

A/CHoneywell Ref

System Component Qty Part No. Des

AZ-81O Digital Air Data Computer 2 7000700-964, 9[C9-864 (CAA or

ASC 61)

FZ-820 Flight Guidance Computer 2 7003974-905(PHASE 2) lo/clo-906 ****_907 ****

~: The -907 FZ-820 Flight Guidance Computers were factory installedon aircraft 1198 and subsequent.

GP-820 Flight Guidance Controller 1 7007546-901/-902* 11

SM-600 Dual Servo (Aileron) 1 4015373-705 12

SB-600 Servo Bracket (Aileron) 1 4015374-905 N/A

SM-600 Dual Servo (Elevator) 1 4015373-704 13

SB-600 Servo Bracket (Elevator) 1 4015”374-904 N/A

RT-300 Radio Altimeter Receiver 2 7001840-922 20/c20Transmitter

AT-222 Antenna (Receive) 2 4007637-002 21/c21

System ComponentsTable 1 22-14-00

Page 1Apr 15/93


A/ctii;:ypl Ref

System Component Qty . Des

AT-222 Antenna (Transmit) 2 4007637-002 22/c22

TM-260 Dual TRIM Servo 1 7000260-602 29/c29

TB-261 Trim Servo Bracket 1 7000261-603 N/A

~: The following four PRIMUS@ 800 Weather Radar System Components werefactory installed on aircraft 1000 thru 1071.

WR-800 Weather Radar Receiver 1 M1585350-34 59Transmitter

WA-800 Antenna Pedestal 1 M1585354 60

FP-900 Flat Plate Radiator 1 M1585377 N/A

WC-81O Weather Radar Controller 2 7006921-311/-3l2* 61/c61

~: The following two PRIMUS@ 870 Weather Radar System components werefactory installed on aircraft 1072 thru 1119.

WU-870 Antenna and Receiver 1 7012640-902 59Transmitter Unit

WC-874 Weather Radar Controller 2 7006921-413/-4l4* 61/c61

~: The following two PRIMUS@ 870 Weather Radar System components werefactory installed on aircraft 1120 and subsequent.

WU-870 Antenna and Receiver 1 7012640-904 59Transmitter Unit

WC-874 Weather Radar Controller 2 7006921-415/-4l6* 61/c61

SG-884 Symbol Generator 3 7008570-903 (PHASE 2) 65/C65/-904 (TCAS 11/

MLS OPTION)-913 (TACAN

OPTION)

DC-884 Display Controller 2 7007540-931/-932 115/cl15(PHASE 2)

-941/-942*(~~};o~~/MLS

-951/-952*(TACAN OPTION)

CD-81O Control Display Unit 2 7007549-901/-902* 120/c120

System ComponentsTable 1 (cent) 22-14-00

Page 2Apr 15/93


A/CRef

System Component Qty Honeywell Part No. Des

NZ-920

NOTE:

PZ-800

NOTE-—.

DL-800

DL-900

SM-81O

PC-880

DU-880

AY-003

DU-880

DU-880

FC-880

NOTE.—.

MD-880

DP-884

DA-884

Navigation Computer 2 7004402-963 (PHASE 2) 121/C121-964 (PHASE 2 WITH AFIS)-976 (9101 SOFTWARE)-978 (9111 = 9101

SOFTWARE WITHOUTGPS BLENDING)

The -976 NZ-920 NAV Computers were factory installed on aircraft1183 and subsequent.

Performance Computer 2 7004609-906 (PHASE 2) 122/C122-910 (9101 SOFTWARE)

The -910 PZ-800 Performance Computers were factory installed on air-craft 1183 and subsequent. The -910 must also be used with ASC 61.

Data Loader 1 7004607-901/-9O2* 123

Da;: Loader 1 7016600-901/-9O2* 123

Servo (Autothrottle) 2 7009025-913 128/C128

Turn Pitch Controller 1 7007990-901/-902* 129

Display Unit (PFD) 2 4053000-902/-9Ol* 130/c130

Inclinometer Kit for PFD 2 7005400-905/-906* - -

Display Unit (ND) 2 4053000-902/-9Ol* 131/c131

Display Unit (EICAS) 2 4053000-902/-9Ol* 132/133

Fault Warning Computer 2 7007484-904(PHASE 2) 134/C134-905 *****-914 *****

The following MD-880 Part No. is for a blank module.Use GAC Part No.

Checklist Module

Dimmer Panel

Data Acquisition

for reorder.

2 7010405-903 --

1 7007543-901/-902* 135

Unit 2 7007580-901 136/137

System Components 22-14-00Table 1 (cent) Page 3

Apr 15/93Use or disclosure of information on this page is subject to the restrictions on the title page of this document.

A/cRef


LASEREF@ II Inertial Reference Unit 2 HG1075AEO3 (PHASE 2) N/AHG1075AEO4 (See Note)HG1075GEO4 (See Note)

NOTE: The HG1075AEO4 LASEREF @ II were factory installed on aircraft 1164and subsequent. The HG1075AEO4 functions identically to the AE03, butthe AE04 has been reconfigured to facilitate updating to the optionalHG1075GEO4. The GE04 has the same functions as the AE03 and AE04, butwill also interface with GPS and provide IRS, GPS, and hybrid GPS/IRSoutputs.

Attitude Heading Reference Unit 1** HG1076AAO1 N/A

Mode Select Unit 1 CG1227AC10/ACOl* N/A

Inertial System Display Unit (ISDU) 1 CGl136AC10/AC20* N/A

LASERTRAKW Navigation Display Unit l*** CG1230AC10/AC20* N/ACG1230AC11/AC21(EFIS DISPLAY OPTION)CG1230AG11/AG21*(GPS READ OUTS)

Engine Pressure Ratio Transmitter 2 LG1189BC03(4063258-3)

O~tional VLF/Omeqa System Components

OZ-800 Receiver Processor Unit 1 7004608-901(Omega/VLF)

AT-800 Antenna Coupler Unit, 1 7011102Teardrop H-Field

AT-801 Ant~;na Coupler Unit, 1 7011103Brick H-Field

AT-803 Ant~;na Coupler Unit, 1 7011100Blade E-Field

Oc)tionalLSZ-850 LicthtninqSensor System Com~onents

LP-850 Lightning Sensor Processor 1 7011822-903

LU-860- Lightning Sensor Controller 1 7012738-905/-906*

N/A

141

142

142

142

145

146

System Components 22-14-00Table 1 (cent) Page 4


A/CRef


AT-850 Lightning Sensor Antenna 1 4057697-901 147(Teardrop)

AT-855 Lig;;ning Sensor Antenna 70146062-901 147(Brick)

OPTIONAL TCZ-91O TRAFFIC ALERT AND COLLISION AVOIDANCE SYSTEM (TCAS II)COMPONENTS

RT-91O TCAS Computer 4066010-902AT-91O TCAS Antenna : 7514060-902

OPTIONAL MLZ-850 MICROWAVE LANDING SYSTEM [MLS) COMPONENTS

ML-850 MLS Receiver 2 7510600-901CM-850 MLS Control/Display Unit 2 7513004-913/914*

OPTIONAL GLOBAL POSITIONING SYSTEM (GPS) COMPONENTS

Global Positioning System lor2 HG2021AB02Sensor Unit (GPSSU)

* The difference between dash numbers is the bezel color. The first dashnumber listed is for a gray bezel unit and the other dash number is fora black bezel unit.

** A third inertial reference unit may be used in place of the AHRU.

*** The LASERTRAKW may be used in place of the ISDU. The optional AC1l/AC21LASERTRAK’”outputs additional ARINC 429 data which is necessary fordisplay of LASERTRAK’”on EFIS. The AG1l/AG21 LASERTRAK’”is available toallow GPS readouts on the LASERTRAKW display.

**** The -906 Flight Guidance Computer (FGC) is an update to the standardphase 2 -905 FGC, which improves monitor function in turbulence andreduces go-around angle to 12 degrees. The -907 FGC is identical tothe -906 FGC with the addition of non-volatile memory flight faultsummary with english readout.

***** The -905 Fault Warning Computer (FWC) adds TCAS 11/MLS, BTMS, andforeign certification recognition. These were factory installed onaircraft 1168 and subsequent. The -914 FWC adds foreign certificationrecognition without TCAS 11/MLS or BTMS.

System ComponentsTable 1 (cent) 22-14-00

Page 4.1/4.2Apr 15/93


I 1

H,m.wllTEME’Czzzl~!ll

~ DISPMV UNIT (FFOl ~ OISFUV ~lT (ND) OIMM DISPUV UNIT @CAS) ma OwLAv UNIT (EICAS) oU.SEoolsFLAvw4rr rFAol ~ OISPUY LHIT (FFO)

oPds4DIMMERPANEL

INERTIALREFEREI$CEUNIT

Q“ ,& ‘=

fflERILALREFERENcE

3[

ARINC wC.074 $vEAT14ER

q -. :4= RADARCONTRCUER

‘$!J

I!: ~.=~:,JJ,,,!t-k 1111111111

GENERATOR

H I I 11~~ “ J I H.----J II 1

1I

vL

1

&AR(NC4i?9TO CtSJSSSIC+EPRT

:.. .

.“%

:.Az&O

AflltW 429 TO 0K31mlSGt, SG2, SG3 AIR WTA

coMPurEFr

14 Ii

PC-SW NANLJAlCONTROLLER

m, ;L-

P

----

~~~ ~ ‘= -~7p

~\ “*

ENGINE PRESSUREENGINE PRESSURERATlo TRANSMImER * !%l_Tr

1::%%:FZ4Zt RATIOTRANSMITTERFLIONT

‘k ~+ ‘ ‘+ ‘“

QUIMNCE GuIMNCE PERFORMANCECOMPUTER COMRJTER

PERFORMANCE NAwaATloNmMPuTEn COFAWTER (XWUTER

O&904 12ATAM.XUISITION UNIT

.+2

T

.- -‘“.,-’ +==-ENGINE .-.5- AT-~ ANTENNA

SENSOR

~:’. r +-..RT401 RAOIO AT-3?2ANTENNAALTIMETERRECEIVEWTRANSMITTER

1 I 1

t t f,, ..,:

[i ‘lb ‘“-”“. .,

f$~$f~ flbs==-> J

AT-=ANTENNA

wow ;* IEr4m4E

w-h SENSOR

8/ ‘wATM \\ ‘“):, : Ib.]\“ ~:.

.-:.: 4

-i

tl:Tlm&ur AT-~

SAA-UMOUAL S“- OUAL TM-SSUOUALSERVO (AILEROU) SERVO (ELIWATOR) SERVO (TRIMI

SM-S1OSERVO - - “ SM-S1OSERVO -- RECEIVEW(AUTOTHROTTLE) THRomE (AUTOTHROTTLE) Tnml

SPZ-8000 System Flow DiagramFiaure 1 22-14-00

.—-—.._..=LE TRANSMITTER ~.1241E-M

d-–

Page 5/6Apr 15/93


1.NOSECOMPARTMENT COMPONENTS:

● WR@(J@WEATHERRADAR RECEIVERTRANSMITTER● WA-800ANTENNAPEDESTAL● FP-sOOFLATPLATEANTENNA~

● wu-870ANTENNAAND RcvR/xMTRuNIT2.INSTRUMENTPANELCOMPONENTS:

● DU-880 DISPLAYUNITS(PFO)

● DU-880 DISPLAY UNITS (ND)● DU-S80 DISPLAY UNITS (EICAS]

3. GIARE SHIELD COMPONENTS:● GP-820 FLlf3HTGUlf3ANCE CONTROLLER● DC-884 DISPLAY CONTROLLER

4. WC-81O OR WC-874 WEATHER RADAR CONTROLLER (SIDE PANEL)

6. PEDESTAL COMPONENTS:● CD-81O CONTROL DISPLAY UNIT● PC-880 TURN PITCH CONTROLLER● DP-884 DIMMER PANEL● MODE SELECTUNIT

& ELECTRONICBAYCOMPONENTS (NOTE):● AZ-81ODIGITALAIRDATACOMPUTER● FZ-820FLIGHTGUIDANCECOMPUTER● SG-884SYMBOLGENERATOR(SG3PILOT’SSIDE)● NZ-9XX NAVIGATION COMPUTER● PZ-800PERFORMANCE COMPUTER● FC-880FAULTWARNING COMPUTER● DA-880DATAACQUISITIONUNIT● INERTIALREFERENCEUNITSOR THIRDAHRU

7.SM-600DUALSERVO(AILERON)8.TM-260DUALTRIMSERVO[UNDERPEDESTAL)9.AT-222ANTENNAS(UNDERFLOOR)10.RT-300RADIOALTIMETERRCVR/XMIT(UNDERFLOOR)11.SM-S1OAUTOTHROTTLESERVO(UNDERFLOOR)12,RUDDERACTUATOR(NONHONEYWEU)13.SM-600DUALSERVO(ELEVATOR)14.ENGINEPRESSURERATIOTRANSMITTER

NOTE:THEREISONE EACH COMPONENTINEACH ELECTRONICBAY,EXCEPTTHETHIRDSG-884OR lRU/At-lRUISINTHE BAY ON THE PILOT’SSIDE.

AD-12490-R4

Component Locations for aTypical“Gulfstream IV Installation

Figure 2 22-14-00Page 7/8

Aug 15/91Use or disclosure of mlorma[lon on VM page /s subject to the restrlc[lons on the Mle page of Ibis document,

2. System Descrii)tion

IThe SPZ-8000 DAFCS consists of the following subsystems which are describedin paragraphs 2.A through 2.N.

● LASEREF@ II Inertial Reference System (IRS)

● ADZ-81O Air Data System

● AA-300 Radio Altimeter System

G EDZ-884

● DFZ-820

. PRIMUS@

c PRIMUS@

6 FMZ-800

Electronic Display System (EDS)

Dual Flight Guidance System

800 Weather Radar System

870 Weather Radar System

Flight Management System (FMS)

. Engine Pressure Ratio System

I

. VLF/Omega System (Optional)

● LSZ-850 Lightning Sensor System (Optional)

I. TCZ-91O Traffic Alert and

● MLZ-850 Microwave Landing

. Global Positioning System

Collision Avoidance System (Optional)

System (Optional)

(Optional)

The SPZ-8000 is a complete automatic flight control system providing fail-operational execution of flight director guidance, autopilot, yaw damper, andtrim functions. The automatic path mode commands are generated by the FZ-820Flight Guidance Computer which integrates the attitude and heading reference,air data, EDS, and FMS into a complete aircraft control system that rovides

Rthe stabilization and control needed to ensure optimum performance t roughoutthe aircraft flight regime.

A central serial communications network provides inter-subsystem communi-cations within the system. The network is denoted by the nomenclatureavionics standard communications bus ASCB). This bus structure usesadvanced communications techniques anJ safety design features to rovide high

Rthrough-put, fail-operationaldata exchan e within the System. T e ASCB7consists of two serial synchronousdigita communications buses. Each bus is

electrically isolated from the other bus. In addition, each bus cancommunicate bidirectionally. Refer to paragraph 3 for a detailed descriptionof the ASCB.

The ASCB interfaces the automatic flight control system with the digital airdata system, the IRS, the EDS, and the FMS systems. Each subsystembroadcasts on the ASCB when directed to transmit by the bus controller, andreturns to an off condition when its time slot expires. The bus controllerfunction is triplex and is contained in each symbol generator. Only one buscontroller is active at a time.

22-14-00Page 9

Apr 15/93Use or disclosure of informationon this page is subject to the restrictions onthe title page of this document.

The system data communication is split between theprivate-line paths provided for specific sensitive

main system bus (ASCB) anddata for which fault

isolation is required. These specific private-line paths include thefollowing:

. IRS attitude and headinq data to SG-884 Symbol Generators (~rivate-lineserial bus)

. FZ-820 Flight(private-line

● GP-820 Flight(private-line

. .

Guidance Computer to GP-820 Flight Guidance Controllerserial bus)

Guidance Controller to FZ-820 Flight Guidance Computerserial bus)

● AZ-81O Air Data Computer to SG-884 Symbol Generators (private-lineserial bus)

● SG-884 Symbol Generators to DU-880 EDS displays (private-line serialbus)

. ADF, NAV, and ILS data to SG-884 Symbol Generators (private-line serialbus)

Information from navigation receivers is not interfaced directly to theSPZ-8000 DAFCS but is pre-processed by the EDS and/or the FMS. The EDS shallfunction as data concentrator and switcher for basic navigation and headingdata and shall provide to the DAFCS source identification logic for currentlydisplayed data. The DAFCS shall receive lateral navigation (roll steeringcommands) and vertical navigation (vertical targets) directly via ASCB fromboth FMS subsystems. However, the DAFCS shall use the FMS navigation dataonly if selected for display by the EDS.

The system displays heading, course, radio bearing, pitch and roll attitude,barometric altitude, selected alert altitude, radio altitude, IAS/MACHtargets, lateral and vertical deviation, to-from indications, DMEindications, engine parameters, and advisory and caution indications.Annunciators denote selected flight mode, altitude alert, decision height,and go-around mode engagement. Pitch and roll steering commands developed bythe FZ-820 Flight Guidance Computer in conjunction with the GP-820 FlightGuidance Controller are displayed by steering pointers to enable the pilot toreach and/or maintain the desired flightpath or attitude.

2. A. LASEREF@ II Inertial Reference System (IRS)

The IRS is comprised of the following components:

● Inertial Reference Units

. Attitude Heading Reference Unit

● Mode Select Unit

● Inertial System Display Unit (Optional)

I22-14-00

Page 10Apr 15/93

Use or disclosure of information on this page issubject to the restrictions on thetitle pageof this document.

I

The IRS is an all attitude inertial sensor system which provides aircraftattitude, heading, and flight dynamics information to the PFD and ND, flightcontrol (DAFCS), weather radar antenna platform, and other aircraft systemsand instruments.

The inertial reference unit (IRU) is the primary component of the IRS andprovides the following digital outputs:

. Primary attitude● Body linear accelerations. Body angular rates● Inertial velocity vectors. Magnetic and true north reference. Present position data● Wind data

The IRU requires a dedicated +24 volt dc backup battery, which the user mustsupply.

The triple-channel mode select unit (MSU) enables the flightcrew to selectthe mode of system operation for three IRUS, provides status indication foreach system, and test initiation for each IRU.

The optional IRU 3 is used in triple IRS installations in place of theattitude heading reference unit (AHRU).

The form, fit, and function of the AHRU are identical to those of the IRUexcept that the AHRU does not provide navigation outputs on the digital businterfaces. The AHRU provides continuous comparison monitoring in which itsinternally computed attitude, angular rate, and acceleration signals arecompared to those same signals transmitted on the ASCB data bus from IRU 1and IRU 2. The AHRU also provides backup attitude, heading, rate, andacceleration outputs.

The inertial system display unit (ISDU) selects data from any one of threeIRUS for display and provides position or heading data to three IRUS.

LASERTRAK@ is an extension of the Honeywell Laseref Inertial Reference System(IRS). The function is resident in a navigation display unit (NDU) whichreplaces the ISDU normally found in IRS installations. The NDU retains theISDU functionality and gives the pilot a means of entering a nine-waypointflight plan. Waypoints are entered using lat/long designations. Flight planprogress is monitored on the NDU or on EFIS.

For additional information on the IRS components,‘referto the manuals listedin INTRODUCTION.

22-14-00Page 11


2. B. ADZ-81O Air Data System

The ADZ-81O Air Data System is comprised of the AZ-81O Digital Air DataComputers.

The AZ-81O Digital Air Data Computer (DADC) is a microprocessor-baseddigital computer which accepts both digital and analog inputs, performsdigital computations, and supplies both digital and analog outputs. Itreceives pitot-static pressures and total air temperature inputs forcomputing the standard air data functions. The DADC provides outputs tothe electronic display system, transponder, flight recorder, flightguidance computer, as well as other elements of the system, such as faultwarning computers and inertial reference system (IRS). Angle of attackinputs have been incorporated and computed AOA outputs are provided. TheDADC provides alerting functions.

c. AA-300 Radio Altimeter System

The AA-300 Radio Altimeter System is comprised of the followingcomponents:

. RT-300 Radio Altimeter Receiver Transmitter

. AT-222 Antenna

The AA-300 Series Radio Altimeter System is a high resolution, short-pulse radio altitude system designed for automatic continuous operationover wide variations of terrain, target reflectivity, weather, andaircraft altitude. The radio altimeter provides a dc output voltage andan auxiliary radio altitude output which are proportional to the aircraftabsolute altitude above terrain.

The precision output is used to drive the PFD RAD ALT display andsupplies altitude information to the flight guidance system.

Proper system operation is indicated by the absolute altitude being inview on the PFD. The RAD ALT display will be blank for absolutealtitudes above 2,500 feet. If a failure occurs, the RAD ALT display onthe PFD will show amber dashes. Momentary signal loss within the usablerange will cause the RAD ALT display to blank momentarily.

The AA-300 may be used in-flight to monitor absolute altitude at anyaltitude up to the maximum range of 2,500 feet, or the pilot may selectan alert altitude with the DH set control and be alerted automaticallywhenever the aircraft reaches that altitude. The AA-300 may also be usedfor displaying ground separation and climb conditions during night orinstrument takeoffs, as well as indicating ground clearance duringapproaches.

Pressing the RAD ALT line select button on the TEST menu of the displaycontroller for confidence testing causes the RAD ALT display on the PFDto read approximately 100 feet altitude.

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2. D. EDZ-884 Electronic Display System (EDS)

The EDZ-884 EDS is comprised of the following components:

_ DU-880

. SG-884

c DC-884

● FC-880

● DA-884

● DP-884

Display Units (PFD, ND, and EICAS)

Symbol Generators (SG1, SG2, SG3)

Display Controllers (DC)

Fault Warning Computers (FWC)

Data Acquisition Units (DAU)

Dimmer Panel

The EDS displays pitch and roll attitude, heading, course orientation,flight path commands, weather presentations, mode and sourceannunciations, air data parameters, engine data, and fault warninginformation.

The primary features the EDS brings to the flight control system aredisplay integration, flexibility, and redundancy. Essential displayinformation from sensor systems, and automatic flight control,navigation, performance, and caution-warning systems is integrated intothe pilot’s prime viewing area. Each symbol generator is capable ofdriving six DU-880 displays, such that in case of a symbol generatorfailure, one of the remaining symbol generators drives the PFD and NDdisplays on both sides and the EICAS displays. In the case of a DU-880Display Unit failure, the PFD takes priority over the ND. In the case ofan EICAS display failure, a compacted EICAS format can be displayed oneither EICAS DU-880 Display Unit.

The symbol generator (SG) functions as the data processor for the displaysystem. It receives digital and discrete inputs, organizes thisinformation into the correct formats as defined by the displaycontroller, and transmits these formats to the display units. All analoginformation is input to the EDS through the DAU. The DAU transmits thisdata to the SG over the ASCB.

The symbol generators in the display system are identical and directlyinterchangeablee. When the display system is in its normal (no failure)configuration, SG1 drives the pilot’s displays, SG2 drives the copilot’sdisplays, and SG3 drives the EICAS displays.

The fault warning computer (FWC) is primarily responsible for supplyingdata to the symbol generators for display of warnings, cautions, andadvisories, engine data, and system pages on the EICAS displays. Itreceives data directly from various aircraft systems and from the ASCBfor other aircraft and avionics systems.

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MAINTENANCE

Honeywell !!!!!!%h.Prioritization and message suppression under specific circumstances willbe performed in the FWC to improve the effectiveness of each warning/advisory to the pilots. The FWC will also function as a continuousengine trend and limit monitoring system with the capability to storeselective engine parameters for later readout and analysis by maintenancepersonnel.

The DP-884 Dimmer Panel provides the CRT brightness control for the sixDU-880 Display units.

2. E. DFZ-820 Dual Flight Guidance System

The DFZ-820 Dual Flight Guidance System is comprised of the followingcomponents:

. FZ-820 Flight Guidance Computers (FGC)

● GP-820 Flight Guidance Controller

● PC-880 Turn

. SM-600 Dual

. TM-260 Dual

Pitch Controller

Servo Drive and SB-600 Bracket (Aileron and Elevator)

Trim Servo and TB-261 Bracket

. Dual-Valve Hydraulic Rudder Actuator (non-Honeywell)

The DFZ-820 Flight Guidance System provides full fail-operational flightdirector, autopilot, yaw damper, and trim. Fail-operational capabilityis provided by redundant flight control functions. Existence of dualattitude/heading and air data sensors is used to full advantage withinthe DFZ-820 System through sensor voting and redundancy managementtechniques. The fail-operational characteristic is extended to includesensor failures.

One servo of each dual servo (aileron, elevator, and trim) and the rudderactuator is connected to a flight guidance computer. Only one computerwill be actively controlling the servos and actuator. The servo oractuator connected to the inactive FGC has a brake applied. Normally,the pilot’s side will be automatically in control. The copilot’s sidecan be manually selected, if desired, or will automatically take controlif there is a disengage type failure in the pilot’s computer. In thisway, fail-operationalcontrol is achieved. If only one FGC is valid, thesystem will disconnect if the remaining FGC fails.

The single GP-820 Flight Guidance Controller has dual circuitry thatprovides the same outputs to each FGC and is used to engage the system,select the operating modes, select the PFD in command, arm theautothrottle system, and set the selected heading, course, verticalspeed, speed targets, and altitude preselect.

The single PC-880 Turn Pitch Controller provides dual pitch wheel andturn knob outputs to the FGC.

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2. F. PRIMUS@ 800 Weather Radar System

The PRIMUS@ 800 Weather Radar System consists of the followingcomponents:

● WR-800

● WC-81O

Q WA-800

c FP-900

Weather Radar Receiver Transmitter

Weather Radar Controller

Antenna Pedestal

Flat Plate Radiator

The PRIMUS@ 800 is an X-Band radar designed for weather detection, groundmapping and avoidance. Weather indications are displayed on the DU-880Navigation Display (ND). Storm intensity levels are displayed in brightcolors contrasted against a deep black background. Areas of heaviestrainfall appear in red, rainfall of medium intensity appears yellow, andareas of weakest rainfall appear in green. After proper evaluation, thepilot can chart his course around these storm areas. The radar can alsobe used for ground mapping. In the MAP mode, prominent landmarks aredisplayed which enables the pilot to identify coastline, hilly ormountainous regions, cities, or even large structures. In ground mappingmode, video levels of increasing reflectivity are displayed as black,cyan, yellow, and magenta.

G. PRIMUS@ 870 Weather Radar System

The PRIMUS@ 870 Weather Radar System consists of the followingcomponents:

. WU-870 Antenna and RCVR/XMTR Unit

. WC-874 Weather Radar Controller

The PRIMUS@ 870 is an X-Band radar designed for weather detection, groundmapping and avoidance. Weather indications are displayed on the DU-880Navigation Display (ND). Storm intensity levels are displayed in brightcolors contrasted against a deep black background. Areas of heaviestrainfall appear in magenta, next heaviest appear in red, rainfall ofmedium intensity appears yellow, and areas of weakest rainfall appear ingreen. Turbulence (TRB) mode is used to detect turbulent air in the 10to 50 NM ranges. Areas of potentially hazardous turbulence are shown ingray white. After proper evaluation, the pilot can chart his coursearound these storm areas. The radar can alsobe used for ground mapping.In the MAP mode, prominent landmarks are displayed which enable the pilotto identify coastline, hilly or mountainous regions, cities, or evenlarge structures. In ground mapping mode, video levels of increasingreflectivity are displayed as black, cyan, yellow, and magenta. A rain

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echo attenuation compensation technique (REACT) mode automaticallyincreases receiver gain as a function of attenuation due to interveningrainfall. At the point where the receiver can no longer detect levelsless than red, a blue field is displayed indicating an out-of-calibrationregion. Target alert (TGT) mode is selected to indicate when level 3(red) or greater weather is present in a sector beyond the currentlydisplayed range. Another feature of the P-870 is automatic tilt whichprovides optimum tilt angle for any selected range.

2. H. FMZ-800 Flight Management System

The FMZ-800 Flight Management System (FMS) consists of the followingcomponents:

. CD-81O Control Display Unit

● NZ-920 Navigation Computers

. DL-800/900 Data Loader

. PZ-800 Performance Computers

. SM-81O Servo (Autothrottle)

The FMS provides lateral and vertical navigation guidance for display andcoupling to the DAFCS. The CD-81O Control Display Unit (CDU) providesthe primary means for pilot interface with the system and displays theselected flight plan data.

The navigation computer can interface with five long range sensors, threevia ARINC 429 buses and two over the ASCB bus. Each navigation computercan also connect to dual Collins Proline 2 or Bendix/King DME receiversand a single VOR receiver. The interface to the IRS, Air Data, EDS, andDAFCS is over the avionics standard communications bus (ASCB). Flightplans are also transferred between navigation computers over the ASCB,while the link to the performance computer and CDU is over an RS-422‘private-line’ interface. To provide high accuracy long rangenavigation, the navigation computer is designed to connect to IRS,Omega/VLF sensors plus VOR/DME. With links to the on-board navigationsensors, the navigation computer develops an FMS position based on ablend or mix of the sensors. The FMS does not directly displaynavigation maps on the CDU; however, the FMS is the source of map datafor other cockpit displays such as EDS. Display of map data is achievedby the utilization of the internal database and ASCB 1/0. A largeportion of the navigation database is subject to updating on a 28-dayinterval. The DL-800/900 Data Loader is used for this purpose.

22-14-00Page 16


The navigation part of the FMS may be considered an area navigationsystem or RNAV. Its fundamental purpose is to provide navigationinformation relative to a selected geographically located point.Navigation management will allow the pilot to define a route from theaircraft present position to any point in the world. The system willoutput advisory information and steering signals to allow the pilot orDAFCS to steer the aircraft along the desired route. Routes are definedfrom the aircraft present position to a destination waypoint via a directgreat circle route or via a series of great circle legs connected byintermediate waypoints.

In addition to providing a lateral steering signal, the navigationcomputer also provides vertical navigation (VNAV). The VNAV modes areVNAV altitude (VALT), VNAV altitude preselect (VASL), VNAV flight levelchange (VFLC), and VNAV vertical path guidance (VPTH). The verticalcommand is proportional to the calculated distance from the desiredvertical path. VNAV allows the pilot to define waypoint altitudes anddescent angles to waypoints and command autopilot to fly the desiredvertical path.

The PZ-800 Performance Computer’s fundamental purpose is to aid the pilotin determining the optimum airspeed/engine setting for his particularflight conditions. In addition, the performance computer functions as anautothrottle computer to directly control the A/C throttles or asguidance to the pilot to optimize thrust management.

The performance portion of the FMZ-800 system has two primary operatingmodes advisory mode only and full active coupled mode. The advisory modecould be considered a flight director in that it advises the pilot onproper airspeeds and engine settings. It is up to the pilot whether hewishes to fly the advisory information or ignore it. The full activemode causes the autopilot and the engine controls to automatically trackthe changing airspeed and throttle advisories. The pilot can reviewother data while the remains coupled to a previous mode.

The computed airspeed and engine settings can be displayed on the EICASdisplay unit or coupled to the SM-81O Servo to control the aircraft’sthrottles to provide automatic tracking of the changing airspeed andengine settings.

In addition to the real time calculations, preflight or flight planningand takeoff calculations are part of the system. “What=if” modes havebeen included so the pilot can determine whether a high altitude is moreappropriate than his current altitude, even if it has a greater headwind.

The joystick is an added feature for entering in a waypoint into theflight plan using a slewable cursor. The new waypoint is defined bylatitude-longitude and can be entered directly into the flight plan.

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2. I. Engine Pressure Ratio System

The engine pressure ratio system is comprised of two engine pressureratio transmitters.

The engine pressure ratio transmitter (EPRT) is a solid-state,microprocessor-controlleddevice. Its primary function is to calculateand output the ratio of engine fan duct pressure divided by engine inlettotal pressure. The EPRT receives fan duct pressure from the on-sideaircraft engine and total pressure (Pt) and calibrated airspeed (CAS)from the on-side AZ-81O Digital Air Data Computer (DADC) over ARINC 429.Data from the cross-side DADC is used when the on-side is invalid. Tocompensate for variations in the actual thrust versus EPR of differentRolls Royce Tay engines, a seven-wire trim plug discrete input isprovided. This allows the engine test people to select a trimappropriate for each engine. The trim plug input is read by the EPRTduring power-up initialization and is used by its software to trim thecomputed EPR. The computed engine pressure ratio is then transmitted tothe DA-884 Data Acquisition Unit (DAU) and to the standby engineinstrument signal conditioner via the two redundant low speed ARINC 429buses. The DAU transmits the EPR data to the SG-884 Symbol Generatorover the ASCB. Calibrated airspeed is used for built-in test only anddoes not affect calculation of EPR.

J. VLF/Omega

The OmegaProcessor

● AT-800● AT-801● AT-803

System (Optional)

Sensor System (0SS) is comprised of the OZ-800 ReceiverUnit and one of the following antennas:

Antenna Coupler Unit, Teardrop H-FieldAntenna Coupler Unit, Brick H-FieldAntenna Coupler Unit, Blade E-Field

The antenna coupler units receive the OMEGA/VLF signals and convert themfor processing by the OZ-800 Receiver Processor Unit (RPU). The RPUreceives the amplified antenna signals and processes them to provideupdated position and velocity information to the navigation computer ofthe FMS. In addition to the antenna signals, the NZ-920 NavigationComputer provides the following input data over an ARINC 429 bus to theRPU:

. Heading and true airspeed

. Initializationdata comprised of LAT/LON, GMT, and the date.

The RPU provides the following output data over an ARINC 429 bus to thenavigation computer.

● Latitude. Longitude. N-S Velocity● E-W Velocity● OMEGA/VLF Station Quality. Estimated Position Error● Status Information● Stations Used● Failure Detection Messages

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For a better understanding of 0SS operation, a description of theOMEGA/VLF network operation and antenna background information is givenin the following paragraphs:

2. J. (1) Omega Navigation Stations

The Omega navigation stations provide a worldwide radio aid tonavigation by transmitting very low frequency radio waves from eightstations scattered around the world. Refer to Table 1.2 for letterand number designation and location of each of the eight stations. A~ilot with an Omega receiver can take advantage of the stableproperties, long-~;waves as a navigat”frequencies 10.2, Ifrequency assignedtracked by the 0SSTo prevent signal “format allows onlya time.

nge, and synchronized format of the Omega radioon aid. Each station transmits four basic3.6, 11 1/3 and 11.05 kHz as well as a uniqueto each station. (This unique frequency can bewith the tunable 13.6 kHz VLF receiver channel).nterference among stations, the transmissionone station to transmit a particular frequency at

(2) VLF Communication Stations

The U.S. Navy operates a worldwide VLF radio communication network.Although not specifically intended as radio-navigation aids, the VLFsignals transmitted by these stations are used by the 0SS tosupplement the Omega station signals. Each VLF station transmits ata different frequency. Table 1.3 lists the locations, frequencies,and transmission strengths of these stations and a similar one inGreat Britain (the term VLF is usually understood to exclude Omega,although the Omega signals are strictly speaking within the VLFband).

22-14-00Page 18.1Apr 15/93


Letter

A

B

c

D

E

F

G

H

~

1

2

3

4

5

6

7

8

Location

Aldra, Norway

Monrovia, Liberia

Haiku, Hawaii, U.S.A.

La Moure, North Dakota, U.S.A.

La Reunion

Golfo Nuevo, Argentina

Australia

Tsushima, Japan

Latitude

66”25’N

06”18’N

21”24’N

46”22’N

20058’S

43”03’s

38029’S

34”37’N

Lonqitude

13”08’E

10”4O’W

157”50’W

98”20’W

55”17’E

65’11’W

146”56’E

129”27’E

Omega StationsTable 1.1

&

1

2

3

4

5

6

7

8

Location

Maine

Japan

Washington

Hawaii

Maryland

Australia

Puerto Rico

Great Britain

Latitude

44”39’N

34”58’N

48°12’N

21”26’N

38”59’N

21049’S

18°23’N

55”22’N

Lonqitude

67”17’W

137”01’E

121”55’W

158”09’W

76”27’W

114”1O’E

67”11’W

1“11’W

Fre~uency

24.0 kHz

17.4 kHz

24.8 kHz

23.4 kHz

21.4 kHz

22.3 kHz

28.5 kHz

16.0 kHz

Power

1026 kW

48 kW

124 kW

588 kW

588 kW

989 kW

100 kW

40 kW

VLF StationsTable 1.2

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2. J. (3) Antenna Background Information

(a)

(b)

There are two basic types of receiving antennas: E-field(electric field) and H-field (magnetic field). Examples ofE-field antennas include long wire, probe and blade; theseantennas respond to the electric-field component of a radiosignal. Such antennas, at OMEGA frequencies, can provideomnidirectional coverage and exhibit good sensitivity.Unfortunately, E-field antennas are also inherently sensitive tothe high electric-fieldsassociated with precipitation-static(P-static). An aircraft flying through clouds containing icecrystals or other precipitation particles can become highlycharged as the result of triboelectric (“frictional”) charging.(Triboelectriccharging occurs whenever two dissimilar materialsare placed in contact and then separated.) The ice crystalsgenerally acquire a positive charge, leaving the aircraft with anegative charge. Corona discharge from some portion of theaircraft then occurs whenever the dc field exceeds a thresholdvalue; each corona discharge pulse results in static.

Loop antennas, if properly shielded and balanced, are sensitiveonly to the magnetic field component (H-field) of anelectromagnetic signal. Loop antennas are therefore lesssusceptible to the electrostatic fields associated with P-staticphenomena. Loop antennas have a directivity pattern, with nullsoccurring along the loop axis. To avoid this null, a secondloop is used at right angles to the first. This configuration isthe familiar “crossed-loopantenna”. Some selection system(switching network) must then be used so the appropriate loopmay (for each OMEGA station at different bearing angles) besequentially employed throughout the OMEGA transmissionsequence. Loop antennas are susceptible to magnetic fieldcomponents produced by the aircraft’s electrical equipment.Strong current impulses produced by engine driven generators,switch closures, relay contacts, transformer saturation effects,etc., can generate magnetic fields with frequency componentsextending into the OMEGA band. These fields are usually fairlylocalized.

NOTE: Some airframe manufacturers don’t recommend theinstallation of H-field antennas.

(c) Finding the best location for installation of theACU is ofparamount importance. The best location for an ACU will varybetween aircraft types and has a high probability of varyingbetween differing configurations of any individual aircrafttype. The most acceptable means of determining the bestlocation for the ACU is to skin map the aircraft.

Skin mapping is selected frequency spectrum survey of practicalantenna locations on an aircraft. By skin mapping, theinstaller determines, over the Omega band, the best location onan aircraft for the installation of ferrite loop ACU.

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Honeywell

I 2. K. LSZ-850 Lightning Sensor System


(Optional)

The LSZ-850 Lightning Sensor System consists of the following components:

. LP-850 Lightning Sensor Processor● AT-850/855 Antenna. LU-860 Lightning Sensor Controller

The lightning sensor system detects lightning activity in the regionaround the aircraft within 100 NM and determines the range and bearing ofeach lightning discharge. A symbol plan display is generated withspecial symbols as shown in Figure 2.1 and these symbols are displayed inconjunction with weather radar displays on the ND.

The LU-860 Li~htnina Sensor Controller contains the switch used for the

I selection of ~he modes of operation.

kWHITE

RATE 1

L. TCZ-91O Traffic Alert

hWHITE kWHITE

RATE 2 RATE 3

Lightning Symbols

BMAGENTA

ALERTAD-13983+5

Figure 2.1

and Collision Avoidance Svstem (Ot)tional)

The TCZ-91O Traffic Alert and Collision Avoof the following components:

● RT-91O TCAS Computer● AT-91O TCAS Antenna

dance System (TCAS) conssts

I TCAS II is an onboard advisory system designed to act as a backup toair traffic control [ATC) radar and the “see and avoid” Procedures.

the

IBy

computing the colsur~ ra~e and altitude of all transponder-equippedaircraft in the surrounding airspace, TCAS II can anticipate a potentialmidair collision much before this has a chance to materialize.

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2. M.

N.

TCAS II continually plot local air traffic on the associated display, andin the event of a conflicting flightpath, guides the pilot toward thecorrect avoidance maneuver. If the intruding aircraft is also equippedwith TCAS, the two systems can communicate their mutual intentionsthrough the Mode S transponders. The coordinated advisories that resultallow the two pilots to execute complimentary avoidance maneuvers.

MLZ-850 Microwave Landing System (Optiona”

The MLZ-850 Microwave Landing System (MLS,components:

● ML-850 MLS Receiver

● CM-850 MLS Control/Display Unit

)

consists of the following

The MLS operates on one of 200 channels between 5031.0 and 5090.7 MHz.The signal format is time multiplexed, that is, each function (azimuth,elevation, basic data, auxiliary data, and back azimuth) is transmittedsequentially on a single carrier frequency. Each function is identifiedby a digitally encoded preamble. The preamble is followed by TO and FROscanning beam signals or more digital data depending on the function.

The ML-850 MLS receiver provides guidance to the azimuth/back azimuth andelevation (glidepath) flight path angles selected on the control unit orautomatically transmitted from the ground station. Guidance is outputfrom the receiver in the form digital deviation signals

Global Positioning System (Optional)

The Global Positioning System consists of the following components:

. Global Positioning System Sensor Unit (GPSSU)

The GPSSU is a two-channel, single-frequency global positioning system(GPS) receiver capable of receiving the L1 (1575.42 MHz) frequencytransmissions from NAVSTAR satellites.

The GPSSU performs the following functions:

. Tracks the L1 coarse acquisition (C/A) code transmitted by the NAVSTARglobal positioning system (GPS) satellites.

c Locks on to the satellite signal.

● Computes the pseudo range from the C/A code. Pseudo range consists ofthe actual range between the satellite and receiver modified byreceiver clock errors.

c Computes the pseudo range rate from the satellite (Doppler). Pseudorange rate consists of the actual range rate modified by receiverclock errors.

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. Decodes the satellite data.

. Computes aircraft position; this is referred to hereafter as thenavigation solution.

. Outputs the aircraft position429 output bus.

c Outputs the aircraft position429 output bus.

The GPSSU interfaces with the fo-

and satellite informationon the ARINC

and satellite

lowing devices

● Global positioning inertial reference units

● Inertial navigation units (INUS).

● Digital air data computers (DADCS).

● Flight management systems (FMSS).

nformation on the ARINC

GPIRUS).

3. Avionics Standard Communications Bus (ASCB) Descri~tion

The ASCB is a communication bus system that allows transmission of data inall directions between subsystems in an aircraft. An example bus system,shown in Figure 3, consists of two pairs of interconnecting wires which formredundant communication paths between subsystems. These two paths are calledBus A and Bus B. Three bus controllers are used to manage all data transferactivity. With dual interconnections and triplex bus controllers, busavailability from the essential to highly essential level is achieved. Thebus controllers reside in each SG-884 Symbol Generator. Only one buscontroller is active at any time. The others act as backup controllers andassume control of the bus when required due to failure of the activecontroller.

Data transfer between users on the bus is controlled by the active buscontroller. Each user is requested by the active bus controller using therequest address of the user to transmit or talk his predefine data messageonto the bus in sequence. During this time, any other users on the bus mayaccept the data message as desired but cannot transmit or talk.

Each user has a defined address or label (user address) and a message, withall transmitted parameters in a defined order. There are no labels onindividual data parameters as in ARINC 429; thereby, bus overhead issignificantly reduced. The only labels in the bus system are the useraddresses, similar to the ARINC 429 equipment identifier. The requestaddresses of the users and the user addresses are listed in Table 2. TheGulfstream IV uses version B ASCB also known as clockless ASCB.

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MAINTENANCE

Honeywell !!4K2!!th.

SG 1Bus

CONTROLLER(ACTIVE)

AFCS 1 t==!

BUS BUSA B

A

\ SG 2BUS

/CONTROLLER(BACKUP)

\

10Hz

IRS140Hz

Fwc 120Hz 4

I

DADC 110Hz

,

tJ

t===

k=+

40Hz

20Hz

10Hz

FSG 3BUS

CONTROLLER(BACKUP)

Example System LJsing the ASCBFigure 3 22-14-00

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I

USER REQUEST ADDRESS USER ADDRESS

IRS 1 82 02IRS 2 83 03DADC 1 06DADC 2 :! 07SG 1 88 08SG 2 8B OBSG 3 8C OcNZ 1 (BASIC) 90 10NZ 1 (BACKGROUND) 92 12NZ 2 (BASIC) 94 14NZ 2 (BACKGROUND) 96 16PZ 1 (BASIC) 98 18AT 1 99 19PZ 1 (BACKGROUND) 9A 1APZ 2 (BASIC) 9C lCAT 2 9D lDPZ 2 (BACKGROUND) 9EFWC 1 A3 ;:FWC 2 A7 27DC 1 AB 2BDC 2 AF 2FFGC 1 BO 30FGC 2 B1 31DAU 1A FO 70DAU lBDAU 2A i; ;;DAU 2B F3 73

ASCB Unit AddressesTable 2

22-14-00Page 21


MAINTENANCE

Honeywell Y&i!#h.The standard bus works on a message basis. Every user transmits his definedmessage on the bus as requested by the controller. Based on the address, allother users can selectively choose to receive the message or ignore it. Thereceiving users must know the order of the data parameters within themessage, just as ARINC 429 users must know the labels of the data they desireto receive. Each data parameter is called a word sequence position (WSP).The word sequence positions start at WSPO, then WSP1, then WSP2, etc. Eachword sequence position contains 16 bits.

Figure 4 illustrates an example of a typical user subsystem. It shows a useraddress defined for the DADC. The defined messaqe content is shown in thebox to the right. Other data in front and in ba~k of the actual data iscontrol and error checking information required in all user messages.

BUS CONTROLLER

{ ‘

REQUESTTRANSMISSION ADDRESS86

DADCNO.1

r,

FLAG RESPONSEADDRESS06

DADCTRANSMISSION

1

PRESSUREALTITUDEBAROALTITUDEALTITUDE RATEINDICATEDAIRSPEEDTRUEAIRSPEEDMACHTOTALAIRTEMPERATURESTATICAIRTEMPERATUREDYNAMIC PRESSURE

CHECKSUMCRCERRORCHECK

~ FLAG AD-15024-R2

Illustrationof a Typical User SubsystemFigure 4

22-14-00Page 22

Aug 15/91Use or disclosure of information on this page is subject to the restrictions on the title page of this document.

Honeywell #&!$~.cEBus controllers operate in active or standby status. The active buscontroller (BC) will detect its own bus control processing faults and removeitself as controller fail-passively. A bus controller will bring both busesto full activity within 150 ms after power is applied and 700 ms for powerinterrupts of greater than 12 ms and less than 200 ms (warm start). If BCNo. 2 sees no activity on both buses after 150 ms from when power is applied,it becomes the active controller. If BC No. 3 sees no activity after 300 ms,it becomes the active controller.

Both Bus A and Bus B are used in the standard bus system. All bus controllerrequests are transmitted simultaneously on both buses. The user subsystems,(LRUS), however, respond with their data on only one bus. User subsystemslisten on both buseson Bus A, right side

Forty times a secondbuses for a group offrame, one bus frametransmit at 40 times

but transmit on only one. Left side subsystems transmitsubsystems transmit on Bus B.

the active bus controller sends out requests on bothusers to transmit data messages. This is called a busevery 25 milliseconds. Some subsystems need notper second; therefore, the complement of subsystems

requested to transmit during a bus frame varies. Some users are asked totransmit every frame, some every other, some every fourth, etc. This allowsupdate rates of 40, 20, 10, and 5 times per second or slower. Figure 3illustrates the operation of the standard bus. System requirements havedictated that a 40-Hz update rate is required for an IRS system, 20 Hz forthe fault warning computer (FWC), and 10 Hz for the FGC and DADC.

The bus controller software is configured to request each subsystem totransmit data at the proper rate. The IRS must transmit every frame, FWCevery other frame, and the FGC and DADC every fourth frame.

Eight frames are defined, with different groups of subsystems transmitting ineach. Table 3 shows the complement of subsystems requested to transmit ineach of eight sequential frames. After frame seven is complete, the sequencerepeats, starting again with frame zero. ASCB applications use eight uniqueframes which repeat, providing update rates of 40, 20, 10, and 5transmissions per second.

22-14-00Page 23


MAINTENANCE

Honeywell W$!#h.

Bus FRAME O FRAME 1 FRAME 2 FRAME 3 FRAME 4 FRAME 5 FRAME 6 FRAME 7

BOTH START START START START START START START STARTBOTH CONTROL CONTROL CONTROL CONTROL CONTROL CONTROL CONTROL CONTROL

BUS A FTIU SG 3 NZl~ PZ l(A)

&

SG 3 NZl~BUS B NZ2~ PZ 2(A) NZ2 ~

BUS A FGC 1 Pz l(P)~ FGC1 Pz l(P) 2BUS B FGC 2 Pz 2(P)~ FGC 2 Pz 2(P)~

BUS A OAOC 1 SG 1 OAOC1 SG 1BUS B OAOC 2 SG 2 OAOC 2 SG 2

BUS A IRS 1 IRS 1 IRS 1 IRS 1 IRS 1 IRS 1 IRS 1 IRS 1BUS B IRS 2 IRS 2 IRS 2 IRS 2 IRS 2 IRS 2 IRS 2 IRS 2

BUS A DAU l(A) OAU l(A) OAU l(A) OAU l(A)BUS B OAU l(B) OAU l(B) OAU l(B) OAU l(B)

BUS A DAU 2(A) DAU 2(A) OAU 2(A) OAU 2(A)BUS B DAU 2(B) OAU 2(B) DAU 2(B) DAU 2(B)

BUS A DC 1 Oc 1 Oc 1 Oc 1BUS B DC 2 DC 2 Oc 2 Oc 2

BUS A FUC 1 WC 1 Fwc 1 FWC 1BUS B FWC 2 FWC 2 FWC 2 FWC 2

BUS A NZl~ NZl~ NZl~ NZl~BUS B NZ2~ NZ2~ NZ2~ NZ2~

BUS A Pz l(P)~ Pz l(P)~ Pz l(P)~ Pz l(P)~BUS B Pz 2(P)~ Pz 2(P)~ Pz 2(P)~ Pz 2(P)~

& Oenotes a.tothrottle data.

~ Denotes background performance data and navigation data.

& Oenotes basic perfcmnsnce data and navigation data.

ASCB Frame Structure Al1owing 40, 20,10 and 5 Hz Update Rates

Table 3

22-14-00Page 24

Aug 15/91Use or disclosure of informationon this page is subject to the restrictions on the title page of Ihls document.

MAINTENANCE

Honeywell &LJLJ!!%M.The bus controller executes a bus frame every 25 milliseconds, 40 times asecond. Two short messages from the bus controller begin each frame, aframe-start message and a control/test message. The frame-start message issimply a ‘wake-up’ call announcing to all users that a new frame is starting.The control/test message is reserved for functions such as identifying theframe number (O through 7 in this example) and controlling maintenance testactivity. Following the control/test message, the bus controller requestsall users to transmit for that particular frame. Figure 5 illustrates atypical bus activity during a frame. Following the start and control/test, arequest for IRS 1 is transmitted on both buses. IRS 1 responds with its dataon Bus A. IRS 2 request is transmitted on both buses. IRS 2 responds withdata on Bus B. This process continues as in Figure 5 until all subsystemshave transmitted their messages. Both buses then go inactive until thebeginning of the next bus frame.

The bus controller repetitively transmits user subsystem requests at theproper times, independent of whether the subsystems actually respond withtheir data messages. User subsystems need not all be in existence on thebus. Requests may be transmitted for subsystems which are optional and notinstalled in a particular application. The bus controller database definesthe length of each user message so that the bus controller may requesttransmission at the proper times, independent of responses.

The ASCB interconnects for the SPZ-8000 DAFCS with EDS are shown in Figure 6.Physical characteristics of the ASCB are listed’below:

. There are two independent ASCBS denoted “An and “B”, each consisting ofone wire pair.

. The ASCB transmission lines shall be Ra.vchem2524E0114 with a thermoradjacket.

● Each ASCB transmission line pair shall125 ohms t 5 ohms. The characteristicpicofarads/foot.

● Each ASCB transmission line pair shall

have a characteristic impedance ofcapacitance shall be 12 i 2

be terminated at its two ends withnoninductive 127-ohm resistors t 1%, 1/4 watt, metal film. The cablelength between the last stub and the termination resistor shall be 24inches.

. The ASCB transmission lines shall have a maximum length betweenterminators of 150 feet.

c Stub lengths at each user pickoff shall not exceed 36 inches. Stubconnections to the main bus shall be accomplished with bus couplers asshown in Volume II, Section 6, InterconnectsTable 501, Figure 3-7.

22-14-00Page 25

Aug 15/91Use or disclosure of reformation on this page is subpcf to the restrictions on the title page of this document,

MAINTENANCEHoneywell $!!!%+.

. The shield connections at each stub shall be accomplished with the buscoupler.

. All bus couplers shall be electrically bonded to the aircraft structure.

c The ASCB transmission lines shall be connected in a daisy chain fashionbetween user subsystems. The cable length between users shall not be lessthan 2 feet.

22-14-00Page 26

Aug 15/91Use or disclosure of information cm this page is subject to the restrictions on the title page of this document.

BUS A ACTIVITY BUS B ACTIVITY

I FRAME START1 FRAME STARTI

TIME

[CONTROUTEST

JI CONTROUTEST I

I REQUESTFW I I REQUEST 171U

lWU MSG. -- -INACTIVE- - -

I REQUEST FGC NO. 1 II REQUEST FGC NO.1 I

FGC NO.1MSG -- - INACTIVE- - -

I REQUEST DADC NO.1 I I REQUEST DADC NO. 1 I

DADC NO.1MSG -- - INACTIVE - - –

I REQUEST IRSNO.1 I I REQUEST IRSNO. 1 I

IRSNO.1MSG -— - INACTIVE- - –

I REQUEST IRSNO.2 I I REQUEST IRSNO. 2 I

--- INACTIVE- – - IRSNO.2 MSG

REQUEST DAU NO. 1(A)IREQUEST DAU NO. 1(A)I

-- - INACTIVE- - -DAU NO. 1(A)MSG

REQUEST DAU NO. 2 (A)

-- - INACTIVE- - -

I REQUEST DAU NO. 2 (A) I

DAU NO. 2 (A)MSG

I REQUEST DC NO. 1 II REQUEST DC NO. 1 I

DC NO. 1MSG -- - INACTIVE- - -

CONTINUEDTO END OF FRAME

CONTINUEDTO END OF FRAME

AD-1 5025-R1

Example of Bus Activity (Frame O)Figure 5 22-14-00

Page 27/28Aug 15/91


c(4mo

1-—.— ——— ——— ——— .—— ——— —.— ——— ——

1

I I 11 IIIIII1IIIIIIIII

IIIIIIIIIIIIIIIII

IIII

II 55A1O

1[DC NO. 2

115J2-A@

I

cl 1S12-AM 1,,.55A11

n

I

LOAD55A47

F“ “——— ——— ——— ——

-r——— ——— —.— ———

F1

IIIIIIII

n55A33

H

55A3a

55A14

1

LEFTTEST PANEL

II

55A9I

~

IIPZ NO, 1

122J1 6-4/5

I55A0

121J1B-26131

IIIIIIIII

iI

IIIIIIIIIIIIIIIL

55A29 C121J1A-l W1l II55A21

II

55A20

IIEl C9Jl B-13/14 55A19

I 55A6 I ~ AZNO. 1

9J1B-13114 55A41 II

LEFTAVIONICS

BAY I

RIGHTAVIONICS

BAY

I 55A4

134Jl B-61/62

IIIiII

IIIIIIIIIIII

I 55A26

55A25 13711 B-4/6

HF -

DA NO. 2 II

i S.7J1/$+6 55Ai7

55A50

H

55A23

I 55A1I

~

SG NO. 1

55A51

BC NO. 1

SG NO. 1 u

65J1B-16117

I 55.4+/// 4 cl JIA-l H/2H

IRU NO 2 II

.—— — ——— —— ——— — ——— ——— —— ——— —— J

1-

,

I 55A21+

lJIA-l H/2H

IRU NO. 1

lJIA-ltu2K 55A37 II I # J

55A13

IIEIJIA-1W2K

55A46

II

SG NO. 3

EMJIB-1w17

SC NO. 3 H

E65J1 B-46147 55A45

~BUS A

~BUs B AD-l&344

SECTION 2COMPONENT DESCRIPTION

1. General

This section provides an illustration, leading particulars, a briefdescri~tion, and a block diaqram or schematic of each component used in theSystem-. The information is ~nly for the specific standard componentsin Section 1, Table 1.

The components are separated into the following subsystems:

Subsvstem Paraqrat)h

LASEREF@ II Inertial Reference System (IRS)

ADZ-81O Air Data System

AA-300 Radio Altimeter System

Reserved Subsystem Not Applicable to G-IV Aircraft

EDZ-884 Electronic Display System

DFZ-820 Dual Flight Guidance System

PRIMUS@ 800 Weather Radar System

PRIMUS@ 870 Weather Radar System

FMZ-800 Flight Management System (FMS)

Engine Pressure Ratio (EPR) System

Optional VLF/Omega System

Optional LSZ-850 Lightning Sensor System

Optional TCZ-91O Traffic Alert and CollisionAvoidance System (TCAS)

Optional MLZ-850 Microwave Landing System (MLS)

Optional Global Positioning System (GPS)

2

3

4

5

6

7

8

8.1

9

10

11

12

13

14

15

NOTE- Refer to the contents section in the front of this manual for a‘“ listing of each component contained in each subsystem.

1isteal

Page 101Apr 15/93


2. LASEREF’”II Inertial Reference Svstem [IRS)

The Gulfstream IV factory-installed inertial reference system includes thefollowing:

. Two inertial reference units (IRUS)

. One attitude and heading reference unit (AHRU)

. One triple-channel mode select unit (MSU)

. One inertial system display unit (IDSU)

At the completion center, the customer has the option to replace the AHRUwith a third IRU, and the ISDU with a LASERTRAK’”. The IRS providespositions, rates, and accelerations in all three body axis (pitch, roll, andyaw), ground track and speed, and wind speed and direction to the rest of theSPZ-8000 system.

A. Inertial Reference Unit (See Figures 2-1 and 2-2, and Tables 2-1 through2-4).

The inertial reference unit (IRU) is the main electronic assembly of theIRS. The IRU contains an inertial sensor assembly, microprocessors,power supplies, and aircraft electronic interfaces.

Accelerometers and laser gyros in the inertial sensor assembly measureaccelerations and angular rates of the aircraft.

FAULT BALL INDICATOR

\

(USEDTOfilTIATE~ ‘ ~ ,TEST MODE)

1 /

AD-1 S478

Inertial Reference UnitFigure 2-1 22-14-00

Page 102Aug 15/91

Use or disclosure of information on this page is subject 10 the restrictions on the title page of this document.

Dimensions (maximum):

Length ...................................... 12.76 in. (324.1 mm)Width ....................................... 12.70 in. (322.6 mm)Height ....................................... 7.62 in. (193.5 mm)

Weight (maximum) .................................. 45.0 lb (21.1 kg)

Power Requirements:

IRU on AC ................................ 115 V, 400 Hz, 135 W MaxDC required when operating on AC ............... 28Vdc, 1.1 WMax

IRUonDC ...................................... 28 V dc, 116W Max

Mating Connector ............................. Cannon BKAD2-313-30001

Mounting ................................... Mounting Rack, HoneywellPart No. 10088502-1OX

Inertial Reference UnitLeading Particulars

Table 2-1

The IRU microprocessors perform computations required for:

● Primary attitude. Present position. Inertial velocity vectors. Magnetic and true north reference. Sensor systematic error compensation

The power supplies receive ac and dc power from the aircraft and backupbattery, supply power to the IRS, and provide switching to primary ac,primary dc, or backup battery power.

The aircraft electronic interfaces convert ARINC and ASCB inputs for useby the IRS. The electronic interfaces also provide IRS outputs in ARINCand ASCB formats for use by the associated aircraft equipment.

A fault ball indicator and a manual INTERFACE TEST switch are mounted onthe front of the IRU and are visible when the IRU is mounted in anavionics rack.

In a triple-IRS installation, IRU 3 provides IRS comparison monitoringfor the flight guidance system and also provides the FMS with a thirdlong range sensor.

The input and output information provided by the IRU is listed inTable 2-2. 22-14-00

Page 103Apr 15/93


Honeywell !%!!fb.”

Pin Name Remarks

JIA-C1l(H) Primary ASCBJIA-C12(L) Clock

JIA-C14(H) Primary ASCBJIA-C15(L) Data

JIA-H1(H) Primary ASCBJIA-H2(L) Data without

line resistance(external buscouplers)

JIA-E13 ASCB Data FieldSelect

JIA-E14 ASCB 2/4 WireSelect

JIA-F1l(H) Secondary ASCBJIA-F12(L) Clock

JIA-F14(H) Secondary ASCBJIA-F15(L) Data

JIA-K1(H) Secondary ASCBJIA-K2(L) Data without

line resistance(external buscouplers)

The IRU transmits and receives with thisport. Pins JIA-H1 and H2 are used whenthe current limiting (short circuitprotection) resistors are placedexternally, as in the G-IV where theyare placed in the bus coupler. Refer toTable 2-3 for 38words/transmit datatransmitted.

The IRU is capable of transmittingeither 23 words/transmission or 38words/transmission based upon thisdiscrete. Ground this pin (short toJIA-A1) for the 38-word transmission.G-IV uses the 38-word transmission.

This discrete instructs the IRU whetherthe installation uses a clock bus ornot. Ground this pin (short to JIA-Al)for clockless. G-IV does not use clockbus.

The IRU receives with this port. PinsJIA-K1 and K2 are used when the currentlimiting resistors are placedexternally. The LRUS received by theprimary and secondary ports are: buscontrollers 1, 2, and 3, FGCS 1 and 2,and IRUS 1 and 2.

NOTE: The FGCS and the IRUS are onlyreceived by IRU No. 3.

Inertial Reference UnitInput/Output Information.

Table 2-2 22-14-00Page 104


HoneywellMAINTENANCEMANUALGULFSTREAMIV

Pin Name Remarks

JIB-G7(H) ARINC 429 \

JIB-G8(L) Output No. 1

JIB-E5(H) ARINC 429JIB-E6(L) Output No. 2

JIB-K12(H) ARINC 429 ,JIB-K13(L) Output No. 3

JIB-F14(H) ARINC 429JIB-F15(L) Output No. 4

JIB-G14(H) ARINC 429JIB-G15(L) Output No. 5

JIB-H14(H) ARINC 429JIB-H15(L) Output !Jo. 6 >

JIB-A8(H) FMS No. 1 ARINC 429’JIB-A9(L) Initialization

Input

JIB-C5(H) FMS tie. 2 ARINC 429JIB-C6(L) Initialization

Input

JIB-A13(H) ISDUARINC 429JIB-A14(L) Initialization

Input

JIB-K5(H) ADCNO. 1 ARINCJIB-K6(L) 429/575 Input

J1B-J1O(H) ADCNO. 2 ARINCJIB-J1l(L) 429/575 Input

1

JIB-J6 ARINC 575/429ADC select

JIB-F1 Mode Discrete No. 1

JIB-F2 Mode Discrete No. 2}

Refer to Table 2-4 for the data that istransmitted on these identical buses.Bit rate = 100 kbits/sec.

The IRU reads the following data fromthese three ports:label 041- set latitudelabel 042- set longitudelabel 043- set heading

Labels 041 and 042 are read for Navinitialization. Label 043 is readto set heading in attitude mode.

The following data is read from thesetwo ports:label 203- pressure altitudelabel 204- bare-corrected altitudelabel 210- true airspeedlabel 212- altitude rate

Gnd (short to JIB-A1) = ARINC 429Open =ARINC 575 (419)

~ Alicm ~ ~Fl: open grid. gnd openF2: open open gnd gnd

Inertial Reference UnitInput/Output Information

Table”2-2 (cent) 22-14-00Page 105

Aug 15/91Use or disclosure of Wormation on this page is subject to the restrictions on the title page of this document.

Honeywell #$(!!ib.cE

Pin Name Remarks

J1B-A1O

JIB-A2

JIB-A3

JIB-J9

JIB-A7

JIA-J9

JIA-A7

JIB-F3JIB-E3JIB-D2JIB-D3JIB-E1JIB-A15JIB-J15

JIA-E6

JIA-E7

JIB-G1

JIA-G9

Remote Test

IRU Orient 1

IRU Orient 2I

SDI No. 2

SDI No. 3t

SDI fro. 4 JAlign AnnunciatorNav Ready AnnunciatorWarn Fault AnnunciatorAttitude AnnunciatorNo Air AnnunciatorBatt Fail AnnunciatorOn Batt Annunciator

Miscompare No. 1

Miscompare No. 2 }

IRU Valid

Charger Inhibit

Gnd = IRU output self-testOpen = normal operation

Handles Handles Handles HandlesAft Forward Left Riqht

A2: open gnd open gndA3: open open gnd gnd

IRU 1 IRU 2 IRU 3B-J9 open gnd gndB-A7 gnd open gndA-J9 open open gndA-A7 open open gnd

1Output logic:

Valid condition- GndInvalid condition- open

1Load capacity: 250 mA

These outputs are only active on IRUNo. 3 or the AHRS. The third unitcompares itself with IRUS No. 1 and 2via ASCB. Aground on JIA-E6 =miscompare between IRUS 1 and 3. Aground on JIA-E7 = miscompare betweenIRUS 2 and 3. The following are theeight parameters which are compared:pitch and roll angles; pitch, roll andyaw rates; and longitudinal, lateral andnormal accelerations. The thresholdlevels for these paramete s are:An les 23deg.,

!1[

Rates a 0.3 + R1 +R2 /64 deg/sec Accel. z 0.1 + AA1 +

AA 1/6 Ig

Open = invalid28 V dc = valid (0.25 A max)

Inhibits the battery charger during thepower-up battery test.Open = charger not inhibited28Vdc= charger inhibit (0.25 A max)

Inertial Reference UnitInput/Output Information

Table 2-2 (cent) 22-14-00Page 106

Aug 15/91use or disclosure of information on this page is subject to the restrictions on the title page of this document.

NN

d.

r+

HSP BIT BIT FUNCTION NOTE FORMAT SCALE RSB APPROXRESOL/LSB POS SENSE

01

FTIU SIWAR--- ------ I;;[;-;[;~----------------l----ly;~;-------------------l--------------------l---l-----------------l-----------------1----1--------1

,1I..................................II----.........................I--------------------l---l---”-------------l-----------------l----l--------ISENSORCONTROL

1 15PACKEDLOGIC

TEST LOGIC 1 = TEST1 14 VAL10 LOGIC 1 ● VALID

6470 FASIRSTL

1 13-1164C0 FASIRSVL

SENSORTYPE 000 ● IRS1I

001 = ANRS010 = OMEGA

11

011 - LORAN100 ■ GPS101 - 111 ● SPARE

i IO-B COUNTER O-7 HEX1 7-0 SENSORAUORESS 02 H - LEFT 03 H - RIGNT 04 II - CENTER

i l------”------------------lIT”0,5~o”pLEMENT......... 1----.............................................l;;-l;:;;;;;--------l;;;-~;---------I;;;;l;i;;i;;;l2 15-0 SIN PITCtl ANGLE +/- 1*O

II II... ...... ......................... ---- -------------------------I 116 10,W00305 l-----------------m--------l.---------------------------------------3 15-0 COS PITCH ANGLE TNO’S COMPLEMENT +/- 1.0 4/- 90 OEG RANGE 6474 RIRCOSTL

II 1----164C4

... ...... ......................... I---.-.-..-.---..---..--0---------------------4 15-0 SIN ROLL ANGLE TuO’S COMPLEMENT

l;;-1;-~;~;&--------l-----------------l----l--------1+1-1.0 . RIGIIT MING 00WN 6476 RIRSINPL

II l----l~;~~;[;;;;---------l;;"";:;------------l;;-lo 00M305 I64C6

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .5 15-0 COS ROLL ANGLE

----------------- . . . . . . . . . . . . . . . . . 1----1--------1. +/-90 OEG RANGE 6478 RIRCOSPL

II-.....................-.....*-----i-”--l:’-i;i;;[”-------------”-l6 15-1 TRUE HEAOING

;---;;~;~li;------li;-l;;;~;;”;-;;;~;-”-l;;;i--”----------l;Al;iiiGKl60

.FLAG LOGIC 1 = VALIO

.

II..- ..-. -.7 15-0

i;;;;i;~;[;;~;;~-------- l----l;i~l;-------------------l ;-;;;;;;-:-;;;;-;;--l;;-l;-;-;~--------- lj;-------------.-l:l:~.[.

II-.. . . . . . . -------"-----"-----"-----l----lfi~;;-;;fi[;~~-------"l+,-goOEG815-1 PITCHANGLE

11510~549,000275 li;ii-i;----------l:%-------l...................---------------------

BO FLAG. . 647E RIRTHETL

I l-"-----------------------l----l::i;i;i~::::-----.--..i--- .- Q---915-1 ROLL ANGLE

;;--i;;-;;;--------- l;;-; ;;;;;;;-;;;;;--- l-----------------l::::l:~::::::I90

. RIGHT UING DOWN 64B0 RIRPN1OLFLAG

. .

II-.. . . . . . . ----"--------------------l----l:::i;;;;~~~:f-----"----l--------------------l---l-------------...-[.-.-.--.--...--.-l!:::[:!!!!:!:l10 15-1 MAGNETICNEAO1NG 0-360 DEG 15 0.0109/0.00549100 FLAG LOGIC 1 = VALIO

EAST 6482 RIRPSIML

II..- ------ . . . . . . . . . . . . . . . . . . . . . . . . . l----l~;;;~[~~;~~l~---------l11 15-1 INERTIAL VERTICAL SPEED

;;--;;;;-~;;fi;----l;;-li-;i-i---------liiiii---------.--l::lR:-l

110.

FLAG LOGIC 1 = VALID● O

II..................................11..-*.........................112 15-4 BOOYPITCH RATE TUO*S CONPLEHENT

;;--;;-;;;;~;------l;;- l;-;;;;;;;-;;;;;--l ;;;-;;---------- l:::y~:!:::$

12 3-1 SPARE. . . 6486 RiRqOL

120 FLAG LOGIC 1 = FLAGII... ...... ------------------------”l----li~;;”;;;;[;~~i---------I

13 15-4 BOOYROLL RATE;;--;;-;~;;;;;------l---l----------------- l-----------------l:~::l~~::::::l

133-1 SPARE. 12 0.01221/0.00076 RIGIIT MING 00MN 648B RIRPOL

130 FLAG LOGIC 1 ■ FLAGII-.. . . . . . . ------------------------- l----l ------------------------- l--------------------l---[-----------------l-----------------1:!:!1:!:!::!:1

A

(D

0m*w

NNI

H3(D

2

i%i?2?)(D

4.

l-t

HSPBIT BITFUNCTION NOTEFORHAT SCALE RSB APPROXRESOL/LSB POS SENSE

I l---------"-------"-------l----lf~~;-~~~~---------l+,-25~EG,~Ec I iFTIUSIMVAR

... ...... -----------------------.................I-----------------l;;;~l~i;~;~--14 15-4 BOO:YY~URATE 12 0,01221/0.00076IIIGNTTURN143-1 64DA140 FLAG LOGIC1 ■ FLAGII

FASYBROLI l~;;;-;;;;~;~;--------l+,2*0G,~-------------------------------------- II-----------------------.................I---------------”-l;;~l;i;;~--

15 15-4 LONGITUDINALACCEL 12 0,000976/0.000061FORHARO153-1 SPARE-. ....-ii b- FLAG LOGIC 1 = VALID

II II I

640C FASNXL..- ..--.. 112 10~0976,00000611 R1GI,T---... ---.. -.*----------- e--- ------------------------- -------------------- --- 1648EIR,RNYL I

16 15-4 LATERALACCEL

----------------- ----------------- ..-. -.-.--.-TUO’S COMPLEMENT +/- 2.0 G’S

163-1●

SPAUE.

160 FLAG LOGIC 1 ● VALIO

II I~ lT”o,~COHPLEM”T I640EFASNYL

... ...... 11210~244,0~0152l;-------”-------l&-&&;-l----------------------------------------------------------------------------------------------17 15-4 NORNALACCEL . +/-5.0G’s173-1 SPARE

. .

170 FLAGII

LOGIC1 ● VALIDII

64E0FASNZLI... .-----.--..-0---------------------------------------------------------------------------------------.................

18 15-1 GROUNDSPEEOII I l;&lfi~~&--l

TUO’SCOMPLEMENT +/- 1029 KNOTS 15 0.06248/0.03124FORUARD180 FLAG LOGIC1 = VALIOII II

64E2FkNZL-..------ 1------------------------------------------------------.--------------------------------------------------------:;~5-1 TR~fA~RACKANGLE

115100,09,000549 IEA5TSEMICIRLE 0-360 DEG

l;;~l;i~fil. ●

IILOGIC1 = VALID

I lfi;;;=~~~--------l+,-goOEG64E4FASTTOL

..-------......................... ---- 1,2IOC04394,0*002746Iup I 1--------1-------------------------------------------------------------2015-4 FLIGNTPATNANGLE 6496RIRGAMAL203-1 SPARE200 FLAGII

LOGIC1 ● VALID..-.-----...............*---------21 15-3 VEW&L ACCEL

l-;:l~;-;~;~;~~---”----I;;--;;;;;---------l;-l;:&;;;&i;;IK--------”------I;81::- I.212-1210 FLAG LOGIC1 = VALIDII II

64E8FASHDDL---...... ........................-.... -------------------------I--------------------li;-l;-~;;;;~;;l;;~~;---------l~i;~l~i~;;;;;l22 15-3 ALONGTRACKACCEL TW’S COMPLEMENT +/-4.0G’S222-1 SPARE

. .

220 FLAGII

LOGtC1 ■ VALID..-.-...- 11........................------.-------................23 15-3 CROSSTRACKACCEL

ITUO’SCOMPLEMENT

;;:-~;;;;;---------li;-l;~;~;;;;;l;iizii-------..-l~l::: I

232-1 SPARE. .

230- FLAG-11

LOGIC1 = VALIDII

64ECFASTRKYL... ...... ......................... -----------------------------24 15-4 TRM&NGLE RATE

I 1,2looo156,0*mo9761-----------------------.................TUO’S COMPLEMENT

-----------------l;i;;l;i;~;;;[l24 3-1

+/-32 DEG/SEC CLOCKt41SE

24 0 FLAG LOGIC 1 = VALIDII II..-------------------------------------------------------------I;;--;:;-;;;---------l;;”1;-;;;;$-;;;;;Iiii;;r--..----l:::l::I

25 15-3 FLIGHTPATNACCEL252-1

TNO’SCOMPLEMENTSPARE

. .

250 FLAGII II

64F0FASNAFPL..-------......................... ------------------------------------------------------------------------------------------.--..*--1 II I Ill

NSPBIT 911FUNCTION NOTEFORNAT SCALE RSB APPROXRESOL/LSBPOSSENSE FTIUS114VARI l"--"--------------"------l----lfi~~~-~~~~~~~--------[i;~i~~-~~-~~-----l~i-l~~~~i~i-~~~-----l~~~~-"---------l~~~~l~~~~~~j-..------

;: ;5i3 PPOSLATITUDE(ISW)PAD LOGIC ZEROS

260- PPOS LAT1TUDE FLAG LOGIC 1 = VALID 611F2RIRLATDL

II..- ------ II. . . . . . . . . . . . . . . . --------- ---- ------------------------- I ;:;~;~-p--l;i-l;;;i;i-;;;----- l;~------l;;l”--l;;l--------I27 15-11 PPOSLONGITUDE(lsb) TWO’S COMPLEMENT -;; ~0-9 PAD

.LOGICZEROS 64F4

PPOSLONGI TUOE VALID LOGIC 1 ● VALID FASLONVL

2? 7 PPOS LAT1TUDE SIGN 1 . SOIJTII O = NORTH27 6-O PPOS LATI lLN)E (msb) RIRLONDL

I l~~~~-~~~~~fi~-~i~~----[ ----l * . ~EST 0. EAST 1--- ------ I l-----------------l-----------------l;;&l-------------------------------------------------------281528 14-0 PPOSLONGITUOE(msw) 64F6II---------;-y;;i~i;;-------------l'---l~~~~~-~&~i~~~{i---------~~;I-i~~~-~~~~------li~i~~~;~Tfi~---l~~~~~l~~~~l----------l~~~~l~i~~~il

29 16-1290 “FLAG LOGIC1 ● VALIDII

64fBFASEDOTLI lfi~~-~&~;~fl-------l~~i;&-;~(-----1i;-t~;~;~-~~----"--1~o~T}l-------------.-.---.-.-.--.--.---0----

:];5-1 N-:L~~LDCiTV164MIR,RWOTLI-----------------..-.--------

.0LOGIC1 = VALID

II II64FAFASNOOTL

-----------------------------------------------------------------------------------31 15-0 SPARE

I II I 164ACI I---------------------.---..-*------------..-.----

64FCII II---------------------------------------------------------------

32 15-0 SPAREI--------------------1I I--------------------................-1--;;1--------1

II II64FE

---------------------------------- I II---------------------------------------------------------------------.................I33

l;&l--------loISCRETESIRS)

3315 1mig sT TUS(LsB)33 14

6500IIRMISL

3313 ADCOR IRUFAULT LOGIC1 = FAULT FASAOIFL33 12 EXCESS[VEMOTION LOGIC1 = EXCESSIVENOTION FASXNOTL33 11 NO IRSINIT1ALIZATIOH LOGIC1 = NO INITIALIZATION FASNIRSL33 10 ALIGNFAULT LoGIC1 ■ FAULT FASALIFL339 BATFAILON BAT LOGIC1 ■ FAIL FASDCFOL338 IRUFAULT LoGIC1 = IRUFAULT FASIRUFL33 1 MC FA~LT LOGIC1 ● ADCINVALID FASAOCFL336 ON BAT LOGIC1 = ON BAT FASONBAL335 BATFAIL LOGICI = BATFAIL FASOCFAL334 ATTITUDEINVALID LOGIC1 = ATTITUDENOTVALIO FASATTIL333 SETHEADING LOGIC1 = SETHEADINGHASBEENINPUT FASSNEOL332 NAVNODE LOGIC1 ● NAV FASNAVL331 ATTITUDEMoDE LOGIC1 ● ATTITUDE FASATTML330 ALIGNMOOE/NOTREADY LoGIC1 = ALIGN FASALILII II---..--...-----..+-........-.----------------------------------.................... *.-................-................. ------------1 II I Ill

USPBIT BIT FUNCTION NOTEFOIMAT SCALEII ;i;~;~;-ii~;;---------”l----l I

RSB APPROXRESOIJLSO POSSENSE FTIUSIWAR... ------ II----------------------------------------------................... .................I I;;;;i--------l3434 15-7 SPARE 6502346 HIGIILATTITUDEALIGN LOGIC1 = ALIGN FASHLAOLi45 HIGHLATIlUOEOPERATION LOGIC1 ● OPEIIATION FASHLML344 IllsCOWING LOGIC1 = IRSCOOLING FASICOLL343 ON DC LOGIC1= ONOC FASONOCL342 NAvROY ~OG~ME#● NAVRDY FASNAVRI.34 1-0 ALIGNSTATUS(BIT1 ■ MSB)It II

o-7MIN 3 1 HIM IIRALISL---..-..- -----------------------------------------------------------------------------I I l-----------------l-----------------l;;;;l-----*35 MAINTENANCETEST35 1536 14351335 1235 1135 1035 935 835 735 635 535 435 335 235 135.0

SPARESPAREA/DOR HIMFAuLTCPUFAULTDISCRETEl/OFAULTHOTFAULTSYSTENTESTSTEHPSENSORFAULTGYROFREQANDBIASFAULTGrKOFAULTACCELRESIDUALFAULTSPAREMEIIN)RYFAULTSENSORLSICFAULT

6504

LOGIC1 ■ FAULTLOGIC1 = FAULTLOGIC1 = FAULTLOGIC1 = FfiULTLOGIC1 = FAULTLOGIC1 ● FAULTLOGIC1 ● FAULTLOGIC1 ■ FAULTLOGIC1 = FAULT

LOGIC1 ● FAULTLOGIC1 ■ FAULT

FASADWLFASCPUFLFASDIOFLFASUINFLFASYTSTLFASTSFAI.FASGFUFLFASGWFLFASACRFL

FASMEWLFASSLSFL

SPAREPONERSUPPLYFAULT LUG[C1 ● FAULT

II IIFASPSUFL.-.------......................... ----......................... ----------------------------------------1 II

36 CHECKSUN SUHOF UORDSI-----------------1;;;;1--------1

II I ICRC6506..-...... ------------------------- -----.----..--0---------------.-.-------..--0----I

3? ERRORCIIECKl---l-----------------l-“--------”------1;;;;1--------1

II.-.------ 11 I6508-------------------------------*.--.----.*--------------------------------... -----------------II I

3tl IXILCFLAIi 7EHEx 111-----------------------------

II..-..-.-.-------------------------II I-------------------------------------------------l---l-----------------l-----------------I----l--------I

NOTES:

+1.0G HHILESTATICk 0.0G UHILESTATIC

MAINTENANCE

Honeywell !i’lN!#h.

DATA WORD

Time To NAVPOSLatitude”POS Longitude”

Ground Speed’TK Angle True”

MAG HDGWhd Speed”Whd Angle”True HDGImagrated Vart Accel

IRS DiscretesPOS Latimde”POS Longituda”Ground Speed”TK Angle True”

True HDGWind Speed”Wind Direction T“TK Angle (MAG~Magnatic !4flG

Drift Angle”F~@htPath Angle”Flight Path Accel”Pitch AnglaRoll Angle

Body Pitch RateBody Roll RateBody Yaw RateBody Long AccelBody Lateral Accel

8ody Nwrnal AccelPlatform HOG”TK Angle Rate”Pitch AIT RateRoll ATT Rate

IRS Maint Diec$eteTime To NAVCycla CounterPotential Vart Spae&Inertial Altitude”

Afong Tk Hrz Accel”Cross TK HRZ Accel”Vert AccelInertial Ven SpeedN-S Velocity”

E-W Velo~Body Normal Accal”Equipment ID

OCTALLABEL

007010011

012013

014015016044265

270310311312313

314316316317320

321322323324325

326327330331332

&335336337

350351354360361

362363364365366

367370371

FORMAT

BCD6CDBco

8CDBCD

BCDBCDBCD6CDBNR

DISBNRBNR8NRBNR

BNRBNRBNRBNRBNR

BNRBNRBNRBNRBNR

BNRBNRBNRBNRBNR

BNR6NRBNRBNRBNR

DISBCDBNRBNRBNR

BNR8NRBNRBNRBNR

BNR6NR01s

SIGNIFBIT

2

:

44

433420

G202020

2020202020

;;

1;20

15

;:1515

;:201515

—

221

;:

1515152020

2015—

DIGITALRANGE

o to 9.990S to BON1BOE to180Wo to 40950 to 359.9

0-359.90 to 2550 to 3590 to 359.9* 256

—* 0.5*loo to 4085*1.O

* 1.00 to 255*1.O*1.O* 1.0

*0.5t 0.5k4* 0.5* 1.0

* 128f 126k 128*4*4

*4*1.O*32* 126& 126

—

0- 9.90-2097151* 32766k 131072

24*4*4~ 32768* 4096

* 4096&B—

RESO-LUTION

0.10.10.1

1.00.1

0.11.01.00.12A.4E-4

—9.54E-79.54E-70.00399.54E-7

9.54E-72.44E-4B.54E-79.54s-79.54E-7

9.54E-79.54E-71.22 E-49.54E-79.54 E-7

0.00390.00390.00391.22E-41.22E-4

1.22E-49.54E-73.05E-50.00390.0039

—

0.1

; .00.125

1.22E-41,22E-41.22E-40.031250.0039

0.00391.22E-4—

Ms8WEIGHT

———

——

————128

:50.520480.5

0.512B0.50.50.5

0.50.520.50.5

64646422

20.5

::64

——10485761638465536

222163642048

20484—

“ Not provided as AHRU ouIpuIs. The AHRU aignais are output wiIh the SSMS set to failure warning (FW),

UNITS

MinDegfMinLlagfbfin

Ktaw

wKtsDagDagFtfSac

—PiradsPiradsKtsPirads

PiradsKtsPiredsPiredsPirsds

PiradsPksdsG“sPiradsPirads

Dag/sDeg/SDeg/SGsGs

GsPirsdsDegJsOagls%fIs

—

MlnCountFtfhlinFt

G“sG’sGsFt/MinKts

KtsGs—

POSITIVESENSE

Alwavs DOS

North’East

Atwaya posDW from N

:W from NAlways posCW from NCW from Nup

—NorthEastAlwaya posCW From N

CW from NAlways posCW from NCW from NCW from N

RrupForwardupRrWtng On

upRtwing dnNose RtForwardRs

upCW from NCwupR! wing dn

—Always poaAlways posupup

FonvardRtwupNorth

Eastup—

TRANSFERRATE(SPS)

22

222250

2510

E25

5050505050

10505050

5050502510

1052

Inertial Reference UnitARINC 429 Output Data

Table 2-4 22-14-00Page 111/112

Aug 15/91Use or disclosure of mformatlon on this page IS sub@ to the restrictions on the title page of this document

F ——— ——— ——— ——— ——— -11A

1

:11

:12

;14

:15

HI

H2

JIAI

F11

F12b IDH

}

CLOCK

L{

Ii

CLOCK

L

{

H

DATA

L

1{

DATAWITHOUT HLINERESISTANCE L

ASCBPRIMARY(TRANSMITANDRECEIVE)

F14

F15

B?

IIK1

K2

H

1DATA

ASCBSECONDARY(RECEIVE ONLY) II

‘%a H

}

DATAWITHOUT

L %%TANCE

ASCB DATA FIELDSELECT

H

}

NO.

L

H

}

NO.

L

H

}

NO.

L

H

}

NO,

L

H

}

NO.

L

H

}

NO.

L

ASCB 2/4 WIRESELECT

{

H

FMS NO. 1

L

{

H

FMS NO. 2

L

,{

H

ISDU

L

{

H

AOC NO. 1

L

{

H

AOC NO. 2

L

lD-lD’-ARINC 423

INITIALIZATION12.5 KHz

HIGHSPEEDARINC113 —

414 —

K4

K5

110 —

Ill —

429100KHz

ARINC 575/42912.5 KHz

DIGITALPROCESSING

JL-

E6 I MISCOMPARE NO. 1

ARINC 575/426ADC SELECT

MODE DISCRETE NO. 1

MODE DISCR=E NO. 2

REMOTE TEST

IRu ORIENT ~

IRU ORIENT 2

SDI ‘

SDI 2

JE7 MIBCOMPARE NO. 2

IIB

1F3 ALIGN ANN

E3 NAV READY ANN

D2 ] WARN FAULT ANN

IEl NO AIR ANN

A7 I415 E?AIT FAIL ANN

)15 ] ON BAT7 ANN

JG1 IRS VALID

sD141A7~ IIA

7G9 CHARGER INHIBIT

JAl PROGRAMPIN GND

.

{

H115 V AC 4WHZ

L

28 V DC ESSENTIAL

28 V DC BACKUP

24 V DC BAITERY

DC GROUNDVARIABLEANNUNCIATORPOWER IN

CHASSIS GROUND

IIB

1Al PROGRAM PIN GND

Ilc

7:6 28 V DC ISDU POWER

J:9 ANNUNCIATOR POWER

-Pcl 1 I=——— ——— ——— ——— ——— J AD-30734#

Honeywell !!!!!kfr.c’

This page intentionally left blank.

22-14-00Page 115

Aug 15/91Use or disclosure Of information on this page is subject to the restrictions on the title page of this document.

Honeywell !!!!!!!ib.cE2. B. Mode Select Unit (See Figures 2-3 and 2-4, and Tables 2-5 and 2-6.)

AD-16472

Mode Select UnitFigure 2-3


Length ........................................ 3.02 in. (76.7 mm)Width ........................................ 5.75 in. (146.0 mm)Height ....................................... 3.375 in. (85.7 mm)

Weight (maximum) .................................... 1.5 lb (0.7 kg)

Power Requirements ..................... 5 V ac and 5 V dc, 2.5WMax28 V dc, 8.3 W Max each channel

Mating Connectors:

J1 .................................................J2 ................................................J3 ................................................

MS3126F18-32SMS3126F18-32SWMS3126F18-32SY

Mounting ......................................... Unit Dzus Fastener

Mode Select UnitLeading Particulars

Table 2-5

22-14-00Page 116


MAINTENANCE

Honeywell W!!#kk.The mode select unit (MSU) provides IRU mode selection and statusindication for three IRUS, and includes a TEST switch for the IRU. Thefollowing paragraphs describe the MSU operation.

2. B. (1) Selection of Basic Modes and Submodes

Modes are selected by setting the MSU mode select switch as follows:

Q OFF-TO-ALIGN - The IRU enters the power-on/built-in testequipment (BITE) submode. When BITE is complete afterapproximately 13 seconds, the IRU enters the alignment mode. TheIRU remains in the alignment mode until the mode select switch isset to OFF, NAV, or ATT. The NAV RDY annunciator lights uponcompletion of alignment.

● OFF-TO-NAV - The IRU enters the power-on/BITE submode, thealignment mode, and upon completion of alignment, the navigatemode.

Q ALIGN-TO-NAV -alignment mode

● NAV-TO-ALIGN -navigate mode.

The IRU enters the navigate mode from theupon completion of alignment.

The IRU enters the align downmode from the

. NAV-TO-ALIGN-TO-NAV - The IRU enters align downmode from thenavigate mode and, after 30 seconds, automatically reenters thenavigate mode upon completion of downmode alignment.

● ALIGN-TO-ATT or NAV-TO-ATT - The IRU enters the erect attitudesubmode for 20 seconds, during which the MSU ALIGN annunciatorlights. The IRU then enters the attitude mode.

● ATT-TO-ALIGN or ATT-TO-NAV - Once the mode select switch has beenset to ATT, the IRU remains in the attitude mode even if the modeselect switch is reset to ALIGN or NAV. The mode select switchmust be set to OFF for at least 3 seconds before the alignment ornavigate mode can be reestablished.

s ATT-, NAV-, or ALIGN-TO-OFF - After a 3-second delay, the IRUenters the power-off submode for approximately 7 seconds. At theend of 10 seconds, the IRU is off.

. ATT-, NAV-, or ALIGN-TO-OFF-TO-ALIGN, -NAV, or -ATT - If the modeselect switch is reset to ALIGN, NAV, or ATT after being in theOFF position for 3 seconds but before the 10-second power-downprocedure has been completed, the IRU completes the power-downprocedures and then reinitiates power-on procedures.

22-14-00Page 117

Aug 15/91Use or dmclosure Of information on this page is subject to the restrictions on the title page of this document.

MAINTENANCE

Honeywell !!it%h.2. B. (2) Annunciators

The following annunciators indicate IRS status:

● ALIGN indicates that the IRU is in the alignment mode. Aflashing ALIGN annunciator indicates incorrect latitude/longitudeentry, excessive aircraft movement during alignment, or mismatchbetween entered and computed latitude.

. ON BATT indicates that backup battery power is being used.

● BATT FAIL indicates that backup battery power is inadequate tosustain IRS operation during backup battery operation (less than21 volts).

. FAULT indicates an IRS fault.

(3) Test Mode

. The test mode is selected by pressing the MSU TEST switch or IRUINTERFACE TEST switch. The test mode can be selected in eitherthe alignment mode or the navigate mode without affecting basicIRS functions. The test mode is inhibited in the attitude mode,and in the navigate mode when the aircraft ground speed exceeds20 knots.

● When either switch is pressed, the IRU outputs a preprogrammedset of fixed digital signals to aircraft instruments.

. The signals are output in three eight-second phases as shown inTable 2-6. During the first phase, the IRU exercises all flagsand annunciators. During the second and third phases, the IRUoutputs fixed signals for displays on cockpit instruments. Atthe completion of 24 seconds, all outputs return to theiroriginal state.

● If either switch is held on, the signals output during the thirdphase continue to be output until the switch is released.

22-14-00Page 118


MAINTENANCE

Honeywell !!W%L.

ANNUNCIATOR/ FIRST SECOND THIRDSIGNAL PHASE PHASE PHASE

ANNUNCIATORS:

ALIGNFAULTON BATTBATT FAILNO AIRNAV RDY

Fault Ball

ISDU DISPLAY

TrackGroundspeedLatitudeLongitudeWind directionWind speedTrue headingTime-to-NAV

OnOnOnOnOnOn

0s

90”200 knN22° 30.0’E22° 30.0’30°100 kn30°0s

0s0s0s0s0s0s

0s

90°200 knN22” 30.0’E22” 30.0’30°100 kn30°0s

0s0s0s0s0s0s

0s

90°200 knN22° 30.0’E22° 30.0’30°100 kn30°0s

OS = Original State

Test Mode OutputsTable 2-6

22-14-00Page 119/120

Aug 15/91Use or disclosure of information on this page is subject to the restrictions on the title page of this document,

MAINTENANCE

Honeywell F!s!+%..

11,—

u

e

●

f

M

NA

a

9

z

P

h

L

s

T

w

v

—

1 MODE SELECT SWITCHSYSTEM 1

rfl ——— —-lLOGIC GROUNDSYS 1 -

OFF ) IMODE SEL 1 SYS 1

5ALN I

1°1 NAV I II

ATT Io

OFFI

n I IIIv.

ALNo

MODE SEL 2 SYS 13 I ~:

NAVo I

IATT o IOFF

+28 VDC SYS 1 21 ALN

9

1NAV o

I*,*BLOWER CONT REL SYS 1

~%ws L–––––

I

NOTES:

A1 250 mA MAX.

A2 PIN ARRANGEMENTS ANO FUNCTIONS ARE IDENTICALFOR CONNECTORS J 1, J2. ANO J3.

A3 WIRES FROM CONNECTOR J 1 PINS S, T, W, AND VARE TIED TOGETHER INTERNAUY WITH RESPECTIVEWIRES FROM CONNECTORS J2 AND J3.

IPANEL LIGHTPLATE

I r —— —— —

I I +@

—— .— —— —. __1

ALIGN

1ANNUNCIATOR

JI A4J1

-4 :

ON BAITANNUNCIATOR

R

BATT FAIL— ANNUNCIATOR

—

@l

FAULTANNUNCIATOR

ALIGN SYS 1 ‘ ‘ ‘ u I 15

T u 14

ON BATT SYS 1 u I 13

T u () 4) I 12

BATT FAIL SYS 1 u , 11

I1 u {) ) 10

FAULT SYS 1 14 ! 9

T u 4) (F 1 1

ANN TEST SYS 1 7

u I 6ANNUN HI 5

A 3 I4

PNL LTG 5 VAC B

PNL LTG RTN 16

A I A4J2

TEST SW 1 3u 8 TEST

TEST SW 2 I 16SWITCH

I

Mode Select UnitSchematic Diagram

Figure 2-4

Q!4HLIGHTPANEL

l————..———— ——————

22-14-00

IIIIIIIIIIIIII 95.a352mc6

82KIX1444 a

J26(X0436 SW 8

Page 121/122Aug 15/91

Use or disclosure of mlormahon on Ih!s page IS subject to the restrictions on the title page of this document.

R&luy~ANCE

Honeywell .U.s,.w.


22-14-00Page 123


2. c.I

Inertial System Display Unit (ISDU) (See Figures 2-5, 2-6, and 2-7, andTables 2-7 through 2-11).

DISPLAY

SELECT

SWITCH

DIMMER

KNOB

DISPLAY

*

@l

Honeywell LASEREF

‘d e

DG/STS

I@@

KEYBOARD

SYSTEM /

DISPLAY

SWITCH

Inertial

CUELIGHTS

AD- I 6474

System Display UnitFigure 2-5

Dimensions:

Length ....................................... 6.58 in. (16.71 cm)Width ......................................... 5.75 in. (14.6 cm)Height ........................................ 4.5 in. (11.43 cm)

Weight ............................................. 5.0 lb (2.27 kg)

Power Requirements .......................... 5Vacordc, 4.5 WMax

Mating Connectors:

J1 ..........................................................................................

:; .............................................

Mounting .........................................

28 V dc, 10 W Max

M83723/75R-1831-NM83723/75R-1831-7M83723/75R-1831-6

Unit Dzus Fastener

Inertial System Display UnitLeading Particulars

Table 2-7 22-14-00Page 124


MAINTENANCE

Honeywell MN!K!bThe ISDU selects data from any one of three IRUS for display and providesinitial position or heading data to

Operator inputs to the ISDU providenavigational data for display. The

● Keyboard. Display● System display (SYS DSPL) switch● Display select (DSPL SEL) switch● Dimmer knob

the IRUS.

position data to the IRU and selectISDU contains:

2. c. (1)

(2)

(3)

Keyboard

(a) The keyboard is used to enter 1atitude and longitude in thealignment mode or magnetic heading in the attitude mode. TheISDU then sends the entered data simultaneously to all IRU inmultiple-channel installations.

(b) The keyboard contains 12 keys. Five of the 12 keys are dualfunction: N/2, W/4, H/5, E/6, and S/8. A dual-function key isused to select either the type of data (latitude, longitude, orheading) or numerical data to be entered. Single-function keysare used to select only numerical data.

(c) The clear [CLR] and enter [ENT] keys contain green cue 1ightswhich, when lit, indicate that operator action is required.[CLR] is used to remove data erroneously entered. [ENTJ isused to send data to the IRU.

Display

The 13-digit alphanumeric split display shows two types ofnavigation data at the same time. The display is separated into onegroup of six digits (positions 1 through 6) and one group of sevendigits (positions 7 through 13). Punctuation marks (located inpositions 3, 5, 6, 10, 12, and 13) light when necessary to indicatedegrees, decimal points, and minutes.

System Display (SYS DSPL) Switch (See Figure 2-6.)

The SYS DSPL switch is used to select the IRU (position 1, 2, or3)from which the displayed data originates. If the switch is set toOFF, the ISDU cannot send or receive data from any IRU.

22-14-00Page 125


SYS DSPL SwitchFigure 2-6

2. C. (4) Display Selector (DSPL SEL) Switch

The DSPL SEL switch has five positions to select data displayed onthe

(a)

(b)

(c)

(d)

ISDU:

TEST selects a display test that lights all display elementsand keyboard cue lights to allow inspection for possiblemalfunctions. The DSPL SEL switch is spring loaded and must beheld in this position.

TK/GS selects track angle in degrees on the left display andgroundspeed in knots on the right.

PPOS selects latitude on the left display and longitude on theright. Both latitude and longitude are displayed in degrees,minutes, and tenths of minutes.

WIND selects wind direction in degrees on the left display andwind speed in knots on the right.

22-14-00Page 126

Aug 15/91Use or disclosure of information on thispage is subjacfto the restr-ictlonson the titlepage of thisdocument.

2. c. (4)

(5)

(6)

(7)

(e) HDG/STS selects heading or alignmentde~endinq ur)onthe current IRU mode.

status for display,Headinq is dis~la.yedin

degrees ~nd”tenths of degrees, and time-to-aiignment’completionis displayed in minutes and tenths of minutes. In thealignment mode, the ISDU displays alignment status (time to NAVready) in the right display. In the navigate mode, the ISDUdisplays true heading in the left display. In the attitudemode, the ISDU displays magnetic heading in the left displayand ATT in the right display.

Dimmer Knob

The dimmer knob is mounted on, and operates independently of, theDSPL SEL switch. As the dimmer knob is rotated clockwise, thedisplay brightens. As the dimmer knob is rotated counterclockwise,the display dims.

Data Outputs

(a) The ISDU simultaneouslytransmits ARINC 429 low-speed (12.5KBPS) initializationdata on three separately buffered digitalbus output ports to ARINC 429 port no. 3 on each IRU.

(b) The data is BCD-transmitted upon manual command in a “burst”mode of 2 to 4 transmissions within 2 seconds.

(c) Refer to table 2-8 for a list of the ARINC 429 data transmittedfrom the ISDU to the IRU.

(d) Set heading data is accepted by the IRU only when in theattitude mode.

Data Inputs

(a) The ISDU receives navigation data from the IRU on the ARINC 429high-speed digital data bus.

(b) Table 2-9 defines the ISDU input data and theircharacteristics.

(c) All input data words, except the IRS Discrete Word, 1abel 270,are in BCD format.

(d) Table 2-10 defines the use and bit functions of the IRSdiscrete word (octal label 270) found in table 2-9.

(e) Table 2-11 defines the use and bit functions of the Time-To-NAVReady discrete word (octal label 351) found in table 2-9.

22-14-00Page 127


MINSIGNIF- UPOATE

OCTAL I CANT OIGITAL RESO- MSB POSITIVE RATESIGNAL LABEL FORMAT CHAR RANGE LUTION WEIGHT UNITS SENSE (SPS)

Set Latitude 041 BCO 5 90 S-90N 0.1 MIN -- DEG NORTH 2MIN

Set Long itude 042 BCD 6 180E-180W 0.1 MIN -- OEG EAST 2MIN

Set Magnet ic 043 8CD 4 0-359.9 0.1 DEG -- OEG CW FROM N 2Head i ng

Inertial System Display UnitARINC 429 Digital Output Data

Table 2-8

MINSIGNIF- UPDATE

OCTAL ICANT OIGITAL RESO- MSB POSITIVE RATESIGNAL LABEL FORMAT CHAR RANGE LUTION WEIGHT UNITS SENSE (SPS)

Present Posit ion 010 BCD 5 90S-90N 0.1 -- OEG NORTH 2Latitude* MIN

Present Posit ion 011 BCD 6 180E-180W 0.1 .- OEG EAST 2Longitude MIN

True Heading 044 Bco 4 0-359.9 0.1 -- DEG CW FROM N 2

Magnet ic Heading 014 BCD 4 0-359.9 0.1 -- DEG CW FROM N 2

Ground Speed* 012 BCD 4 0-2000 1.0 -- KTS ALWAYS POS 2

Track Angle - 013 8C0 4 0-359.9 0.1 -- DEG CW FROM N 2True*

W indspeed* 015 BCD 3 0-256 1.0 -- KTS ALWAYS POS 2

Wind Oirection - 016 BCD 3 0-359 1.0 -- DEG CW FROM N 2True*

IRS Discretes 270** Dis -- -- -- -- -- -- 2

Time to NAV ROY 351** BCD 2 0-9.9 0.1 -- MIN ALWAYS POS 2*

*Not provided by AHRU.**Refer to table 2-10.

●**Refer to table 2-11.

Inertial System Display UnitARINC 429 Digital Input Data

Table 2-9

22-14-00Page 128


BIT(S) FUNCTION O STATE 1 STATE

o-7 Label bits .- --

8 Source destination identifier Open(SDI) #2

Gnd

9 SDI #1 Open Gnd

10 Align mode Not align mode Align mode

11 Attitude mode Not attitude mode Attitude mode

12 NAV mode Not NAV mode NAV mode

13 Set heading Not set heading Set heading

14 Attitude invalid (IRU FAULT Attitude valid Attitude not validannunciator on)

15 BATT fail Not BATT fail BATT fail

16 ON BATT Not ON BATT ON BATT

17 Air data input invalid Air data valid Air data invalid

18 IRU fault (BITE) Not IRU fault IRU fault

19 BATT fail (ON BATT) Not BATT fail BATT fail

20 Alignment fault hb~l~lignment Alignment fault

21 No initialization Initialized Not initialized

22 Excessive motion (Align mode) No40i~~cessive Excessive motion

23 Air data computer (ADC) or t’4g’41$DCor IRU ADC or IRU faultIRU fault

24 Not used

25-27 Time to NAV RDY:

BITSMinutes 25X27

10-6 1116-5 p;5-44-3 001;-: 110

;;:1:0

NAV RDY 000

28 Not used -- --

29-30 SSM -- --

31 Parity, odd --●

--

I ARINC 429 IRU Discrete Word (Octal Label 270)Table 2-10

22-14-00Page 128.1Apr 15/93

Use or disclosure of information on this page issubject to the restrictionson the titlepage of thisdocument.

BIT(S) I FUNCTION I REMARKS

o-7 Labels bits I --8-13 INot used I

--

14-17 Time to NAV Least significant digit (LSD) inbinary coded decimal (BCD)

18-21 Time to NAV Most significant digit (MSD) in BCD

22-31 INot used --

ARINC 429 Time to NAV Ready Discrete Word (Octal Label 351)Table 2-11

22-14-00Page 128.2

Apr 15/93


r—lIIIII

ISDU

I

IIIL—-

A1J1

H1 4

4L

}5 V AC PANEL LIGHTING

2

3+28 V DC POWER

4i RETURN

7 4 CLR-ENT DIMMING A 2

8 CHASSIS GROUND

11 4 ADATA DISPLAYTEST 3(RESERVED)

24 + a-% ISDU BUS (OUTPUT) HI

I [ ,SDU BUS (OljTpUT) LO* # :}

ARINC 429

25LO-SPEED

26+,=. ISDU (INPUT) HI

4

[ I ,SDU (lNpUT) LO4 3 8 }

ARINC 429

27HI-SPEED

L

A1 PIN ARRANGEMENT IS IDENTICAL ON

A

CONNECTORS J2 AND J3.

2 GROUND FOR DIM AND OPEN FOR BRIGHT.

A 3 GROUND FOR ILLUMINATION AND OPENFOR OFF.

AD-31303#

ISD1.1Wiring DiagramFigure 2-7

22-14-00Page 128.3Apr 15/93

Use or disclosure of information on thispage issubject to the restrictionson the titlepage of thisdocument.

2. D. Optional LASERTRAKW Navigation Display Unit (NDU)2-9, and Tables 2-12 through 2-16).

[See

FROMTOWPT DATA ENTRYDISPLAY KEYBOARD

ON/OFFSWITCH I DATA DISPLAY /

/

DATA SELECT ISPECIALFUNCTIONKEYBOARD KEYBOARD

SYSTEM SELECT FAULTKEYBOARD ANNUNCIATOR

871211-1

Navigation Display Unit

Figure 2-8

Figures 2-8,

Dimensions:

Length ........................................ 6.0 in. (15.24 cm)Width ......................................... 5.75 in. (14.6 cm)Height ........................................ 4.5in. (11.43 cm)

Weight ............................................. 3.81b(I.72 kg)

Power Requirements ............................ 5Vacordc, 3WMax28 V dc, 10 WMax

Mating Connectors:

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . M83723/75R-1831-N:; ::::::::::::................................. M83723/75R-1831-7J3 ............................................. M83723/75R-1831-6

Mounting ......................................... Unit Dzus Fastener

Navigation Display UnitLeading Particulars

Table 2-12 22-14-00Page 128.4Apr 15/93


2. D. The LASERTRAK’”Navigation Display Unit (NDU) is a combined navigationcomputer and display unit that can be used to initialize as many as threeglobal positioning inertial reference units (GPIRUS) and to displaynavigation data based on a nine-waypoint flight plan and selected globalpositioning inertial reference system (GPIRS) inputs.

The NDU provides the following functions:

. Entry of position and heading for up to three inertial sensors.● Entry of position for as many as nine waypoints.● Selection of FROM TO leg.● Display of selected ARINC 429 data from each inertial sensor.. Position initialization.. Self-test of front panel operation.

The NDU contains:

● Data entry keyboard● Special function keyboard● SYSTEM SELECT keyboard● Data select keyboard. Displays. Fault annunciator. ON/OFF switch

(1) Keyboards - The NDU contains data entry, special function, SYSTEMSELECT, and data select keyboards.

a. Data Entry Keyboard

~ The data entry keyboard consists of 12-keys that are used toenter, display, modify, or transmit initialization, legchange, and flight plan data.

~ The [ENT] key is used to accept entered data that is beingdisplayed on the data display. The [CLR] key is used toremove the entered data and clear the display. Both keyscontain green cue lights that indicate when operator actionis required.

~ The back key [0 BCK] is used to select the number andposition of the previous waypoint for display when the WPTcue light is lit.

b. Special Function Keyboard - The special function keyboardconsists of [M T], [BRT DIM], and [TST] keys.

~ M T - The magnetic/true north key contains two green cuelights that indicate north reference. When the M cue lightis lit, the NDU displays angular data referenced to magneticnorth. When the T cue light is lit, the NDU displays angulardata referenced to true north.

22-14-00Page 128.5Apr 15/93


2 D. (1) b. ~ BRT DIM - The bright/dim key is used to control the lightningof the two DNU displays. When this key is first pressed and-held, the display intensity increases until the brightestlevel is reached. When this key is pressed and held a secondtime, the display intensity decreases until the dimmest levelis reached.

~ TST - The test key is used to start a test of all NDUannunciators, cue lights, and displays.

c. SYSTEM SELECT Keyboard

~ The SYSTEM SELECT keyboard consists of the [1], [2], and [3]keys that are used to select data for display from inertialreference unit (IRU) 1, IRU 2, or IRU 3. Each key contains agreen cue light that lights when the data from that IRU hasbeen selected for computation and display. Since the NDUdisplays data from one IRU at a time, only one cue light is

at a time.

Unless otherwise indicated, IRU 3 refers to the IRU orattitude/heading reference unit (AHRU) in the number 3IRU position.

~ Some inertial reference systems (IRS) installationswith theLASERTRAK NDU may include only one or two IRU. If a SYSTEMSELECT key is pressed and no IRU is present, the NDU displaysdashes.

d. Data Select Keyboard

~ The data select keyboard consists of eight keys that are usedto select data for display. Each key contains a cue lightthat when lit indicates what type of data has been selectedfor display. When the [WPT], [LEG CHG], [POS], or [HDG STS]key is pressed, the operator can enter new data or modify thedata that the NDU is displaying. If data is selected fordisplay from an attitude-heading reference unit (AHRU), theNDU displays only heading and status. Data select keys aredescribed below.

~ WPT - The waypoint key is used to build and display a flightplan consisting of one to nine waypoints. When this key isrepeatedly pressed, the NDU displays the successive flightplan waypoints by number and position.

~ LEG CHG - The leg change key is used to define or modify thecurrent flight plan leg by displaying the FROM and TOwaypoint numbers. The current leg must be defined beforenavigation data can be defined and displayed.

22-14-00Page 128.6Apr 15/93


2. D. (1) d. ~ XTK DTK - This key is used to select and display thecrosstrack position in nautical miles and the desired track.

~ DIS TIM - The distance/time key is used to select and displaythe distance and time to the TO waypoint based upon presentposition and current ground speed.

Q TK GS - This key is used to select and display the currenttrack and groundspeed.

~ Pos - The position key is used to initialize and display thepresent position of the aircraft.

Q WD WS - This key is used to select and display the currentwind direction and windspeed.

~ HDG STS - The heading/status key is used to display thecurrent aircraft heading. When the IRU is in the align mode,the NDU displays alignment status (time remaining until navmode entry). When the IRU is in the nav mode, the NDUdisplays heading in the left display, and when the IRU is inthe attitude mode, the NDU displays magnetic heading in theleft display and “ATT” in the right.

(2) :::;]:;s - The NDU contains a data display and a FROM TO WPT.

a. Data Display - The NDU data display consists of two displays: asix-position display and a seven-position display, each havingdegree, decimal, and minute indicators.

b. FROM TO WPT Display - These three elements display the FROM andTO waypoint numbers of the current leg and the number of thewaypoint defined by the position being displayed.

(3) Fault Annunciator - The NDU FAULT annunciator lights when NDUbuilt-in-test equipment (BITE) detects internal failures.

(4) ON/OFF Switch - The ON switch is used to power on and power offNDU.

the

22-14-00Page 128.7Apr 15/93


2. D. (5) Data

(a)

(b)

(c)

(d)

(6) Data

(a)

(b)

(c)

SIGNAL

Set Latitude

Set Longitude

Set MagneticHead i ng

Crosstrack

Status Word

Honeywell

outputs

The NDU simultaneously


transmits ARINC 429 low-speed (12.5KPBS) data on three separately buffered digital bus outputports.

Refer to table 2-13 for a list of the ARINC 429 datatransmitted from the NDU.

The NDU does not output set latitude, set longitude, and setmagnetic heading signals (table 2-13) at a repetitive rate;instead, the NDU outputs these signals as a “burst” of two tofour transmissions at a 5/s rate when the data is entered bythe pilot.

Table 2-14 defines the use and bit functions of the status worddiscrete (octal label 275) found in table 2-13.

Inputs

The NDU receives data from IRU 1, IRU 2, and IRU 3 throughports 1, 2, and 3 (connectors Jl, J2, and J3), respectively.

Table 2-15 defines the NDUcharacteristics.

Table 2-16 defines the usediscrete word (octal label

Table 2-17 defines the useReady discrete word (octal

OCTALLABEL—

041

042

043

116

275*

FORMAT—

Bco

BCD

BCO

BNR

DIS

SIGNIF-ICANTCHAR

5

6

4

15

--

* Refer to Table 2-14 for octal label format.

OIGITALRANGE

90s to90N

180E to180W

o to359.9

+ 128

--

input data and their

and bit functions of the IRS270) found in table 2-15.

and bit functions of the Time to NAVlabel 351) found in table 2-15.

RESO-LUTION

0.1 MIN

0.1 MIN

0.1 DEG

0.004nm

--

MS8WEIGHT—

--

--

--

64

--

UNITS—

OEGMIN

OEGMIN

DEG

Nm

--

POSITIVESENSE

NORTH

EAST

CW FROM N

Fly left

--

MINUPOATERATE(SPS)

--

--

--

25

5

Navigation Display UnitARINC 429 Digital Output Data

Table 2-13

22-14-00Page 128.8Apr 15/93


DTTIC\ CllNrTTf3N RFMflQKCVll {a) I Ullu 1 Lull ,\L# ,n, \,\d

o-7 ARINC label --

8-9 Not used --

10 Waypoint alert O = No waypoint alert1 = Waypoint alert

11-21 Not used --

22-23 TO/FROM 00 = Not FROM or TO01 = TO10 = FROM11 = Invalid

24-26 Not used --

27 Mag/True north o = Magnetic1 = True

28 CDI flag o = Invalid1 = Valid

29-30 Sign status matrix (SSM) 00 = Undefined01 = No computed data (NCD)10 = Functional test11 = Normal

31 Parity, odd --

ARINC 429 Status Word Discrete (Octal Label 275)Table 2-14

22-14-00Page 128.9Apr 15/93


MINSIGNIF- UPDATE

OCTAL ICANT DIGITAL RESO- MSB POSITIVE RATESIGNAL LABEL CODE CHAR RANGE LUTION WEIGHT UNITS SENSE (SPS)

IRS Discretes 270* DIS -- -- -- -- -- -. 1

Pos Latitude 310 BNR 20 *0.5 0.000001 0.5 Deg/lBO N from 5(6NR) o“

Pos Longitude 311 BNR 20 tl.o 0.000001 0.5 Deg/lBO E from 5(6NR) o“

Ground Speed 312 BNR 15 0 to 4096 0.125 2046 KtS A 1ways 10(6NR) pos

TK Angle True 313 BNR 15 tl.o 0.00003 0.5 Deg/180 CW from 25(fINR) N

True Hdg (BNR) 314 BNR 15 *1. O 0.00003 0.5 Oeg/180 CW from 25N

W indspeed 315 BNR B O to 255 1.0 128 KtS Always 10Pos

Wind Oirection - 316 BNR 8 il.o 0.0039 0.5 Deg/180 CW from 10True N

TK Angle - Mag 317 BNR 15 il.o 0.00003 0.5 Deg/180 CW from 25N

Magnetic HDG 320 BNR 15 *1.O 0.00003 0.5 Deg/lBO CW from(6NR)

25N

Time to NAV Ready 351 BCO 2 Oto 9.9 0.1 -- Min -- 1**

*Refer to table 2-16.**Refer to table 2-17.

Navigation Display UnitARINC 429 Digital Input Data

Table 2-15

22-14-00Page 128.10Apr 15/93

Use or disclosure of information on thispage issubject to the restrictionson the titlepage of thisdocument,

BIT(S) FUNCTION O STATE 1 STATE

o-7 Label bits .- --

8 Source destination identifier Open Gnd(SDI) #2

9 SDI #1 Open Gnd

10 Align mode Not align mode Align mode

11 Attitude mode Not attitude mode Attitude mode

12 NAV mode Not NAV mode NAV mode

13 Set heading Not set heading Set heading

14 Attitude invalid (IRU FAULT Attitude valid Attitude not validannunciator on)

15 BATT fail Not BATT fail BATT fail

16 ON BATT Not ON BATT ON BATT

17 Air data input invalid Air data valid Air data invalid

18 IRU fault (BITE) Not IRU fault IRU fault

19 BATT fail (ON BATT) Not BATT fail BATT fail

20 Alignment fault t!~!l;lignment Alignment fault

21 No initialization Initialized Not initialized

22 Excessive motion (Align mode) No!Oig;cessive Excessive motion

23 Air data computer (ADC) or hJg;l~DC or IRU ADC or IRU faultIRU fault

24 Not used

25-27 Time to NAV RDY:

BITSMinutes 25 26 27

10-6 1116-5 011

101H 0013-2 1102-1 010

100NA;-[DY 000

28 Not used -- --

29-30 SSM -- --

31 Parity, odd -- --

ARINC 429 IRU Discrete Word [Octal Label 270)Table 2-16’

,

22-14-00Page 128.11Apr 15/93

Use or disclosureof informationon thispage issubjectto the restrictionsonthe titlepage of thisdocument.

BIT(S)

o-7

8-13

14-17

18-21

22-31

FUNCTION

Labels bits

Not used

Time to NAV

Time to NAV

Not used

REMARKS

--

--

Least significant digit (LSD) inbinary coded decimal (BCD)

Most significant digit (MSD) in BCD

--

ARINC 429 Time to NAV Ready Discrete Word (Octal Label 351)Table 2-17

22-14-00Page 128.12Apr 15/93

Use or disclosure of information on this page issubject to the restrictionson the title page of this document.

w5V LIGHTINGCIRCUIT

G1230AC NDU

~Jl

5V/H

5vjL

+ 28 V/FWVR

+ 28 V/RTN

BRT/DIM

CASE/GND

ON/OFFTEST/L

ARINC/OUT/H

ARINC/OUT/L

ARINC/lN/H

ARINC/lN/L

1

2

3

4

7

8

911

24

25

26

27

I i

II

\/

J2

m

TO IRU 2

DJ3

TO IRU 3

OR AHRU

B

A13

A14

G7

G8

1-

J1

+28 VDC ISDUPWR

lN/PRT/H 1,LSDUlN/PRT/L INPUT

OUT/PRT/H 1OUTPUT

BUSOUT/PRT/L 429H-11, I

lopCDIPOWER CONTROL

,

CDI

r>IiII t~J

I fp&

NOTES:

A1 NDU POWER ISSUPPLIEDFROM IRUWHEN IRUISOPERATING.

A 2 IFINPUTISGROUND, KEYBCIARD CUE LIGHTSARE DIM WHEN ON. IFINPUT ISOPEN,KEYBOARD CUE LIGHTSARE BRIGHT WHEN ON.

A 3 IFINPUTISGROUND, NDIJENTERS TEST MODE.

NDU Wiring DiagramFigure 2-9

95-83081303

22-14-00Page 128.13/128.14

Apr 15/93


MAINTENANCE

Honeywell !YW$h.


22-14-00Page 129


MAINTENANCE

Honeywell W!+A&.3. ADZ-81O Air Data Svstem

AZ-81O Digital Air Data Computer (See Figures 3-1 and 3-2, and Tables 3-1,3-2, and 3-3.)

AD-3037-R1

AZ-81O Digital Air Data ComputerFigure 3-1

22-14-00Page 130


MAINTENANCE

Honeywell !W%!h.


Length (including hand”Width ................Height ...............

e) ....................... 15.76 in. (400.3 mm).......................... 3.51 in. (89.15 mm).......................... 7.62 in. (193.5 mm)

Weight (maximum) ........................................ 9.9 lb (4.5 kg)

Power Requirements .................................... 28 V dc, 40 WMax

Mating Connector:

J1 ........................................ DPX2MA-A106P-A1O6P-33B-OO31

Mating Pneumatic Connectors:

Pitot (straight) .......................................... 40007-1B24*Static (straight) ......................................... 40007-1A26*Pitot (90° elbow) ........................................ 40007-1B24E*Static (90° elbow) ....................................... 40007-1A26E*

*All part numbers are American Safety Flight Systems.

Mounting ................................ Tray, Honeywel1 Part No. 7007974

AZ-81O Digital Air Data ComputerLeading Particulars

Table 3-1

The AZ-81O Digital Air Data Computer (DADC) is a microprocessor-based digitalcomputer which accepts both digital and analog inputs, performs digitalcomputations, and supplies both digital and analog outputs. It receivespitot-static pressures and total air temperature inputs for computing thestandard air data functions. The air data equations are solved directlyusing a 16/32-bit arithmetic microprocessor under the control of an 8-bitgeneral purpose microprocessor. The DADC provides outputs to the electronicdisplay system, transponder, flight recorder, flight guidance computer, aswell as other elements of the system, such as fault warning computers andinertial reference system (IRS). Angle of attack inputs have beenincorporated, and computed AOA outputs are provided. The DADC performsaltitude preselect operation and alerting functions. The DADC also usesangle of attack to process static source error correction for the standardair data outputs.

Analog and discrete outputs are shown on Figure 3-2. Digital outputstransmitted on the ASCB are listed in Table 3-2, and ARINC 429 outputs arelisted in Table 3-3.

22-14-00Page 131

Aug 15/91Use or disclosure of information on this page is subject to the restncllons on the tttle page of this document.

NN

1

USPBIT 011FUNCTION NOTEFORKAT SCALE RSBAPPROXRESOL/LSOPOSSENSE

01I

FTIUSIHVAR... ...... ~;i;-;i;;----------------l----l;;-;~;-------------------l--------------------l---l---------------”-l-----------------1----1--------1

I l;~;;-;;~;~-------------l----l;~;;~;-~&~-------------l-----------------"--l---l-----------"-----l-----------------l~;l--------l.........“11 15 TEST LOGIC1 ● TEST(SELFTESTANO14AINT.TEST) 5ACUFASAIICTL1 14 VAL1O LOGIC1 = VALIO FASAUCVL1 13-11 SPARE1 10-8 COUNTER O-7 HEX1 1-0 DAOCAL)L)RESS 06 H - LEFT O? H - RIGIITII... ...... .........................l----li~~pi~~~~~~--------li;--~~~~~-~~~~-----l;~-;~-~-;------------1~~---------------1~~~~1--------I

2 15-0 PRESSUREALIITUUE - . .

II II5AC2RADIIPL

..-...... ......................... -------------------------------------------------... ................. .................3 15-0 BAROALTIlUOE

I 116120, 20FT ITHO’S COMPLEMENT

l;;~;l--------l+/-65536FEET . . UP

II II5AC4 RADIRICL

. . . ------ . . . . . . . . . . . . . . . . . . . . . . . . . 1 I 1-----------------1------.....-.4------------------------------------..4 15-2 ALTITUUERATE TUO’S COMPLEMENT

;;---------------I;;;;l;fi;;i-1+/-32768FT PERHIM 14 4.0/1.0FPH

41 BAROMT FLAG LOGIC 1 ■ VALID40 PRESSUREALTFLAG LOGIC1 ● VAL1O

5AC6FASBALTL

II l----lfi;fii;~~;~l+,--------l+,~024~“o,~ 11410,25,003,25IFASPALII.

..- . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . -------- --- . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .S 15-2 CALIBRATE AIRSPEED

1#-------i. FORUARO

51 SPARE.0

5AC8 :::~::~L50 CAS FLAG LOGIC 1 = VALID

I l--”---------------”------l Iwo,sCWPLEHENT... ...... ----.........................6 15-2 TRUEAIRSPEED

I;;--i;;-~~~----”-l;;-l;-;;;;;;;”;l;”---l;;;~~~;----------l;Mll~M;~;i-l-SPAltE

.*

:: TAS FLAG5ACA

I l--------------------"----l----l;;-:&;;i"--------lFASIAS1

. . . ------7 15-2 MACH

II-------------------- --- . . . . . . . . . . . . . . . . . I+/- 2.0 MACH

-------------”---l#;~~p14 0.000244/O.0DO061FORMARD

71 SPARE?.0

5ACCWICII FLAO LOGIC1 ■ VAL19 FAS14NL

I l"------------------------l----i;;;;-~fii~;E;~-------l-........815-3 TO:;~R~lRTEMP

;;--;;;-;;;~-------li;-l;;;;;;~-w;;i-"-l;&--------------Iitiil;iiii;i-la 2-1

. .CAPC

no- TEMP-FLAG(TAT,SAT)II

LOGIC1 ● VALID... ...... 11....................---------.........................9 15-3 STATICAIRTEMP

I....................TUO’S COMPLEMENT

92-0+/-256DEGC

SPAREII----------------------------------1--”-l;~;;;;i;;;---”-----l:;--;;;;;-;;~-----

10 15-2 PRESELECTEDALTITUOE101 IN-MOTIONFLAG LOGICO ■ IN-H3TION10.0 , FLAG .LOGIC1 = VALIO

a nbs

‘l;;-l;&;--;;----;;---l;;~--------------1;;;;1:;”.. .

‘l;;-l;;;;-;-;;-------l~---------”-----I::li=ixil.5A02FAOCRANL

FASAPSLII-.. . . . . . . I l-------------------------l--------------------[---l-----------------l-----------.........................----

&

h)N1

USPBIT 011FUNCTION NOTEFORNAT SCALE RSBAPPROXRESOL/LSBPOSSEHSE FTIUS114VARI l---"---------------------l----lfi~T~-~&j~[~~~--------[Jj1-;~~~-~~fi~------ii~-i~-;~~;~-~~i~~----l~~~~~~~----------1~~~~l~..-------

1I 15-2 HAXALLOUABLCAIRSPCEO . .11 1-0 SPAREI li~~~-~~~~~~----”---”-l lTM~,s~~pLE~ENT

5AI14... ...... I-------------------------------------------------I1610 ~,953 ,N,,G Ip”sA*ltspEE” 1;;;;1--------1---................. .................12 15-o +/-64 IN-M . -

II II5A06

.-. ------ . . . ---------------------- .-*. ------------------------- I 113 100439,0*M549I-----------------------................. 15A981RMAOA*LI13 15-3 TR~fA~~A TuO’S COMPLEWNT

................-----........+/-180OEG UP

132-1.

5A08 FASAOAL130 AOA FLAGII

LOGIC1 = VALID... ...... II I........................--------------------------------------------------l;;-l;&-’-------’-;;-l&--------------”l;~;~l;;~;;-;;l14 15-4 NORHALIZEDAOA TUO*S COMPLEMENT +/-2.o(RATIO)143-0 SPARE

. .

II II5AOA

... ...... ......................... ----------------------------------------------------.................15 15-0 SPARE

I II I-----------------1;;;;1--------I

II II5AOC

--- . ----- --------------------- ---- ---- ------------------------- -------------------- --- ----------------- . . . . . . . . . . . . . . . . .I II I i;&l--------l16 15-0 SPARE

1

II -------------------------1 I5AOE

~“-- ------ 1 I l--"--------------l----------"------l;&i----"---...-----------------------------.-..--...--+--------1715-O SPARE

II II5AE0

..-.....-......................... 1 11----------------------------------------------------.................18 15-0 SPARE

t-----------------I;M;I--------I

II II5AE2

... ...... ........................-----.........................I-“------------------l---l-------"---------l-----------------l;--J--------I19 15-0 SPARE

II --------”--”-------------l----l;;;:;-:-.’-~;~”TSAE4

.-...-*-- 1.........-....................I;;-l;-&;;;;--------l20 15-0 TOTALPRESSURE

;;;-~i~;;;;;-----l;~;l;~~;; I+/-64 IN-IIG ●

II II5AE6

... ------......................... ----------------------------------------.-.-.-.>----................-;! :5-2 MAXALLONABLEMCH

I II ITHO’S COMPLEMENT

-----------------l;~;l;;-- I+/-2.0PAcllSPARE

14 0.0uU244/O.000061FORUARO

210 ALTITUOEALERT5AE8

LOGIC1 = LAMPONl----l;;;;i~;;;”;;”--------I;;--;;:;;;;;;<---Ii;-l;~i-i;~;------l

FAOLALRL. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

22[ 15-0 ‘BARO SET ( IN-IIG);;;-;;;;;;;;----”l;;~~l;~;;;l~-l

- . .

II. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . l---- l------------------------- l--------------------l--- 1-----------------[ ----------------- 1:::%------1

USP BIT 011 FUNCTION NOTEFORMAT SCALE RSBAPPROXRESOL/LSllPOSSENSE FTIUSIMVAR--- -------------------------------l----l;;;;~;i;~;;;;;;---------l;;:-;;;;-;;---------l;;-l;-;;;;;;-;~----l;;~;;~;;;;”;---”-l;;;;l;~;;;~[l23’15-4 ‘BAWS:T (HILLIBARS)233-1

.5AEC

230 BAROSETFLAG LOGIC1 = VAL1O FASBSETLI l-------------------------l----l-------------------------l--------------------l---l----------"------l-----------------l;;~..-------

~f ~;-1I SPAREVmo EXCEEOEO LOGIC 1 = EXCEEOEO 5AEE

24 9-8 FLAPPOS CODE 00. 0 uEG FADDF2L9249-8 ol ● 10 oEG FADOFIL824 9-8 10-20 DEG24 9-8 11 -39 oEG24 7 CAUINPRESSNREVALID LOGIC i = VALID246 FLAP POS VALID LOGIC1 = VALID2i 5-o SPAREI l~;~;;~~~;;-~&-;-’;---l----l;~;----------”..-------

25I............-......*----------l---l-----------------l-----------------1;;;;1--------I

II.........fi;;-;;;---------”------l----l;~-~-~;-------------l-----------”--------l---l-----------------1-----------------1:1--------I26

I lERkoR----------------------------------I----l:;;----------------------l--------------------l---l-----------------l-----------------FJ--------l27

5AF4-----------------l----l--------III......... .........................l----l~~;~~;-------------------l--------------------l---l-----------------

28 NOLC FLAG

MAINTENANCE

Honeywell KfN%&.

Paramter

PressAlt

km km Alt

Ml

x

Ann

rAs

rAT

\lt Me

MT

klacted Altibd

formalized A!M

“Otal Pressure

i@nralt ID

Cktal (fbte 5Label I@i!lteCc&secmd.203

2C4

2a5

208

207

210

211

212

213

102

244

242

371

234—a5

10

20

10

10

10

10

2.5

20

2.5

5

20

10

2.5

10

10

Tranmissia OK&

32313028 2827262524232221201918 17 1615 14 13 12 11 10987654321

P

P

P

P

P

P

P

P

P

P

P

P

P

*

P

P

—

—

s Ksb ----------- --- lab x

s ro b ----------- --- -1*X

0 ri6b------ --see fbta2 --- lsbxx

0 lr6b ----------- -lsbx xxx

0 lmb----------lsb Xxx xxv

o IIE b ----------- -- lsb xxx

s lrab---------lsbx Xxx xxx

s mb---------lsb Xxxx xxx

s Irab---------lsbx Xxx xxx

s ro b ----------- --- lsb AM

s lmb---------lsbx Xxx xxx

s 116 b ----------- --- lsbxx

- --- 00100100 0000110

~ W Fiel~

X= Sa?f&el; .=8imsry Point fbte~

~Siq Bit: O=+ l=-

~ Stitus : 11 = oats Valid

CO= Ma Irwalid

01 = Self Test

W Parity Oeck Bit

Id 21md 31sd 21sd lsd

~ Oats Field

= kilral Point

~ Status : IX = Osta Irwalid . psitim

01 = Self Test

Data Imalid = NOTransnissim

—

—)(—)[—

11000001

00100001

10100001

01100001

11100001

00010001

10010001

01010001

11010001

01000010

00000101

01000101

10011111

+Latel w

Ckstinstim Ccd

““W”= All Call

00111001

10111001

Note 1 : X = Ckm’tcare. my cmt.ain valid kss si~ificant &ta.Note 2 : This binary pint is for millina& ulit.M.e3:V= Werspduamingdih

H = Crakiq nutim discreteA = PLTAlert di~

Note4: ~ ~- MC LOCATIffl1 0 Ri@t Sideo 1 tit sick

fide 5: Mnters listd pruvide tf-a trananissim qxia&s p sEcmd._timel UFdsterates are given in * applicable rkcriptim sactims.

AZ-81O Digital Air Data ComputerARINC 429 Outputs

Table 3-3- 22-14-00Page 135/136

Aug 15/91Use or disclosure of reformation on this page is subject to the restrictions on the title page of fhis document.

Honeywell !!!!!;~.c’

II SL,.

%’:5!53””’SARO DISABLE 36 <r-

sSEC DISABLE 39 <-~

SELF TEST 52 <-=

~194 RED

95 ‘W4 ;~~~MISC

OUTPUTS6 YEL

REGISTER 97 AOA TEST SEL (SL)

96 AOA TEST SEL (15K)

ALTIT”OESWITCH 77 @- ●

ALERTER SEL S3

LEF7 RIGHT PROG 100

k%TloN[:i:E

A

bI ’10

75 ALTALERT LIGHT

76 ALTALERT HORN (’ND)

49 ALTALERT HORN (26 V)a ARITHMETIC

PROCESSOR(9511) I1“0.456

(E

AOA 81TEST 82‘NO ~

JIB

ALT SEL

(:’

103[H)

SLEW (L)●

CONTROL lo4

II

H12JIA63 CASIN PRESSURERATIO (WIpEF

78 PRESSURE ALTSIGNALWA

CONVERTER (H) ’18

)53 AOA REF(L) ~

DISC~’C~E

COND;;:NING

MULTIPLEXING

ANALOGTO

DIGITALCON-

vERTER

~

I101

102

103

104

105

106

JIB AOA60 - PROBE 4

INTERFACE

AOA SIGNAL(WIPER)

l— IAIRCRAFTINTERLOCKKEYING

3(H)

‘1A

1

11 PRIMARY(L) ,2 ASCB

ASCO(H) JIB

113 SECONDARY(L) ,4 ASCB

I

B

(H) .IIB

(L)

1

26 OUTPUT27 NO. 1

(H)70 OUTPUT

(L) 71 “o, zARINC

429(H)

(L)

1

W OUTPUTNO, 3

[H) 3132 OuTpuT

(L) 33 NO. 4

3 PROGRAMMEMOFIY(PROM) I

‘%=33Et-BAROCORRECTIONPOTENTIOMETER

=/

AIRCRAFTCONSTANTS:VMO. SSEC.

ALT SWITCH,A/S SWITCH,ETC. (PROM)

28 W3CSUPPLY

(+-JI1 POWER

SIGNAL 5

TSUPPLY

GROUNO 6 I

Oc7<:

GROUND .q< I J

CHASSIS gGROUND ,0 AD 13)42

AZ-81O Digital Air Data ComputerBlock DiagramFigure 3-2 22-14-00

Page 137/138Aug 15/91

Use or dwclosure of mformatlon cm thw page IS subject to the restrictions on the hlle page of this document.


22-14-00Page 139

Aug-15/91Use or disclosure of information on this pege is subject to the restrictions on the title page of this document,

4. AA-300 Radio Altimeter System

A. RT-300 Radio Altimeter Receiver Transmitter (See Figures 4-1 and 4-2, andTable 4-l.)

RT-300 Radio Altimeter Receiver TransmitterFigure 4-1


Length ..................................... 11.07 in. (281.2 mm)Width ....................................... 4.56 in. (115.8 mm)Height ...................................... 4.09 in. (104.0 mm)


Power Requirements .................................. 28Vdc, 0.7A

Transmitter characteristics (nominal):

Type ..................................... Short Pulse ModulationRF frequency ........................................,... 4.3 GHzPeak power .................................................. 5W

Receiver characteristics (nominal):

Type ............................................ SuperheterodyneIF frequency ............................................. 60 MHz

RT-300 Radio Altimeter Receiver TransmitterLeading Particulars

Table 4-1

22-14-00Page-140

Aug 15/91Use or disclosure Of information on this page is subject to the restrictions on the title page of thts document.

::Wl;:lANCE

Hone~dl .IJLFST.EAMIV

Displayed Operational Altitude ● ............................... 0 - 2500 ft

Data Outputs/Accuracy:

Precision output ....................... DC analog voltage (0 - 2500 ft)Gradient: -4.OmV dc/ft

O alt=OvoltAccuracy: o-loo ft, t3ft

100 - 500 ft, i 3%500 - 2500 ft, t 4%

Auxiliary Output .......... DC analog voltage (O - 2500 ft)Gradient: Per ARINC characteristic 552,

Alt =0.02h +0.4 Vdc below 480 ft

and 10+10Lnh+20Vdc500

above 480 ftAccuracy: O - 100 ft, i 4 ft

100 - 500 ft, f 4%500 - 2500 ft, t 5%

Altitude Trips............. 100 MA current sink provided at andbelow trip points indicated below:

Tri~ Point Accuracy

50 ft t4ft250 ft *11)ft500 ft ~16 ft1200 ft *6O ft

Mating Connectors:

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . MS3116F16-26S;; - TRANSMIT..............● .......................... GRFF 4007-0002

(GRFF Connectors, GRFF Division, Solitron Devices, Inc)J3 - RECEIVE.......................................... GRFF 4007-0002

Mounting ...................................................... Hard Mount

RT-300 Radio Altimeter Receiver TransmitterLeading ParticularsTable 4-1 (cent)

22-14-00Page 141


MAINTENANCE

Honeywell !!!!&!!Sh.The radio altimeter receiver transmitter provides a dc output voltagewhich is proportional to the aircraft absolute altitude above terrain.In addition, it provides radio altitude trip points, an indicator warningflag output, and an auxiliary radio altitude output.

The precision output is used to drive the PFD RAD ALT display andsupplies absolute altitude information to the flight guidance system.

The RT-300 transmitter section is an all solid state, two-transistor unitoperating in a master oscillator, power amplifier configuration. Thetransmitter is collector pulse modulated by a solid state modulator. Asample of the transmitted signal is detected and applied to the processorassembly as a To (system sync) pulse to initiate the timing sequence.

The r-f ground return signal from the receive antenna is input to thereceiver subsystem where it is amplified and converted to produce an i-fsignal. The i-f signal is amplified and detected to produce a videopulse which is further amplified and applied to the processor.

Within the processor, the detected video pulse is compared to the systemsync pulse, and the time difference between these two pulses isdetermined. This time difference is converted to a dc analog voltage(internal range voltage) proportional to the shortest range to theground. The internal range voltage is fed to the output assembly forfurther processing. An automatic gain control (AGC) voltage is derivedin the range tracker and fed back to the receiver assembly to maintainthe amplitude of the video signal at a constant level.

The processor also contains the search and acquisition circuits whichenable the system to initially acquire the ground return signal orreacquire the signal should tracking be interrupted. The search mode isinitiated upon loss of signal. A sweep signal is generated which causesthe system to search for the earliest return signal. This processcontinues until track is reestablished. Once track has been properlyestablished, the search mode is inhibited. A search valid is output tothe output networks to cause the indicator needle to disappear from viewduring the search phase. A sensitivity time control (STC) voltage is fedback to the receiver assembly to gain-program the receiver as a functionof altitude. Gain programming prevents acquisition of the direct antennaleakage during the interval when the transmitter is on.

The internal range voltage is processed within the output networks toproduce dc outputs for driving the indicator needle and other aircraftsystems requiring radio altitude information. Four altitude trip outputswhich supply a ground at or below the preset altitude are provided. Theflag warn signal controls the indicator flag to provide warning of systemmalfunction. In addition, a track invalid signal is available whichindicates if the RT-300 is not in the track mode. Two test inhibitinputs are provided to permit defeat of the pilot-activated self-testwhen the flight control system is engaged.

The self-contained power supply operates from the aircraft 27.5 V dc bus.It provides all RT-300 operating voltages.

22-14-00Page 142


MAINTENANCE

Honeywell !%t%h.

J2v

J3

TO TRANSMIT ~ TO RECEIVEANTENNA ANTENNA

---- -——

m /77

TRANSMITTERI , To PULSE VIDEO

MODULATORRECEIVER

AGC VOLTAGESPROCESSOR

STC VOLTAGE

TRACKJ1 VALID J1

iI

1]W ALTOUTPUT(EH)

SELF TEST T II

X AUXOUTPUT

I 1;Y FLAG WARNING

I 1>F TRACKINVALID

IN OUTPUTCOMMON

ITEST INHIBIT NO.1 D OUTPUT +

I

ASSEMBLY i

ITEST INHIBIT NO.2 B

I1

>

I 1IOUTPUTTEST E

I I )

I A 1>

I ! 28VDCFILTERED 1

u 1200FTTRIPV 250 FT TRIPR 50FTTRIPL 500 FT TRIPP ALT TRIPCOMMON

IL

+ +30 VDC i+28 VDC +15 VDCINPUT PWR q

Z +15 VDC OUTPUT

‘1POWER ~ +5 VDC 1>

I SUPPLY-15VDC

I

PWR GND Q I ~ – 15 VDC OUTPUT

-“’15VDZ:RT-300 Radio Altimeter Receiver Transmitter

Block DiagramFigure 4-2 22-14-00

Page 143Aug 15/91


MAINTENANCE

Honeywell !!K$!!Rh.4. B. AT-222 Radio Altimeter Antenna (See Figure 4-3 and Table 4-2.)

One AT-222 antenna is used to receive information from the RT-300 andanother is used to transmit information to the RT-300.

)1 I

4

AO-14n5 @

AT-222 Radio Altimeter AntennaFigure 4-3


Width (along contoured mounting surface) .... 7.00 in. (177.81Mn)Height ...................................... 2.75 in. (69.85 mm)

Weight (maximum) ................................. 1.25 lb (0.57 kg)

Power Requirements .● ............................. 100 W average max

Mating Connector:

J1 ................................................... TNC - maleGRFF 4007-0002 (Straight)

GRFF 4100-0001 (Right Angle)

Mounting .......................................... Flush mounted toaircraft skin

AT-222 Radio Altimeter AntennaLeading Particulars

Table 4-2

22-14-00Page 144


MAINTENANCE

Honeywell !i!!i$%h.5. paraurat)h5 is not amlicable to this svstem.

22-14-00Page 145

Aug 15/91Use or disclosure of information on this page is subject to the restrtctlons on the title page of this document.

Honeywell !?%!!%.CE6. EDZ-884 Electronic DisI)lay System

The EDZ-884 Electronic Display System consists of six DU-880 Display Units,three SG-884 Symbol Generators, two DC-884 Display Controllers, two DA-884Data Acquisition Units, one DP-884 Dimmer Panel, and two FC-880 Fault WarningComputers. The system displays flight attitudes, headings, courseorientation, flightpath co~ands, weather and mapping pr&entations, mode andsource annunciations, engine performance data, crew alerting messages, andchecklists.

A. DU-880 Display Unit (See Figures 6-1 through 6-4 and Table 6-l.)

AO-30251

DU-880 Display UnitFigure 6-1

22-14-00Page-146

Aug 15/91Use or disclosure Of reformation on this page is subject to the restrictions on the title page of this document.


Length ..................................... 14.00 in. (355.60 mm)Width ....................................... 8.20 in. (208.28 mm)Height ...................................... 8.75 in. (222.25 mm)


Power Requirements:

Primary ....................................... 28 V dc, 215 WmaxLighting ................................. 5V acordc, 1.5Wmax

28 V dc, 1.5Wmax

Mating Connector J1 ............................. DPXBMA-A106-33P-O415

DU-880 Display UnitLeading Particulars

Table 6-1

The DU-880 Display Unit is a large format 8 x 8-inch display which uses ahigh resolution color CRT to display attitude, heading, altitude,airspeed, vertical speed, air data parameters, navigation information,weather mapping, engine data, crew alerting information, and systempages. The six-unit configuration is arranged in the cockpit as shown inFigure 6-2 with the primary display function identified. The displaysare interchangeable except when used as a primary flight display (PFD)and an inclinometer is attached to the bezel. In addition to the circuitbreaker panel, power may be removed from the display system by thedisplay power panel located in the overhead on the pilot’s side of thecockpit. This panel is shown in Figure 6-3. The leading particulars ofthe DU-880 are listed in Table 6-1. The electronic display systemoperation is described with reference to Figure 6-4. Figure 6-5 containsa block diagram of the DU-880.

Each of the six display units interfaceswith the symbol generatorsthrough a I-MHz serial digital bus. Each display unit has three l-MHzserial digital bus inputs, each input is connected to one of the threesymbol generators. In addition, each display unit (C)U)has a WX businput, however, only the two navigation displays (DUS 2 and 5) use thisinput, which is connected to all three symbol generators. Bus sourceselect discretes are input to the display units from the reversionarycontroller to select which SG will drive a particular DU.

22-14-00Page-147


Honeywell !#!!!G.cE

EHz El[#11510PFDDU6

nCAS/SWDU4

AD-6827-RI

Display Unit Cockpit ConfigurationFigure 6-2

r-------------------- -— —-— —---- -.-—— . 1

iIIII1III~

iIIIIII

PILOT

DISPLAYSEICAS COPILOT

OFF OFF

7

KOFF

L---e ---------------—-- —- —-- ——— —----— J

Display Power PanelFigure 6-3

22-14-00Page 148

Aug 15/91Use or disclosure of information on this page is subject 10 the restrictions on the title page of this document,

MAINTENANCE

Honeywell U&!#h.Each display unit transmits data to the fault warning computer (FWC) overa private-line (ARINC 429) bus, referred to as the DU wraparound bus.~~;m~;ntent of the data varies depending on DU position and display

. All six DUS transmit activity counter information. The FWCmonitors for activity on the DU wraparound bus by way of the activitycounter. The CHECK DU message is activated whenever the FWC fails to seeany activity on the wraparound bus for DUS 1, 2, 5, and 6. The messageis displayed as appropriate for the DU’S positional location (1, 2, 5,and 6) in the aircraft.

The ENG annunciation is activated on the PFD whenever the FWC fails tosee any activity on the wraparound bus for DU3 (or DU4 with compactedEICAS format). The DU4 annunciation is activated on the PFD whenever theFWC fails to see any activity on the wraparound bus for DU4 when notdisplaying compacted EICAS format.

The DUS displaying the PFD format additionally transmit attitude (pitchand roll), speed (CAS or Mach), and altitude information. The FWCcompares the attitude data received from the PFD’s wraparound bus withthe pitch and roll ARINC 429 data received from the IRS’S. It alsocompares the speed and altitude data received from PFD’s wraparound buswith the CAS or Mach and altitude ASCB data from the DADC’S. If thedifference between the display and sensor data exceed the appropriatethreshold for any one parameter, the CHECK DU(X) message is activated;where (X) is position of DU in the aircraft which is displayingmiscompared PFD data. If the miscompare occurs on both PFDs, then thered CHECK PFD message is displayed. The DUS displaying engine dataformat additionally transmit left and right EPR, left and right TGT, andleft and right N1 and N2 information. The FWC compares the engine datareceived from DU3 (or DU4 with compacted EICAS format) with the ASCB datafrom the data acquisition units (DAUS). If the difference between thedisplay and sensor data exceed the appropriate threshold, for any oneparameter, the ENG annunciation is activated on the PFD.

Provisions for external light sensors are included as inputs to the DUS,however, short-term light changes are adequately corrected for by theDU’S own light sensor located on the front of the DU, while long-termchanges in light conditions is accounted for by the DP-884 Dimmer Panel.The display unit compares the adjacent light sensor (ALS) input (thedisplay light sensor (DLS) output from the adjacent tubes) with its ownlight sensor and uses the larger value to program brightness so that thetubes track each other. Weather radar returns displayed on the twonavigational displays (NDs) can be dimmed independently by using thebrightness knob on the weather radar controllers. A 10-minute warmupperiod is required after power is initially applied to the DUS beforecolor stabilization is complete. Degaussing procedures are provided inthe Gulfstream aircraft maintenance manual.

22-14-00Page 149

Aug 15/91Use or disclosure of information on this pege is subject to the restrictions on the title page of this document.

MAINTENANCE

Honeywell !’&ii!!!#h.Over-temperature protection is provided by internal thermistors. Shoulda DU’S internal temperature increase to a preset level, the rastergenerator will be shut off until the temperature drops back below thistrip point. Should the DU’S internal temperature continue to increaseafter the raster generator has been shut off, an even higher temperaturetrip point will shut down the low voltage power supply (LVPS). At thesame time a discrete is provided to the FWC, allowing an overtemp messageto be displayed. The LVPS will remain off until the temperature dropssufficiently below the higher trip point.

22-14-00Page 150


MAINTENANCE

Honeywell !!!!!!%h.

I II II 1

7 WI DU2

m{<

IRS1

- {u

IRS1

= {= “

IRS1

ARINC IRS2SG1

ARINC IRS2 ARINC IRS2

IRS3 IRS2 42s IRS3

DADC DADC DADCARINC 429 ARINC 42a ARINC 429

m

d

DADC Fwc MU

DU ARINC 429 WRAPAROUNDt 4A4t4

OU ARINC 422 WRAPAROUND

3~\

a 3 3 3\~ g g

A

r-

OU ARINC 42S WRAPAROUNOS:

\

STANDARD =ARINC 429 324

325

{n341

USEROEFINEO

342

ARINC 422

~345● PARAMETER SLEW RATE

ALTITUDEMAcnAIRSPEEDPITCHROLL=TIVtTY COUNTER

LEFr EPRRIGHT EPRLEIT TGTRIGHT TGTLEFr N1RtGHT N1

THRFSHO~ MAXIMUM RATES

200FT ● 20,mo FrMN0.01 M ● .01rniSECam ● 10 KTS/SEC6 DEGREES ● 4QVSEC6 OEGREES ● 40%SEC

0.5

50W

5%

AKX?047S-R1

EDZ-884 Electronic Display System InterfaceFigure 6-4 22-14-00

Page 151/152Aug 15/91


,,r—————_ —————————————————1

MICROSYMBOL

GENERATOR

DISPUYPROCESSOR

i“G’TNcoLoRN’oEOY I I

{

(H)

WXSUS1 IL)TERM

{

(H)

w Bus 2 (L)TERM

{

(H)

WXWS3 (L)

TERM

{

(H)

BRIGHTNESS mT (WI

rL)

SOFTWARE ENASLEFLT TEST ONLY {

10NO.1

ID NO.2

PORT SEL A

PORT SEL B

{BuRST OUT ;

04JVALID

{

(H)

BuS 1.1 MHZ PRlM4RY (L)TERM

{

[H)

BLS 2.1 MHz, ALT 1 (3-)TSRM

{

[H)

BUS3.1MHZ, ALT2 (3.)TSRM

{

(H)

GUS 4, 1 MHz. AI-T3 iL)TERM

r‘s ‘T (L)

ALS4:

RE~TEL’’”T=”=’Ro

CXl;~~ROUNDr C)

{

(H)

WX DIMMING (W)(L)

\

Ccus

wO::oySENSOR

HV ENASLE

FOCUS CONTROL -

DTEMP FwR VALID

cu POWER DOWN 32

REMOTE LIGHT SENSOR GND 40

t

H) 53 I}

OUVOLTAGESREMOTE LIGHT SENSOR PWR

(L] E4

ou OVER TSMP 65 LOW

{

101 V*RE

28VDC 102 suPFIY103

{

104

2SVOCRTN 105

10s

BUSORwx INPUT

HIGH BPEEODIFFERENTIALINPUT TERMINATION

rm——————— ——<

[

I 1s0o REcEIVER

ilK n

TERM. >

L ——— —.—— ——— ——— ——. ——— ——— — -iAD-33252-RI #

DU-880 Display UnitBlock DiagramFigure 6-5 22-14-00

Page 153/154Apr 15/93


MAINTENANCE

Honeywell !!!k!%fh.


22-14-00Page 155

Aug 15/91Use or disclosure Of reformation on this page is subject to the restrictions on the title page of this document.

6.

Honeywell !i$!!!~.c’B. SG-884 Symbol Generator (See Figures 6-6 and 6-7, and Tables 6-2 and

6-3.)

AD-11941

SG-884 SymbolFigure

Generator6-6


Length ...................................... 14.75 in. (374.7 mm)Width ........................................ 4.91 in. (124.7 mm)Height ....................................... 7.62 in. (193.5 mm)

Weight (maximum) ................................... 15.4 lb (7.0 kg)

Power Requirements ................................ 28 Vdc, 85 WMax

Mating Connector:

J1 ....................................... DPX2MA-67S-67S-33B-0001

Mounting ....................... Tray, Honeywel1 Part No. 7003272-903

SG-884 Symbol GeneratorLeading Particulars

Table 6-2 22-14-00Page 156


The SG-884 Symbol Generator (SG) functions as the data processor for thedisplay system. It receives digital and discrete inputs, organizes thisinformation into the correct formats as defined by the DC-884 DisplayController, and transmits these formats to the DU-880 Display Units. TheSG is also used as the bus controller for the system and also outputsdata over the ASCB to other components in the system as specified inTable 6-3. Table 6-3.1 lists the ASCB transmitted data changes whenusing the -904 Symbol Generator for the TCAS and MLS option.

The SG contains provisions to drive six display units simultaneously,with four independent formats. This requirement exists to provide forredundancy in the event of an SG failure. The available formats that theSG is capable of creating include:

. Primary Flight Display Format

. Map Format● Plan Format● Compass Format● Engine Instrument Format● System Display/CautionAdvisory Format● Compacted Engine Format

The symbol generators in the display system are identical and directlyinterchangeablee. When the display system is in its normal (no failure)configuration, SG1 drives the pilot’s displays, SG2 drives the copilot’sdisplays, and SG3 drives the EICAS displays.

The symbol generators interface with the following components, and adescription of each interface is provided in the following paragraphs:

DC-884 Display ControllerFC-880 Fault Warning ComputerDA-884 Data Acquisition UnitFZ-820 Flight Guidance ComputerNZ-920 Navigation ComputerAZ-81O Digital Air Data ComputerInertial Reference SystemNavigation SensorsJoystickPZ-800 Performance ComputerLASERTRAKW

22-14-00Page 157


6. B. (1) DC-884 Display Controller Interface

The DC-884 Display Controllers provide all mode and PFD, ND, andEICAS display format select information. The symbol generatorsutilize the DC-884 for control of the PFD and ND formats as follows:

PFD/ND Source Display Controller

SG1 DC1SG2 DC2SG3 DC1 or DC2

The symbol generators utilize both display controllers on an equalpriority basis for control of the EICAS formats except forselections for the system page, Vspeeds from the FLT REF menu, FWC,DAU and AT selections. For these exceptions, the symbol generatorsonly look at DC No. 1 (DC No. 1 retransmits DC No. 2 data on ASCB).If power is lost to DC1, then the SGS will utilize DC No. 2.

(2) FC-880 Fault Warning Computer Interface

The FC-880 Fault Warning Computers provide all data necessary forthe caution advisory portion of the EICAS. This data is provided inan ASCII form over the ASCB. The fault warning computer provides upto 24 messages for display with up to 18 characters per message.

The FWC also provides the necessary data to display checklistinformation, as well as exceedance displays.

The DC-884 Display Controller defines which fault warning computeris utilized for display purposes. The symbol generators monitor thefault warning computers through comparison of checksums transmittedby each fault warning computer. A detected fault warning computermiscompare (red messages only) is displayed on the PFD as an amberFWC.

(3) DA-884 Data Acquisition Unit Interface

The DA-884 Data Acquisition Units provide radio altitude and allengine and aircraft system parameters.

The data acquisition unit broadcasts its data field simultaneouslyon both ASCB A and ASCB B. The pilot controls which channel isdisplayed by the display controller’s SENSOR mode. The symbolgenerator will disregard the channel of the data acquisition unitnot selected for display. Comparison monitoring of the two dataacquisition unit channels shall be accomplished by the fault warningcomputers.

22-14-00Page 158


6. B. (4) FZ-820 Flight Guidance Computer Interface

The FZ-820 Flight Guidance Computers provide flight director modestatus, command bar data, and PFD priority select information. ThePFD priority select information determines which PFD (pilot orcopilot) is to display command bar data. Flight director modestatus is provided on both PFDs.

(5) NZ-920 Navigation Computer Interface

The NZ-920 Navigation Computer transmits basic navigation data fordisplay as part of its primary data field on the ASCB. This dataincludes current navigation data and TO waypoint guidanceinformation. In addition to the basic navigation data, thenavigation computer supplies the following information as part ofits background data field:

● 51 Flight Plan Waypoints- 1 previous- 50 current

. 10 Immediate Area Navaids

G 9 Immediate Area Airports

. Current Position

. TO Waypoint Curved

. Holding Pattern

● Vertical Profile

Path

(6) AZ-81O Digital Air Data Computer Interface

The AZ-81O Digital Air Data Computers provide altitude, airspeed,normalized angle of attack, and vertical speed ARINC 429 informationfor display on the PFDs. The symbol generator provides comparisonmonitoring of the data, and the DC-884 Display Controller determineswhich computer’s data is displayed.

(7) Inertial Reference System Interface

The inertial reference system provides pitch and roll attitude andheading ARINC 429 information for display. The symbol generatorprovides comparison monitoring of the data and the DC-884 DisplayController determines which inertial reference sensor to utilize.

22-14-00Page 159


Hone~ell

6. B. (8) Navigation Sensors Interface


The following independent navigation sensors are interfaced with thesymbol generators via ARINC 429:

● VORG ILS/MLS● DME● ADF● LASEI?TRAK’”(optional). TCAS (optional)

(a) VOR

The VOR navigation sensors provide the symbol generators withbearing and frequency information for display. The symbolgenerators generate lateral deviation for both display andtransmission on the ASCB.

(b) ILS/MLS

The ILS navigation sensors provide the symbol generators withlateral deviation, vertical deviation, and frequencyinformation for display. The symbol generator transmitslateral and vertical deviation to the FZ-820 Flight GuidanceComputer over the ASCB.

(c) DME

The DME sensors provide distance and frequency information fordisplay. Distance information is transmitted over the ASCB forutilization by the FZ-820 Flight Guidance Computers.

(d) ADF

The ADF sensors provide the symbol generators with bearing andfrequency information for display purposes only.

(e) LASERTRAKW

The LASERTRAKN Navigation Display Unit may be configured toprovide additional outputs for display purposes only.

(f) TCAS

The optional TCAS Computer provides the symbol generators withtraffic advisories, resolution advisories, performanceresolution advisories, and map data for display.

(9) PZ-800 Performance Computer Interface

The PZ-800 Performance Computer provides autothrottle mode statusand priority select information. This information is provided onboth”PFDs. 22-14-00

Page 160Apr 15/93


uSP BIT BIT FUNCTION

II

NOTE FORNAT SCALE RSB APPROXRESOL/LS9 POS SENSE FTIU SIMVAR..- ------ . . . . . . . . . . . . . . ..-.. *----- I----lfi;;i;;------------------l--------------------l---l -----------------l ----------------- 1----1--------1--j------l:;-G;;-------------l----l;;;;~;;fi;-------------l--------------------i---l--”-------------”l---------”-------l;~;;l--------l

1 15 lEST LOGIC 1 ■ TEST1 14 VALID LOGIC 1 = VALID

5C50

1 13-11 BACKUPSTATUS 000 ● SPARE5CA0

001 - SPARE:1

010 ● SPARE

1011 = NOT BACKUP100 ● BACKINGUP SG1

11

101 ● BACKIHGUP SG2

1110 ■ BACKINGUP SG3

1 1O-B COUNTER111 = BACKINGBOTH SG’S

I 7-oO-7 HEX

EFISAOORESS 08 H - LEFT OC N- RIGHT OB H - CENTER

21I-.. ...... ......................---SOURCEIDENTIFIER

l----l---------------*---------l--------------------l---l-----------------l-----------------I;--ilFi;;iiil

2 15 FLAG LOGIC1 ● VAL1O2 14-8 DISPLAYEOATTITUOE 00 H.VG1 01 H ■ VG2

5C5202 N=VG3 5CA2IEFOATTL

2 03 H= INS 04 ti=lNS1 05 II=INS206 H=lNS307 H ● ATT1 08 N - ATT2 09 N=ATT3 OA N-ATT3

: OC H= IRS 0011■ lRS1 OE N=IRS2 OF N=lRS327-0 SELECTEOND NAVSOURCE (SEEUSP5 FORDEFINITION) lEFStlAUL

31... ...-.*I-------------------------l----l~------------------------l II.................... ---.................

SOURCEIDENTIFIERI-----------------1;;-1--------I

3 15 FLAG LOGIC 1 = VALIO3 14-8 OISPLAVEOHEAOING oAH=MGl DC H= MAG20EN~MAG3 1211+AG4

5C54

3 OBN=TRUl OOH=TRU2 0FH=TRU35CA4

llH=TIUl43 13H-OG13

14Ii● OG 2 i5N=OG3-17 }1● NOG1 18H ■ NOG2 ]911.IIDG3

16H=OG41AN=tlDG4

3 lBN - IRS ICII= IRS1 IDII=IRS2 lEII=IRS3331 SPARE

08 H= CHP1 09 N= CF!P2

36-1 PORT103

OB II= PORTA OA N = CS PORTA

3009 II= PORTB OB N ● CS PORTII

I l"-::-fl:!~:::-----------l----l:-:-fl:--:-:-::::--..--.-l-.-..----.....-....-[..-[--------- I................. ................. II.... ........

HSP 011 BIT FUNCTION NOTE FORNAT SCALE RSB APPltOXRESOL/LSBPOS SENSE FTIU SIMVAR.

“-;l------l;~;;;-i;;;i;i;;--------l----l-------------------------l--------------------l---l-----------------l------------SELECTEOSRN 000 ● SRN NOT SELECTED 5C56 IEFSSRNL4 1s-13

44444444 124 114 10-84

i444441-644

(i46-342-044444

ool.~010 ● 1011 ● P100 = 1101 = t110 ■ s111 ■ t

fts sOURcE 1 ■ ANILltN/SRN SELECTEO 1 = LRtSELECTEOLRN Ooo=l

001-1010= !011 ● (

SHNHAROUAREID

LRN BLRN A

lR/LOC 5CA6INTO VOR/LOCTRANSITION.sINTO tU.STRANSITION10SSSIOE VUN/LOC~AREIOSSSIOEUS.LN3O ■ ARINC429 FEFSHLSLSELECTED O = SRN SELECTEO FEFLRSRLINNOT SELECTED IEFSLRNLINSELECTEO(SEE 2-O))AREtOSSSIOELRN A

100 ■ SPARE101 = LRN B (sEE5-3)110 = SPARE111 = CROSSSIOELRN B00 ● ASCBol . ARINc42910. PL ;I11 . ~ALoGWo . spARE001 ● L ASCB010 = ANMOG HITH ARINC56B011 ■ ARINC561 (6 MIRE]100 ● ARINC429 HklllSPEED101 ● ARINC429 LOU SPEEO110 ● R ASCB

IEFSRNHL

IEFLRNBLIEFLRNAL

A*mo 4

751

I111 ● ANALUGtllTHARINCS61

... ...... ..........---------------SOURCEIOENTIFIEII

l----l-------------------------l--------------------l---l-----------------l-----------------lii&l---

5 15 TRANSITIONFLAG 1. TRANslTloN 0. MT TRANslT1oN5 14-B OISPLAYEDARHEONAV (SEENEXTBYTEFORDEFINITION)57-0 SELECTEDPFONAVSWtlCE NAV SWRCE DEFINITION:

imH. lNs 21 H = INS 1 22 H=INS2 23 H-INS324 II● NLS 25 H = MLS I 26 II=HLS 22? H = VLF

:28 H ■ VLF 1 29 II=VLF2 2A H=VLF3

2B H = RNAV 36 H ● RNV 1 37 II=RNV25

N 534 II● FMS 2C H ■ FMS 1 2D lt=FMS2 35 H=FNS332 H = ILS 1 33 II= lLS 2 32 H=NAV1 39 H=NAV2

N:3A H = TAC 3B H = TAC 1 3C H=TAC23D H ● ONS

t3E H = SKE

I l-------------------------l----l---------.---------------l--:!-t!!!!---------1------ ------

!i& 6“ -EFIS DISPLAYSOURCES “ ‘,

6 15-3 SPARE$~za 62 MACH/CASSELECT LOGIC 1 ● HACHa 61

~~ IDAU2 SELECT LOGIC 1 ■ DAU 21J.-

0 -:l:-----IDAU’‘ELECTLOGIC 1 ■ DAU lB

-------------------------l----l-------------------------l--------------------1---’?WQmo-m

5C585CAB I

----------------- ----------------+ ;-~1--------lI5C5A5CAA

I. . . . . . . . --------- . . . . . . . . . . . . . . . . . 1----1 --------1

tiSP 8[1 BIT FUNCTION NOTE FORMAT SCALE

71I

RSll llPPROKRESOL/f.SB POS SENSE;;~fi~;~;-;;;;i;;-;~;~~l----l-------------------------l--------------------l---i-----"-----------1-----------------1;;;;1--”-----I

FTIUSIHVAR---------

1 15-11 SPAME 5C5C1 10 IRS LOGIC I = FAIL 5CAC

Vls LOGIC 1 ● FAIL;: RAOALT LOGIC 1 ● FAIL77 BAROALT LOGlC1- FAIL76 lAS LOGlC1 = FAIL15 CO$@ARATOR140NCLEAR LOGIC1 ● CLEAR14 LOCIAZ LOGlC1 - FAIL

GS/EL LOGIC1 ● FAIL;: tlEAOINO LOGIC1 - FAIL71 ROLL LOGlC1 = FAIL70 PITCH LOGIC1 - FAILII.-....... -.--................---.-l----l;~;i~i;;[;---------------l;-:-;;;-;;;---------li;-l;-~i;;-;;;;;---l;;;;-~i;;;-;~;---l;~;;l~~;;;~;[l

8 15-1 HEAOING/TRKEIHUIR . .5C5E

80 FLAG LOGIC 1- VALIOI l;;i;i;;-;;i~;;----------l"---l;i~fi;----------"--------l~-:-;;;;-;;~;-------l;;l;:;;;;;;;;-;;----l;;;;;-;;;;;~-----l;~i;

5CAE FEFJITKEL.-. ------

9 15-49 3-1 SPARE90

5C60DISPLAYFLAG LUGIC1 ● DISPLAYED 5CB0

I l;i~;[;;;;-;;;;;;;;;------l"---l-------------------------l---"---------"------l---l----------"------l------------I--- ------. .. .

1010 15-14 TO-FHOt4OATA 00 = NONE

01 . To5C62REFTOFRL5CB2

10. FRON11 . (IN(I$EO

10 13 PFO VALIO LOGIC 1 ● VALIII10 )2 NO VALIO LOGfC 1 ● VAL1O FEFPLHoL10 11-8 SPSRES10 1 01SPLAYEDRAOALT

FEFTOFllL1.RA2:0.RA1

106 CROSSSIDEVENTCAP LOG1C1 ● CAPTURE[G/SOR EL) FEFCSVCL105 OISPLAYEODAOCL/R 1 = RIGHT O ● LEFT FEFllAOCL104 t4AP/PLANMOUESELECTED LOGIC1 ● SELECTEII FEFPLMDL103 TO-FROMFLAG LOGIC1 = VAL1O102 OUTERt4ARKER

FEFTOFRLLOGIC1 ■ BEACON FEFOTNKL

10 1 141UOLE14ARKER LOGIC1 = BEACON100 INNERMARKER

FEF140NKLLOGIC1 = BEACON

I l--------------------=-"--l----l;~~;;~[;~i;---------l--------------------l--.l.---.----..----.-l---------------------1:::!!!!!---------11 15-4 ARHEOLATOEV I113-1

+/-2911.9HICROAHPS 120.146/0.00915mA FLYRIGIITCMOSPARE

5C14REFALOVL

110 FLAG5C64

LOGIC1 ■ VAL1OI l-------------------------l----l-------------------------l--------------------l---l-----------------l-----------------l::---------

2u(D

n

t-t

-12s

s..4.

NN1

IUSPBIT BITFUNCTION NOTEFORMAT SCALE RSB APPROXRESOL/LSBPOSSENSE FTIUSIHVAR

‘::’~5i4‘M$;:R:ERTOEVII..---.... --------------------------------------------------------------------------... -----------------1 II I-----------------l;:;;l;;;fifiI

TUO’S COMPLEMENT +/-399.2NICROlWIPS12 0.195/0.0122mA FLYUPCND5C66

120- FLAG LOGIC1 ■ VALID 5CB6FEFAVOVL,3115-4l;;[~;f;~~~;;~-~;fi--”-l ISE”,C*RCLE... ------ -----------------------------I I ~21oeo*78,0,m549I----------------------------------------~~-----------l;~i;l~;~~~~~[l

0-360 DEGi3 3-1 SPARE 5C68130 FLAG LOGIC 1 = VALID 5CB0 FEFSCRSLI l;;-;~fifi--------------l----lsEH*clRcLE..------- I---------------------------------------------l-i;l-;:;;;;;;-~;;;--l~;;;-------------l;;i;i;;;;;;~:l

14 15-4 0- 360 DEG .:! ; SPARE 5C6A

TUNE-TO-NAV LOGIC 1 ■ TUNED 5CBAFEFTTNL141 TUNE-TO-LOC LOGIC1 ● TUNED FEFTTLL140 BEARINGFLAG LOGIC1 ● VALID FEFVURGL xII... ------ II I--.-.--.*-------------------------------------------------------------------------------------11 I-----------------liiiiliiii;iil o

::~5-4 OISPLAYEOLATIJEVIATION3 TUO’S CONPLE14ENT +/-29E.9HICROANPS 12 0.146/0.00915mA FLYRIGIITC140BEARING0 FLAG LOGIC1 = VM1O SEENOTE 3 SEE NOTE 3 5C6C FEFBNEOL a

152 BEARINGO FLAG LOGIC 1 = VALID 5CBCFEFBEqOL!151 ACTUA1/CALCULATED ~ . calculate 0. AcWAL FEFACBGL a150 LAT DEV FLAG LOGIC 1 ● VALID FEFLDEVLII..------- 11 I..------------------------------------------------------------------------II I----0---------------.................l;~-&-:”1

16 16-4 DIj~~LA;EDVERT OEVIATION 3 TUO’SCOMPLEMENT +/- 399.2MICROANPS 12 0.195/0.0122mA FLY UP CHll 2:::1 SEE NOTE 3 SEE NOTE 3 5C6E

FLAG LOGIC 1 ● VALIO 5CBE FEFVDEVL ‘mII..- -------------------------------II I---- ---------------------------------------------l-fil-i:~;:&;;;;---l~f;;;~;l;;;;ll”;l;l--”-----l =

:; :5;4 Dl:~:$:EDRAOIOALTITUOE TUO’S COWLE*NT +/- 2047 FEET5C70

170- FLAG LOGIC 1 = VALID 5CC0c)~~

II II-.. ...... ----------------------------- ------------------------------------------------ ----------------------------------1 II I 1;;;;1--------1 &18 OESIGNATORLAT 118 15-12 TEN’SOF MINUTES eco18 11-8

5C72 4D~MINUTES BCD 5CC2

frz

181-4 TENTH’S OF HINUTES BCD183-2 SPARE ~g181 POLARITY180

~ . s@jTti0. NORTHOATAVAL1O(LAT1 S LAT2 ) LOGIC=lDATAINVALIO m

IA

II---.-------------------------------------------------------------------------------I 1---1 I 15C241 I19 ESIGNATORLAT2...........------................. ----........

1915-12 HUNOREO’S OF DEGREES BCD 5C1419 11-8 TEN’S OF OEGREES BCD191-4

5CC4DEGREES BCO

193-1 SPARE190 ENTERBUTTONPRESSED L~lC 1 ● PRESSEOII II..-*....- ......................... ---------------------------------------------------------------------................. ----1 II I 15C261--------I20 DESIGNATORLDN 1

20 15-12 TEN’SOF tYINUTES BCO2011-8

5C76NINUTES BCD

201-4 TENTH’S OF MINUTES Bco5CC6

203-2 Ft4sSEL 00. FMs NOT sEL (BIT382)20 01 = LEFT FNS SEL20 10. R[GHTFMs sEL20 11. uNoEFINED201 POLARITY 1 ■ HEST O ● EAST20,0 DATA VALID(LON 14 LON,2 ) ,LOGIC=1 DATA INVALID

t..- .---------------------------------- ---------------------.....t I 1-------. -.-.--..-1 . . . .. .&... ----.-l ,-:1 .- I.....~,.--- ---------- ---

USP BIT BIT FUNCTION NOTE FORMAT SCALE RSD APPROXRESOL/LSBPOS SENSE FTIU SIMVAR---l------l;~;;;~~;;;-i~~-;---------l----l-------------------------l--------------------l---l-----------------l-----------------l;i2121 15-12 HUNDRED’S OF DEGREES Bco21 11-8 TEN’S OF DEGREES UCD

5C78

21 )-4 DEGMEES BCD5CCB

21 3-o SPARE---1------l;~-~;;;~~;;-------------l----l-------------------------l--------------------l---l-----------------l-----------------liii

:; 15 FLAG LOGIC 1 = INVALIDDATA 5C7A FEFINIEDL22 14 RANGE 1 “ 0-3999.0 0 ■ o-399.9 5CCA FEFDMERL22 13-12 IIUNDMED’S/TtlOUSAND’S BCD22 11-8 TEN’S/HUNDKED’S IICD22 1-4 ONE’S/TEN’S Bcl)223-0 TENTWsjoNi’s BCilII---------......................---l----l;;;;;;;:;~---------------l---------"----------l;;l;-;;;;;-~;;;""-l;;;;;-;;;;-;;;;-l~;;;l-------

23 15-4 01SPLAYEIIIIOLL O - 360 OEG23 3-1 SPARE

. .

230 FLAG5C7C

LOGIC 1 = VALIIIII

5CCC... ...... -------------------------l----l;;fi;;;;;;---"-----------l;-:-;;;-;;-"-----"-l;;-l;;;;;;-~;;;"--l-----------------l;;;;l------24 1S-4 DISPLAYEOPITCH24 3-1 SPARE

. . NOSEUP5C7E

240 FLAG LOGIC 1 - VALID... ------..-----------------------l----l;;fifi;;”;--”------------l:--;;;~~---------l;;l;:;;;~;;:~;;;---l;;;;-------------I:;;liiiiiiiil25115-4‘COURSEERROR(SRN) .25 3-1 SPARE 5CB0250- FLAGII

LOGIC 1 = VALID... ...... .........................l----l;;;;;;;;”;---”-----------I26 15-4 DISPLAYEDHEADING

;--;-&;-------”-l;;-l;-;;;;;;-;;;---l;;;~-------------I:;;I::;:LI

263-1.

SPARE,

260 FLAG LOGIC 1 ● VALID5C82

II..- .------------------------------l-"--lfi;;;Fiii;ii---------l--------------------l---l-----------..----liiiii-------.----lol!!!!!!!!l27 15-4 OfSPLAYEDVERTICALSPEED +/- 65536FT PER 141N15 2.0/1.0FPM21 3-1 SPARE 5CL14210 FLAG LOGIC 1 ● VALID27 (OR MAINTENANCETEST)

5CD4

II.-. --------------------------.....1----l;~;;-~;;~;~;;---------l-"------------------lii-li-i;-i-iiii;-----[------------.-..-1---.1---.--..128 15-6 MLS SELECTEOELEV ANGLE ~ 51.2DEG285 SELECTEDELEV ANGLEFLAG LOGIC 1 = VALID

● . 5C36

284 AZ SOURCE LOGlCO ■ AZ5C86

28 3-O MLS GROUTH LOGIC 1 - BACKAZ5CU6

II--- ----.--------------------------l----l-------------------------l--------------------l---l-----------------[--------------.--l----l--------l29 15-11 MLS GROUTH29 10-5 APP AZ TO THSIIHOLDDIST BINARY

5638

29 4 01STANCEFLAG0-6300 METERS 6 98.4/0.096

LOGIC 1 = VALIO5C88

29 3-o SPARE5c00

l-- I----------------------------------30 15-2 OISPLAVEOAIRSPEEO

1----l~~;;-;;;;i~;~~;---------l;;--;;;;-;;;;;--"---l;;-l;-;;;;;:;;;;;----l~;;;;;i----------l;;;~l--------[30 1

-SPARE

.61WA

300 FLAG LOGIC 1 - VALIO--“..5CDA

I l-------------------------i----l~~;;;-~;~;i;j~ii---------l---------31 15-1 DISPLAVEOBARU ALT

;;--;;;;;-~~;;------l;~-l;-;;;:~-~;-------l-----------------liii;l--------l

310 FLAGUP

LOGIC 1 = VALIO●

5C8C

I l-------------------------l----l-------------------------l--------------------I---l... ------ 5CDCI.......---------------------------1----1--------1

WP BIT BITFUNCT!ON NOTEFORMT SCALE RSB APPROXRESOL/LSDPOSSENSE FTIuSIHVARI lfi;;;-~~----------------l-"--l~H-~~~s-------------l-----------------"--l---l-----------------l-----------------l;~;~l---.-.......

325C8E5CDE

I l;;;~&~;;--------------l----l;:---------"-----"-----l-"------------------l---l------------""---l-------"---------l;~;l--------l... ......33

5C90

I l~;;~~----------------l--"-l;;-~;----------"--------l--------------------l""-l-----------------l-----------------l--"-l"----5CE0

..-------34I l-;j----------------------l----l-------------------------l--------------------l---l-----------------l-----------------l--.-.----.-NOTES: uSE UNROUNOEOSELECTEbHEAI!ING ANO SELECTEO COURiEIN THE CALCULATION‘OF:‘

., .

NoGSELECTBUGDISPLAYEDPOSITIONCRSSELECTRUGOISPLAYEDPOSITIONLATDEVPOINTEROISPLAYEOPOSITIONNSPII“lllIG/TllKERRON”UiP i3 “SEL’OCRS SRN o

IIUSP 25 “CRS ERROR SRN “NSP 15 “oISPLAYEO LAT l)EVO (VOR)

2)

3)

NNENHLS IS 01SPLAYED,USESELECTEOAZIMJTHAS RECEIVEOSELBUGPOSITION& USP258 NSP 13

FORARINC429NLS,Ie (HSP4, BIT12=0)ANO(USP4, BITSFACTOR, RSEl, ANORESOLUTIONARE:

SCALE RSII APPROXRESOL/LSfl

~ 2400UA 15 (BITS15-1] 0.146/O.073242

(1.SB=l.0 OEG) FROM NLS CWRSE

15-13● 011 OR111) THESCALE

n

$wo

w

WSPBIT BIT FUNCTI@4 ROTEFORMAT SCALE RSBRESOLUTI(XVL5B Pt3SSENSE FTIU SIMVAR--- l--.. -- !------------------------- l . . ---------------------------- l- . . ..--... - . . . . . . . ..j... ~... - . . . . ..- . . ---- l . . .._. ~_-.. __. -~.._.... I

4 ‘ 15-13 SELECTEDSRNI

000 = SRNNOTSELECTED ‘001 = MLS010 = LRN TO VUR/LOCTRANSISTICN011 = SPARE100 ❑ LRN TO MLS TRANSISTICN101 = CROSSSIDE MLS TRANSISTI~110 = SPARE111 = SPARE

--- I--. --. f.. ----------------------- 1... - I .-.. - . . . ----------------- ~-..-.---..-..--....-l--. ~----------------------------------- ~_~_-_.l

5 ‘ 15-8 PREVIEWNAV SUIRCE ‘ (SEE NEXTBYTE FU7 OEFITION)I

NAV SCLRCEDEFINITION5 7-0 SELECTEDPFO NAVSUJRCE 20 11 = INS

24 NOTUSEO27 11 = VLF

. . . I . . . . .. !------------------------- ~------------------------------ l-------- . . . . . . . ..-.. j --------------------- l.- . . . . ..--- . . ..-. ~-... j..-..... I

6’15 NOVALIO LOGIC 1 = NOVALIO14-8 SPARE7 MLS2 FAIL LOGIC 1 ❑ MLS2 FAIL6 MLSI FAIL LCM31C1 = MLS1 FAIL5 SPARE

TCASFAIL LCGIC 1 = TCASFAIL6; TOGAWE LOGIC 1 = TOGAACTIVE

--- l------ f ------------------------- ~------------------------------ I.. --... -.--. - . . . . . ..l. -. l . . . . . . . . . . . . . . . . . ~.. -... - . . ..-... -.l.. --l -------- I

7’15-12 SPAREI II I I

11 MLS LOGIC 1 = FAIL73 GS/GP LCKIC 1 = FAIL

. . . l . . . . .. ~. . . . . . . . . . . . . . . . . . . . . . . . . l ------------------------------ : ------------------------------------------ ~.. -.---- . ..-. -.-. ~. . ..~.. -.....l

11 ‘15-4 fMUDE PREVIEWNAVI I

..- /-. -... /.-_._... ..-. _.. . . . . . ~------------------------------ l ------------------------------------------ ~.. -.-. -... - . . . . ..l-...l -------- I

1215-4 CWSIDEADFBRGI I

nI I I I. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . l-. --. -.-. --. - . . . . . ..~.. -{-------I I I I

---------- ~-------- ..-... --. {-.. -[-. --. -..lI

BITA5 1

(16)>

(16)>

(20)>

(7)e

(6)~

(14)-

DISCRETE A9

I # r 1 I

_l GFm,cwm hl ADDRESSR DECODER kll“L-Eu-_l

r I Hll

r-----d 1111

~ IBUFFER BUFFER

- I

IIIII

4+ WxMONITOR

(+ DISCRETERECEIVERS

TIMINGMONITOR

t-iCLOCK

GENERATOR t--t

1:1

I I

w

16BIT BIDIRECTIONALDATALOCAL BUS

I

ARINC Al

—14BIT + * 7 DUAL

2 TO 1 MUX LINE 6 BIT CHANNEL

* (12T06 RECEIVERS 429/419PLUS 2 (6) MUX 429

ARINC 429) + + SERIAL Bus

10 BUFFER

RECEIVERS

I L--Jl

t--d, I I II

RAM2k X 16BIT Id

I I

Ill ASCB A2

L.uBUFFER

RAM (MAILBOX) BUFFER

n> L

GDIRECTMEMORYACCESS

60166 RAM (LOCAL) PROM (PROGRAM)CPU SkX 16BIT 16kX 16BIT

L

I

- I (LOCAL) k II

16BIT BIDIRECTIONALDATAGLOBAL BUS

RBUFFER

ISERIAL 1/0 + MANCHESTER ~

CONT ASCB CODERl~ DECODER -

MUX ANDBUFFER

ASCB 1

ASCB 2

AAHRS-232

AD.15187 @ .R1

SG-884 Symbol GeneratorBlock Diagram

Figure 6-7 (Sheet 1) 22-14-00Paae 167/168- -a.

Aug 15/91Use or disclosure of mformatlon on this page IS subject to the restrlcttons on the Iltle page of this document

I IRAM RAM

16 BIT OATA ADDRESSBIDIRECTIONAL POINTER

‘2k Xee11

IP

< BUFFER 60166 LOCALCPU

3UFFER

16 BIT BIDIRECTIONALBUFFER

DATA GLOBAL BUSSERIAL~

+C~UNl-lMTION

4 C~LERADDRESS STATEDECODER MACHINE

6 BIT UNIDIRECDATA GLOBA

RAM PROM16K X 16 BIT 32k X 16 BIT

CPLI A1O

1-1

BUFFER 60166 LOCALCPU BUFFER

16 BIT DATABIDIRECTIONAL

WX IN -

v LOCAL BUS 2

RAM PROM16kx 16 BIT 32kx 16 BIT

I !CPU A12 I

l—

-1MANCHESTERENCODER

I

-~~CHANNEL 1

--— -—.. — .—-— -—-— -—-— ——.

SAME AS CHANNEL 1 L DISPLAY UNIT

DISPLAY UNIT INTERFACE Al1CHANNEL 2

CHANNEL 2

+ ADDREss I

DISPLAY UNITCHANNEL 1

vDATA + STATE

ADDRESSGENERATOFiI RAM

LATCH MACHINE 7 BUFFER 16kx4BlT,

t,~1

I I I I 1I

WX INTERFACE A13

I BUS [

‘26’”-Blls A

BIJS B

4DIFFERENTIAL MONITOR

LINE + MICRO4 ~ DRtVER

Im CODE

+

+MoNtToR

MONITOR

DRIVER VLslDATA BASE

+<+ PROM

+ , I#?

+ & — — TRANBMllTER +TRANSMITTER

4DRIVER

VLslDATA BASE

+ PROM

STANDARD BUS CONTROLLER M

I 1

I POWER SUPPLYr-

CLEAR

CONTROL CIRCUITS

I A6 00K AHEAD

+ ++ WX OUT+ POWER SUPPLY ++5 VDC

+26 VDC~ TRANSFORMER CIRCUITS ++15VDCA7 +-15 VDC

SG-884 Symbol GeneratorBlock Diagram

Figure 6-7 (Sheet 2) 22-14-00Page 169/170

Aug 15/91Use or disclosure of mformallon on Ihm page is subject 10 Ihe restrtctlons on the title page of this document.

MAINTENANCE

Honeywell !!Hi!+A&.


22-14-00Page 171

Aug 15/91Use or disclosure of information on this page is subject 10 the restriaions on the title pege of this document.

6. C. DC-884 Display Controller (See Figures 6-8 through 6-23, and Tables 6-4and 6-5.)

DC-884 Display ControllerFigure 6-8


Length ...................................... 10.58 in. (268.61 mm)Width ....................................... 13.02 in. (330.71 mm)Height ........................................ 2.78 in. (70.61 mm)


Power Requirements:

Primary ........................................... 28 V dc, 35.0 WLighting .................................. 5 V ac ordc, l.OWmax

Mating Connectors:

J1 ................................................. MS3126F22-55SWJ2 ................................................. MS3126F18-32SW

Mounting ......................................... Unit Dzus Fasteners

DC-884 Display ControllerLeading Particulars

Table 6-4

22-14-00Page 172


MAINTENANCE

Honeywell M!!!fi!i!.The DC-884 Display Controller (DC) provides the flightcrew with a concisemethod with which to control the entire electronic display system. Thecontroller consists of 10 lighted function keys, each with an “ON”annunciation, and a CRT, which, along with 10 line-select keys, providesthe medium for menu selection. The CRT and the function keys aresunlight readable. In addition to the function keys, the controllercontains three knobs which are used to set the value of a menu parameter(SET), set the barometric pressure (BARO), and adjust the CRT brightness(BRT). The display controller compares inputs from the display units anduses the larger value to program brightness.

Course select information from the flight guidance controller isconditioned, digitized, and routed to the symbol generators via the ASCBbus. This information is used by the SGS to calculate left-rightdeviation for display and the flight guidance computer.

The GIV system is configured to operate with two DCS installed in thecontrol panel, directly in front of each pilot, above the primary EFISdisplays. Figure 6-8 shows the layout of the display controller. Eachdisplay controller is comprised of the following functions:

. Baro Set Control

. Menu Driven CRT Display

. Display Controller CRT Brightness Control

. 10 Menu Defined Line Select Keys

. DC Menu Parameter Set Control

. 10 Display Function Keys

- MAP Mode- COMP Mode- PLAN Mode- NAV Mode- SENSOR Mode- FLT REF Mode- TRS Mode- SYSTEM Mode- TEST Mode- DISP Mode

The following paragraphs describe each mode controlled by the functionkeys. Digital outputs transmitted on theASCB are listed in Table 6-5.

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6. C. (1) DC Modes - General

Before discussions of the specific DC modes can begin, somediscussion on the display conventions used must be given to enhancethe descriptions that follow.

(a) DC Parameter “Selected For Display” - A selected parameter i.s

displayed on EFIS/EICAS only when the item is boxed on theappropriate menu. The absence of a box means that the item isnot being displayed.

(b) DC Parameter “Selected For Set” - The first selection of a lineselect key by a settable parameter causes the parameter to beshaded (i.e., inverse video). When inverse video is displayed,the DC set knob can be used to change the value accordingly.In the normal set sequence of a given parameter, “selected forset” is followed by “selected for display”.

(2) DCMenu Declutter Mode

The DC power-up state is the menu declutter mode as shown in Figure6-9. This mode may be selected at any time by depressing the active(annunciated) function key. Selection of the declutter option hasno effect on any of the menu selected items.

(3) MAP Mode

Generally, the MAP function is comprised of the following items:

. Bearing Pointer Source Selection

. Flight Plan Declutter Options● Wind Display. Vertical Profile. Optional TCAS (Refer to paragraph 6.C.(15)(b).

(a) MAP Function Key Selection - Selection of the MAP mode causesthe function key to be annunciated along with the MAP menu tobe displayed on the DC CRT (Figure 6-10). If the ND is in theCOMP (compass) or PLAN modes, the ND will configure to the MAPdisplay format.

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MAINTENANCE

Honeywell !%!$AA.

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Honeywell !!!!!!~.c’6. C. (3) (b) Bearing Pointer Source Selection - The first row of this menu

is dedicated to selection of the bearing circle sources. Thesecond row of the menu is comprised of the bearing diamondsources. Selection of the left line select keys move the boxacross the row (left to right) through the bearing sources.With the box on the extreme right-hand source, the nextselection of the left line key, removes the box from the menu.This is how bearing pointers are removed from the display(PFD/ND). Another selection of this key, will bring the box upon the left side of the row. The right line select keys in thefirst two rows operate the same except that the box moves fromright to left.

Selection of the AUTO mode forces the particular bearingpointer source to be the same as the active navigation source.The pilot’s AUTO selection appears on the bearing circle (No. 1sources) pointer only as shown in Figure 6-10. The copilot’sAUTO selectionpointer only.

The bearing po

appears on the bear

nter power-up defau

ng diamond (No. 2 sources)

ts are:

Pilot - AUTOCopilot - AUTO

(c) Waypoint Declutter Mode - The selection for this mode isdisplayed on the DC CRT as ID WAYPT. This selection has threeindependent modes as described:

I ID WAYPT -

mIDWAYPT -

IDWAYPT -

This is the power-up mode for this parameter.With the box surrounding the entire mode, theactive flight plan waypoints and IDs aredisplayed on the MAP display format.

This is the next selection of the line key.With the box surrounding only the WAYPTlabel, the active flight plan waypoints are~;:~):yed without IDs on the MAP display

.

This is the next selection of the line key.With both the ID and WAYPT labels unboxed,the active flight plan waypoints and IDs arenot displayed on the MAP display format.

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MAINTENANCE

Honeywell !!!W!4A&.6. C. (3) (d) Navaid Declutter Mode - The selection for

displayed on theDC CRT as ID NAVAID. Thindependent modes as described:

ID NAVAID -

EEl-

this mode iss selection has three

This is the Dower-uD mode for this Darameter.With both the ID and NAVAID unboxed; theactive flight plan navaids and IDs are notdisplayed on the MAP display format.

This is the next selection of the line key.With the box surrounding the entire mode, theactive flight plan navaids and IDs aredisplayed on the MAP display format.

ID NAVAID - This is the next selection of the line key.With the box surrounding only the NAVAIDlabel, the active flight plan navaids arecll~~~:yedwithout IDs on the MAP display

.

(e) Airport Declutter Mode - The selection for this mode isdisplayed on the DC CRT as ID AIRPT. This selection has threeindependent modes as described:

ID AIRPT - This is the power-up mode for this parameter.With both the ID and AIRPT unboxed, theactive flight plan airports and IDs are notdisplayed on the MAP display format.

I.w!Y_l- This is the next selection of the line key.With the box surrounding the entire mode, theactive flight plan airports and IDs aredisplayed on the MAP display format.

uID AIRPT - This is the next selection of the line key.With the box surrounding only the AIRPTlabel, the active flight plan airports are$:~~~:yed without IDs on the MAP display

.

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Honeywell !$!!!~.c’6. C. (3) (f) WIND Display Selection - The WIND format can be selected for

display in either the vector or X-Y format as shown in Figure6-10. The power-up default is X-Y format. By selecting theline key, the box moves left to right as shown below:

First SecondPush Push

r

> XY > VECTOR > OFF

1Third Push

(9) Vertical Profile - The vertical profile mode can be selectedfor display on the MAP mode format. The power-up default forthis mode is VERT PROF selected for display (boxed). Thisfunction is an alternate action selection on the DC (i.e.,boxed/unboxed).

(h) MAP Selections General - Bearing pointer and wind displayselections made on this menu also change the selections on theCOFIPmode menu, and will be reflected on the COMP display ifsubsequently selected. Flight plan declutter selections madeon this menu are also changed on the PLAN mode menu, and willbe reflected on the PLAN display if subsequently selected.

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MAINTENANCE

Honeywell !i!!!!!%h.6. C. (4) COMP Mode Menu

The COMP function is comprised of the following items:

● Bearing Pointer Source Selection. Wind Display

(5)

(a)

(b)

(c)

COMP Function Key Selection - Pressing the COMP key causes thefunction key to be annunciated along with the COMP menu to bedisplayed on the DCCRT (Figure 6-11). IftheND is in the MAPor PLAN modes, the ND will configure to the COMP displayformat.

Bearing Pointer Source Selection - The bearing pointer sourceselection operation is the same as the MAP mode discussed inparagraph 6.C.(3). Selections of bearing sources on this menualso change the selections of bearing sources on the MAP menu,and will be reflected on the MAP display if subsequentlyselected.

WIND Display Selection - The WIND display format selectionoperation is the same as the MAP mode discussed in paragraph6.C.(3). Wind display selections made on this menu also changethe selections on the MAP mode menu, and will be reflected onthe MAP display if subsequently selected.

PLAN Mode Menu

The PLAN function is comprised of the following items:

. Flight Plan Scroll - FORE/BACK

. Map Declutter Options

. Wind Display

(a) PLAN Function Key Selection - Pressing the PLAN key causes thefunction key to be annunciated along with the PLAN menu to bedisplayed on the DCCRT (Figure 6-12). IftheND is in the MAPor COMP modes, the ND will configure to the PLAN displayformat.

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I

I

6. C. (5) (b)

(c)

(d)

Flight Plan Scroll - The pilot and copilot have the ability toscroll the active flight plan in either the forward (FORE) orbackward (BACK) direction. The line select key may be singlestepped (one push at a time) or held down for a slew of theflight plan. A box is drawn around the BACK or FORE labels asappropriate as long as the line key is held down.

Flight Plan Declutter Options - The ID WAYPT, ID NAVAID, and IDAIRPT operation is the same as the MAP mode discussed inparagraph 6.C.(3). Flight plan declutter selections made onthis menu are also changed on the MAP mode menu, and will bereflected on the MAP display if subsequently selected.

WIND Display Selection - WIND can be selected for display onthe PLAN mode format. The power-up default for this mode isWIND not selected for display (unboxed). This function is analternate-action selection on the DC (i.e., boxed/unknown).The format of the WIND display is fixed on the PLAN modeformat.

(6) NAV Mode

The NAV menu, as shown in Figure 6-13, is comprised of the followingfunctions:

● Active Navigation Source Selection

- Flight Management System (FMS) 1 or 2- NAV 1/2 (VOR or ILS)- MLS 1/2 (Option)- ILS 1/2 (Option)- LASERTRAKW (LTRK) (Option)

. Preview Mode Source Selection

- FMS 1/2- NAV 1/2- MLS 1/2 (Option)- ILS 1/2 (Option)

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6. C. (6) (a) NAV Function Key Selection - Selection of the NAV mode causesthe function key to be annunciated along with the NAV menu tobe displayed on the DC CRT (Figure 6-13). The NAV modefunction is used to select the active navigation sourcedisplayed on the PFD. A NAV function key selection has noeffect on any of the display formats. NAV source selectionsare made using the menu driven line select keys and controlonly the on-side displays.

(b) Active Navigation Source Selection - The first four lines ofthe left side of this menu are dedicated to selection of thePFD navigation source. Selections in this mode effect the on-side displays only. The power-up default in this mode is asfollows:

Pilot - NAV 1Copilot - NAV 2

There is always one source selected for display. Alternateactivations of the same line key moves the box between the 1and 2 selections of the same source. Selection of a differentline key moves the box to the number 1 or 2 source of thatline.

All of the possible sources are shown in Figure 6-13. The DCmust be wired in the aircraft to get the MLS, ILS and LTRKselections shown on the NAV menu. Although the aircraft may bewired to show MLS on the NAV menu, the optional TCAS/MLS DC-884Display Controller, Part No. 7007540-941/942,must be used toselect MLS as the active NAV source. See paragraph 6.C.(15).

(c) PREVIEW Mode Source Selections - When the PREVIEW line selectkey is pushed, the PREVIEW submenu is displayed on the DC CRTas shown in Figure 6-14. The power-up default of this mode isall sources deselected. The NAV sources on this page areidentical to those on the main NAV menu. Multiple activationsof the same line select key moves the box left to right asshown:

First SecondPush Push

r>OFF >1 >21

Third Push

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IWJ;:PANCE

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Honeywell !!!!!!!by’Selection of a short range NAV source (NAV, ILS, or MLS) bringsup the CRS SET window adjacent to the NAV 1/2 selection. Thewindow is initialized to O degrees. The window comes up shaded(inverse video) and boxed. The pilot can then change thecourse of the previewed source through the DC parameter setknob. The line key adjacent the CRS SET window has nofunction. This window is only shown (activated) when a shortrange NAV is being previewed. If the previewed source is anFMS, desired track is automatically set by the FMS and the CRSSET window is removed from the preview menu.

The RETURN key brings up the main NAV menu. The previewselection, on the main NAV menu, is boxed if a NAV source isselected for display on the submenu. To clear a PREVIEWproblem, all sources on the submenu must be deselected.

6. C. (7) SENSOR Mode

The SENSOR mode function is used to change the normal display sensorconfiguration in the event of failure conditions. The SENSOR menuis shown in Figure 6-15.

(a) SENSOR Function Key Selection - Selection of the SENSOR modecauses the function key to be annunciated along with the SENSORmenu to be displayed on the DC CRT (Figure 6-15). The SENSORmode menu is used to select the active sensors displayed on thePFD and ND. A SENSOR function key selection has no effect onany of the display formats. SENSOR source selections are madeusing the menu driven line select keys.

(b) Sensor Source Selections - The following source selectioneffect the on-side displays only:

. Inertial Reference System (IRS) 1, 2, or 3 (Option)

. Digital Air Data Computer (DADC) 1 or 2

. Radio Altimeter (RADALT) 1 or 2

The following selections are controlled by either side displaycontroller:

. Flight Guidance Computer (FGG) 1 or 2● Data Acquisition Unit (DAU) 1 channel A or B. Data Acquisition Unit (DAU) 2 channel A or B. Fault Warning Computer (FWC) 1 or 2. Autothrottle Computer (A/T) 1 or 2

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There is always one source selected for display. Alternateactivations of a line key moves the box between the 1 and 2(and 3 if applicable) selections of the source. The pilot’sand copilot’s power-up defaults are as follows:

PILOT COPILOT

IRS 1 IRS 2DADC 1 DADC 2RAD ALT 1 RAD ALT 2DAU 1 A DAU 1 ADAU 2 B DAU 2 BFWC 1 FWC 1A/T 1 A/T 1

The FGC power-up selection is dependent on which FGC is poweredup first. The selection will be the same on both pilot’s DCS.

All of the possible sources are shown inmust be wired in the aircraft to get theon the SENSOR menu.

6. C. (8) FLT REF Mode

The FLT REF mode function is comprised of thereference data (Figure 6-16):

. Vspeed Reference Bugs

- Vl, VR, V2, VFS, VSE, VREF

Figure 6-15. The DCIRS 3 selection shown

following settable

. Performance System Auto Vspeed Selection Mode

● Radio Altitude Set

. AOA Reference Bug

(a) FLT REF Function Key Selection - Selection of the FLT REFmodecauses the function key to be annunciated along with the FLTREF menu to be displayed on the DC CRT (Figure 6-16). The FLTREF mode menu is used to set and select, for display, importantreference data on the PFD. A FLT REF function key selectionhas no effect on any of the display formats. FLT REF dataselections are made using the menu driven line select keys andDC parameter set knob.

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v, : 124 RAD ALT: [mlVFS : 130 AUTO VSPDIv SE : m v REF : 130

AD-20868

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6. C. (8) (b) Automatic Vspeed (AUTO VSPD) Mode Operation - The power-updefault for Vspeed selections is to the AUTO VSPD mode. Inthis mode, Vspeed information is automatically displayed on thedisplay controller from the FMS performance computer. Thesource of this data is the on-side FMS except when thecross-side FMS is selected as the active navigation source onthe NAV menu. The source of the Vspeed information isdisplayed on the DC menu as an AUTO VSPD1 or 2 as appropriate.

If the Vspeed data is valid, and the aircraft configurationmatches the selections made during FMS performanceinitialization, the Vspeed information is boxed and displayedappropriately on the PFD. After takeoff, VREF is automaticallyselected by the FMS for display (takeoff Vspeeds are deselectedat this time).

If the data is invalid, the digital readout on the DC menu isshown with dashes. If the aircraft configuration does notmatch the performance computer configuration, the Vspeeds aredisplayed with an asterisk but are not boxed on the menu. Bothconditions will result in the amber VSPD annunciation.

Selection of any line select key adjacent to a Vspeed willcancel the AUTO VSPD mode on both pilot and copilot DCS. Thevalues remain displayed and can set manually as discussed inparagraph 6.C.(8)(C).

(c) Manual Vspeed Set Function - The first selection of a line keycauses reverse video to appear around the parameter. The valuecan then be set between the ranges of 80 and 250 knots. The DCallows the values to be set and displayed with the followingrestrictions:

. The Vspeed order is always maintained as follows:

- VI, VR, V*, vF~,v~E

. VI can never be set higher than V~, and V~ can never be setlower than V1.

. Vz can never be set any closer to V~ than 4 knots.

● VF~ can never be set any closer to V2 than 4 knots.

● V~E can never be set any closer to V~~than 4 knots.

● VREFset for display cancels (unboxed) all other displayedVspeeds.

● Displayed V~EFis cancelled (unboxed)with the selection fordisplay of any one of the other Vspeeds.

. Vspeeds are settable from either side DC.

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6. C. (8) (d) AOA Reference Bug - The%VS line key causes the%VS parameter. The AOA

(9)


first selection of either the AOA orinverse video to surround the AOA andand %VS are set together using the DC

parameter set knob. The range of set values is 0.2 to-0.59 AOAand 1.5 to 1.3%VS. For values greater than 1.5% VS, the %VSwindow is blanked. The second selection of either line selectkey boxes both the AOA and%VS. Only the AOA value isdisplayed (the%VS setting is for reference only) on the PFD.The power-up default for these parameters is 0.59 AOA and 1.3%VS. The selection for display affects the on-side PFD only.

(e) RAD ALT Reference Select - The first selection of the line keycauses inverse video to surround the last set value. The valuecan then be set in the range of O to 2500 feet. The selection,for display, of this parameter effects only the on-side PFDformat.

Thrust Reference System (TRS) Mode Menu

The TRS mode is comprised of the following selections and settablereference data (Figure 6-17 and 6-18):

. Selection of Performance System Computed EPR Limits

Takeoff (TO) or Go-Around (GA) EPRReduced Takeoff (FLEX) EPRClimb (CLB) EPRCruise (CRZ) EPRMaximum Continuous (MCT) EPR

● Performance System AUTO EPR Selection Mode

. MAN EPR Limit Set Mode

. DUAL/SPLITMAN EPR Limit Set Capability

(a) TRS Function Key Selection - Selection of the TRS mode causesthe function key to be annunciated along with the TRS menu tobe displayed on the DC CRT (Figure 6-17). The TRS mode menu isused to set and select for display important reference data onthe engine display. A TRS function key selection has no effecton any of the display formats. TRS data selections are madeusing the menu driven line select keys and DC parameter setknob.

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MAINTENANCE

Honeywell WB!%L.

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6. C. (9) (b)

(c)

(d)

TRS Menu General - The main menu (Figure 6-17) is used toselect the displayed performance computer calculated EPR data.The submenu (Figure 6-18) is used to set and select the EPRlimit manually. Both the pilot and copilot displays areidentical on this menu. Selections made on either DC are shownon a last entered basis.

EPR Limit Source Annunciation - Data displayed on the main TRSmenu is sourced from the priority PFD commanded and FMSperformance computer. The data source is labeled on the menuwith a 1 or 2 designator following the AUTO selection. This isthe source of the data regardless of the selected mode.

FMS (PZ) EPR Limit Mode Selections - EPR limit data isdisplayed on the main menu of the TRS mode. The data islabeled as follows:

Mode Label

Takeoff or Go-Around TO or GAClimb Mode CLBCruise Mode CRZMaximum Continuous MCTReduced Takeoff FLEX

The reduced takeoff EPR, labeled FLEX, is displayed only whenactivated on the CD-81O Control Display Unit. This value istreated as a target on the engine display (green bug) asopposed to a limit (white tickmark). When selected, the FLEXtarget value and the TO limit value are both selected.

The power-up default is AUTO mode. This mode selects theappropriate rating based on phase of flight and is accomplishedby the priority FMS-PZ. The AUTO remains selected in thismode, and the active rating is also selected (boxed) on themenu. Manually selecting a rating on this menu is accomplishedby depressing the appropriate line select key. This cancelsthe AUTO mode. Invalid data is shown by displaying a dashedline through the digital value.

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6. C. (9) (e) Manual EPR Set/Selection - Selection of the MAN line select keyon the main menu causes the TRS submenu to be displayed (Figure6-18). With DUAL mode selected, selection of either lineselect key adjacent to the MAN EPR values show inverse videoaround the last set EPR value. Both values are set via the DCset knob. This value is settable from either DC on a lastentered basis. Subsequent selections for set will cause thedigital data to be boxed and inverse video. The first entryinto this menu is shown inverse video only. This allows an EPRrating to always be selected.

(f) SPLIT EPRSet Mode - With the SPLIT mode selected, the left andright EPR values can be set independently. Selection of theline key nearest the EPR value, effects only that value in the“select for set’’/’’selectfor display” process.

(10) SYSTEM Mode

The SYSTEM menu, as shown in Figure 6-19, is comprised of thesystem page display selections:

● Hydraulic System Page (HYD). Fuel System Page (FUEL). APU/Bleed System Page (APU/BLEED). Engine Start Page. Engine/APU Exceedance Page (EXCEEDANCES). Waypoint Listing (WAYPT LIST). Checklist

NOTE: For optional TCAS; refer to paragraph 6.C.(15)(b).

(a)

(b)

(c)

SYSTEM Function Key Selection - Selection of the SYSTEM modecauses the function key to be annunciated along with theSYSTEM menu to be displayed on the DC CRT (Figure 6-19). TheSYSTEM mode function is used to select various system pagesfor display. A SYSTEM function key selection has no effect onany of the display formats. SYSTEM page selections are madeusing the menu driven line select keys.

System Page Selections - Alternate activations of any singleline select key toggles between system page displayed andundisplayed. The power-up default is no system page selectedfor display. The system page selections are made from eitherside display controller.

Checklist Function - Selection of the CHECKLIST line key callsup the CHECKLIST submenu as shown in Figure 6-20.

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AD-11792

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MAINTENANCE

Honeywell %N!#h.6. C. (10) (d) Procedure Selection - In this menu the pilot or copilot can

select either normal (NORM), emergency (EMER), or abnormal(ABNORM) checklist procedures. Only one of these selectionscan be active (boxed) at a time.

(e) Cursor Manipulation Functions - The following selections aremomentary and are boxed as long the line select key isdepressed.

● LINE ADV - Moves the CHECKLIST cursor forward through thechecklist by line.

● LINE BACK - Moves the CHECKLIST cursor backwards throughthe checklist by line.

● PAGE ADV - Moves the CHECKLIST cursor forward through thechecklist by page.

● PAGE BACK - Moves the CHECKLIST cursor backwards throughthe checklist by page.

c ENTER - Means by which a particular procedure isselected or checked off complete.

● RECALL - Returns to the first unchecked procedurewithin a given CHECKLIST.

(f) RETURN - Selection of this key returns the main SYSTEM menu tothe DC CRT with the CHECKLIST selection boxed. To deselectthe CHECKLIST system page, another system page must beselected.

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6. C. (11) TEST Mode

The TEST mode function is used to select the following crewinitiated tests (Figure 6-21).

●

●

●

●

●

●

●

●

Electronic Flight Instrument System (EFIS)Engine Instrument and Crew Alerting SYSTEM (EICAS)Radio Altimeter (RADALT)Autopilot Disconnect Logic (A/P DISC)Angle of Attack System (AOA)Autothrottle Disconnect Logic (A/T DISC)Windshear System Test (Option)Maintenance Test

(a) TEST Function Key Selection - Selection of the TEST modecauses the function key to be annunciated along with the TESTmenu to be displayed on the DC CRT (Figure 6-21). The TESTmode function is used to activate crew initiated subsystemtests. A TEST function key selection has no effect on any ofthe display formats. TEST selections are made using the menudriven line select keys.

(b) Crew Initiated System Tests - The individual line keys of thismenu select the self-test mode of the named subsystem. Theactivated test is boxed as long as the line key is depressed.Tests are initiated for the on-side displays only. EICAStests can be selected from either side DC.

(12) DISP Mode

The DISP menu, Figure 6-22, is displayed on the DC CRT in thismode. This menu gives the following display choices.

. Flight Director Command Bar Presentation

- Single Cue (SC) Command Bar- Cross Pointer (CP) Command Bar

. CAS/Mach Display Selection

● Metric Altitude Display Selection

c Baro Set Display Units

- Inches of Mercury (IN)- Millibars (MB)

● Bearing Pointer Declutter Mode (PFD)

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MAINTENANCE

Honeywell WiYAAW

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MAINTENANCE

Honeywell !!R%AA.6. C. (12) (a)

(b)

(c)

(d)

(e)

(f)

DISP Function Key Selection - Selection of the DISP modecauses the function ke.vto be annunciated alonq with the DISPmenu to be displayed oh the DC CRT (Figure 6-22). The DISPmode function is used to configure display parameters of theon-side PFD which are not changed very often. A DISP functionkey selection has no effect on any of the display formats.DISP selections are made using the menu driven line selectkeys.

Flight Director Command Bar Format - By alternate-actionselections, the pilot and copilot can change their respectivePFD command bar symbology between single cue (SC) or crosspointers (CP). One of these selections is boxed at all times.The power-up default is the last selected command bar format.

CAS/Mach Selection - Above 25,000 feet, the speed display canbe selected between the CAS and Mach tapes. The CAS/Machselection is momentary and remains boxed as long as the linekey is depressed. This selection affects only the on-sidePFD.

Metric Altitude Selection - The metric altitude display can beselected on the on-side PFD by using this line select key.The power-up default mode is the last selected state of thisdisplay.

Baro Set Units Selection - The pilot and copilot can selectthe units, IN or MB for the baro altitude set data on theirrespective PFDs. The power-up default is the last selecteddata format.

Bearing Pointer Declutter Selection - Bearing pointers may beremoved from the PFD by selecting the BRG line key on the DISPmenu. Selections made in this mode affect the on-side PFDonly.

(13) CRT Dim Knob

The crew has control over the brightness of the DC CRT via the BRTknob.

(14) Baro Set Knob

Each pilot’s BARO set knob is tied directly to the on-side DADC.When the pilots are displaying cross-side DADC data on their PFD,they do not have control over the displayed baro setting from theirrespective DC. The baro set function is independent from the DCand does not require the DC to work to set the data.

22-14-00Page 196

Aug 15/91Use or disclosure of information on this page is subjecf to the restrictions on the title page of this document.

6. C. (15) Optional DC-884 Display Controller, Part No. 7007540-941/942, forMLS/TCAS Installations

(a) MLS Source Selection

The -941/-942 DC-884 Display Controller is required when a MLSis installed. The DC-884 provides a MLS installed grid/openprogramming discrete at pin J1-d. A ground indicates an MLSis installed. This discrete has priority over the ILSinstalled discrete.

MLS and Bendix ILS will not be simultaneously present in anaircraft due to symbol generator port limitations. Therefore,if the MLS and ILS discretes are both set, the DC shouldignore the ILS installed discrete and consider the system asMLS installed. This will aid installers of aircraft thatformerly possessed a Bendix ILS should they neglect to removethe ILS installed discretes from the DC.

The DC-884 also provides a grid/openMLS select discrete at pinJ1-FF. Control priority over the MLS select discrete is asfollows:

● On-side active NAV● Cross-side active NAV● On-side preview NAV. Cross-side preview NAV

This implies that if the pilot has selected NAV 1, copilotselection of MLS 1 will not toggle the pilot’s discrete.Table 6-4.1 lists the primary cases for this MLS selectdiscrete.

MLS 1 Select MLS 2 Select

Toggled on Toggled on

DC-884 NAV Source MLS 1 Select Preview of MLS 1 MLS 2 Select Preview of MLS 2

Pilot’s DC NAV 2 x

copilot’s Dc MLS 2 x x

Pilot’s DC MLS 2 x

Copilot’s DC NAV 2 x

Pilot’s DC MLS 1 x x x

Copilot’s DC MLS 1 x x x







DC-884 Display Controller MLS Output Discrete LogicTable 6-4.1

22-14-00Page 196.1Apr 15/93

Use or disclosure of information on thispage issubject to the restrictionson the titlepage of thisdocument,

A grid/openremote tune inhibit discrete is output at pinJ1-GG. When toggled, this discrete shall be GND for 500 t 50msec and then return to open state. This discrete toggleswhen the on side MLS is exited. This discrete is used toforce the gables NAV control head to tune the DME in acontinuous label stream rather than in a burst tune mode.This will allow the DME to be retuned when moving from NAVtuning source to MLS tuning source, and then back to NAV, upon

the return to NAV. If this were not done, the DME would notbe returned by the NAV (burst tune only tunes on initialfrequency selection) and an “F” could be displayed on the NAVcontrol head.

NAV Mode Menu

Figure 6-22.1 shows the NAV mode menu when the optional MLSis installed. Alternate activation of the left No. 3 lineselect key moves the box between the 1 and 2 selections ofthe MLS source. On the pilot’s side the first push of thekey boxes the 1 and on the copilot’s side the 2 is boxed.

Preview Mode Submenu

Figure 6-22.2 shows the preview mode submenu with MLS. Theleft No. 4 line select key is used to select the MLS sourceas follows:

. First, push boxes MLS 1 (pilot’s side) or MLS 2(copilot’s side)

. Second, push boxes MLS 2 (pilot’s side) or MLS 1(copilot’s side)

c Third, push removes the box from MLS 1 or 2.

Selection of any other source (NAV or FMS) causes the MLSsource to be unboxed and deselected.

The selection of the #2 NAV source and preview NAV sourcefirst on the copilot side is a product enhancementunassociated with MLS integration. It simply allows thecopilot to select his on-sic NAV/preview source with asingle key stroke rather than double clicking the lineselect key, as is presently required.

The course set window on the NAV preview paqe will be.-removed while MLS is selected. -

22-14-00Page 196.2Apr 15/93


I 1 ~FMS 1 2 1 I

=+N’vE+IMLS

I1

2t

I1

❑ ▼✍I r I

I 1 PREVIEW t 1

AD-34320

NAV Mode Menu with MLS SelectedFigure 6-22.1

1 1

I 1

IPREVIEWIFMS 1

NAV 1MLS 1

22

❑RETURN

, r 1

I r I

t-

AD-34321

Preview Mode Submenu with MLS SelectedFigure 6-22.2

22-14-00Page 196;3Apr 15/93

Use or disclosureof information on thispage issubject to the restrictionson the titlepage of thisdocument.

6. C. (15) (b) TCAS Mode Selection

The DC-884 provides two places to select TCAS pages: thenavigation map mode menu and the system mode menu. Thenavigation map mode menu allows the on-side pilot or copilot toselect TCAS display for the on-side navigation display. Thesystem mode menu allows selection of the TCAS format on thesystem menu. The TCAS system menu is selectable by either thepilot or copilot from their DC-884. Also, the TCAS computercan automatically select the TCAS system page for displaywhenever it determines an intruder to be a traffic advisory(TA) or a resolution advisory (RA). This is accomplishedthrough tying the TCAS TA and RA (preventive and corrective)lamp driver discretes to a system page select discrete on theDC-884.

Navigation Map Mode Menu (See figure 6-22.3)

While TCAS is installed, the navigation map menu provides aTCAS selection on right line select key No. 3 The TCASselection will be ON when boxed and OFF when unboxed.Power-up default for the navigation page menu will be tothe ON (boxed) selection. On the ND map menu, TCASselection will not deselect any other mode.

System Mode Menu (See figure 6-22.4)

While TCAS is installed, the system mode menu provides aTCAS page selection on left line select key No. 5. TCASselections will be ON when boxed and OFF when unboxed.Selection of the TCAS system page by either the pilot orcopilot causes it to be displayed on both DCS. TCAS systempage will be selected on a last entered basis. Power-updefault for the system page menu will be to the OFF(unboxed) selection. When TCAS is selected on the systemmenu, all other selections will be unboxed.

Navigation Display Compass and Plan Modes

TCAS selections will not be available in the ND compassmode. TCAS selections will not be available in the ND planmode.

Automatic TCAS System Page Select

When the TCAS system page select pin (J1-m) transitions toground (falling edge), the TCAS select mode on the systemmode menu will be boxed. The TCAS selection on the systempage shall be capable of being deselected (another modeselected - checklist, waypoint list, etc.) even if thediscrete remains in the grounded state. This discrete willbe debounced 200 msec.

DC Pin Assignments for TCAS

Pin 115/Cl15Jl-j is TCAS installed. Pin 115/Cl15Jl-m isTCAS system page select.

22-14-00Page 196.4Apr 15/93


a BGo FMSI VORI ADFI AUTO 1+-1

BGO FMS2 VOR2 ADF2 I I

=%lmmnl mI D NAVA I D[ [VERT PRo Fi=

~+1 1“1 WIND XY VECT I—EIr !

AD-34322

MAP Mode Menu with TCAS SelectedFigure 6-22.3

I 1

I 1

1I J

HYD ENG START

FUEL EXCEEDANCES

APU/BLEED WAYPT L I ST

tf I

m CHECKLIST r 1 I

AD-34323

System Mode Menu with TCAS SelectedFigure 6-22.4

22-14-00Page 196.5/196.6


USP Bit BIT FUNCTION NOTEFuRMAT SCALE RSBAPPROXRESOL/LSBPOSSENSE FTIUSIHVARII... ...... 11...............-..-.*----...--------------------------I--------------------l---1-----------------l-----------------1----1--------1

HDLCFLAG 7E HEX:1 I... ...... ~-~~~;;~----------””---lIPACKEO~wlc I I l-----------------l---------------"-1~~;~i--------i----------------------------------------------------

1 15 TEST LDGIC1 = TEST 65501 14 DC VALID1 13

LOGIC1 - VALIDTRANSMIT/l!ECElVE LOGIC1 ■ TRANSMIT

: :f~l SPARECOUNTER O-711EX

1 7-; DC ADDRESS 2U H - PILOT 2F N - COPILOT

21I II..-..----......................... -----------------------------I‘-------------------1---1-----------------‘-----”-----------l;;;~i--------1HAVFONMATSELECT

2 1s SPARE 65522 14 mAP/colwMINDS LOGIC1 ● ON2 13 AIRPORTID LOGIC1 = OH212 MIMIXV / vECTOR LOGIC1 ● XV2 11 HAP LOGIC1 = SELECT2 10 cOMPASS LOGIC1 = SELECT29 PLAN LoGIC1 ● SELECT28 PLANMINDS LOGIC1 = ON2 WO;~LUITER21 LOGIC1 ● ON26 UAVPOINTS LOGIC1 = ON25 AIRPORTS LOGIC1 = ON24 UAVPOINTIO LOGIC1 = ON

NAVAIO10 LOGIC1 ● ON:; VERTICALPROFILE LOGIC1 ■ ON21 FPLN SCROLLFORE LOGIC 1 = SCkuLL20 FPLN SCROLLBACK

31

LOGIC 1 ■ SCROLL--- ......I II I----------------------------- -------------------------.................... --- .................

BEARINGSELECTII I-----------------l;;fil--------l

3 IIEARIHGO316 w

6554LOGIC 1 ● SELECT

3 14 ;l#/DC 11 ONI.V) LOOIC 1 ● SELECT313 LOGIC 1 ● SELECT312 ADF 1 LOGIC 1 ● SELECT

“311 FNCS 1 LOGIC 1 ● SELECT::0 SPARE

SPARE38 SPARE3 BEAR:NG(>3? LOGIC 1 ● SELECT36 VOR 2 LOGIC 1 = SELECT35 ADF 234

LOGIC 1 = SELECTWCs 2 LOGlC 1 = SELECT

33 SPARE32 SPARE31 SPARE30 AUTO (DC #2 ONLV) LOGIC 1 = SELECTI l-------------------------l----l-------------------------l... ......

11----------------------- --..--.---.--*---I-----------------l----l--------l

-.

USP BIT 611 FUNCTION

41

NOTE FORHAT SCALE RSB APPROXRESOL/LSB POS SENSE FTIU SIFWARI..- -------------------------------l----l-------------------------l--------------------l---l-----------------l-----------------1;;-;1--------1TESTtUN)E

4 Is EFIS LOGIC1 ● TEST4 14 ElCAS

6556LOGIC1 = TEST

4 13 RADIOAI.T LOGIC1 ● TEST4 12 HINGSIIEAR LOGIC1 = TEST4 II AOAS1 LOGIC1 ■ TEST4 10 AOAAll LOGIC1 ● TEST49 :; ~;): LOGIC1 = TEST48 LOGIC1 ● TEST::-5 SPARE

HAINTTESTENABLE LOGIC1 ● EHABLEO43 UOUGNO LOGIC1 ● UOH42 MAINTTESTSELECT LOGIC1 - SELECT4 l-o SPAflE

51..-------I II......................... ..-*........................-

SY~Sl~,~OEI--------------------l---1-----------------l-----------------I;;i;l--------l

5 15-12 HEXCODE5 00 = HYD

655a

55

01 ● AC PDUER

502. Dc PouER

503. MAYPTLlsT

504 . FuEL

s05 = APU/BLEED06. ENGEXCEEUANCE

5 01. cllEcKL[sT00. s~RY

:5

09. Em sT~TOF . oFF

55 11-10 SPARE

OA - OE - SPARE

CHECKLISTSELECT:9 ENERGEHCY LOGIC 1 ● EMERGENCY58 ALINORHAL LOGIC 1 ● ADHORNAL

NORMAL LOGIC 1 - NORMAL:: LINE ADVANCE LoGIC 1 ● ADVANCE55 PAGEADVANCE LOGIC 1 ● ADVANCE54 RECALL LOGIC 1 = RECALL53 ENTER LOGIC 1 ● ENTER52 IHDEX LOGIC 1 ● SELECT51 LINE IIACK LOGIC 1 = llACK50 PAGEBACK LOGIC 1 ● BACKI l-------------------------l----l-------------------------l--------------------l---l... ......

1----------------------------------1----1--------1

NSP BIT BIT FUNCTION NOTE FORNAT SCALE RSIIAPPIIOXRESOL/LSt)pOs sENSE FTIU SIWARI l;~;;~-;~;;------------l----t-------------------------l-----"--"-----------l-"-l.-. ...... I-----------------.......----------1;;;;1--------166 15-14 INS 00-lRSl 655A6 01 ● IJIS2

10. IMS 3; 11 . #oT usED6 13 W&c,;$z Loolc 1 ● OAOC 16 12 LOGIC 1 ● FW i6 11 IIAUIOALT 1/2 LOOIC 1 ● HAOIOALT i6 10 I)AU1 A/t) LOOIC 1 * DAU IA69 I)AuZ A/ii L061C 1 ● DAU 2A68 SPAUE6? AFLX 1/2 LWIC 1 = AFCS 166 AWOTHtlOITLEl/Z LOOIC 1 ● AUTUTIIROTTLE1665-0 SPARE

,1--- ------1;tiii-ii;-fi;;--"---- l----l--"----------"---------"-l--------------------l---l---------"-------l-----------------t;;;;l---------

? )5 NAv 1 LOGIC 1 ● SELECT 655C1 14 NAv 2 LOOIC 1 D SELECT1 13 FKS 11 lZ

LOGIC 1 ● SELECTWCs 2

J 11LOUlC 1 ● SELECT

HLS 1 L061C i ● SELECT1 10 NLs 2 100IC 1 ● SELECT

11S 1 L(N31CI ● sELECT;: 11S 27?

&OOfC 1 = SELECTLTRK LOOIC 1 ● SELECT

16-0 SPAUE

tli--- ------tfi;;;;~;;;;;:---" --t----t----------------------"--l--------"-----------l---1-----------------1“----------------l;;;;1--------I

8 t5 NAu i l~IC 1 = SELECT 655E8 14 NAv 2 100IC 1 ● SELECT8 13 FNCS 18 }2

100IC 1 ● SELECTFMCS 2 LOOIC 1 = SELECT

8 11 HLs 1 LOOIC 1 ● sELECT8 10 NLs 2M9

LOUlC 1 = SELECTILS 1

88WJIC 1 ● sELECT

11S281

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86 SPAUEas SPARE84 SPAKE

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w- I l-------------------------l----l-------------------------l--------------------l---t-----------------1----------------”1----1--------1--- ------

m

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IvNI

IIISP BIT BIT FUNCTION NOTE,FORMAT SCALE RSB APPROXRESOL/LSB POS SENSE FTJU SJJWAR.--1------I-------------------------I---- ........................-I--------------------l---l-----------------l-----------------l----i--------I9’ ‘THRUSTREFERENCEsrsf40DE’ ‘ “6520” “9 15 AUTOV SPEEIISELECT LOGICI ● SELECT 65609 J4 AUTOEPRSELECT LOGJC 1 ● SELECT FOCTRCL9 13 TO/GAEPR SELECT LOGIC 1- SELECT FDCTGATL9 12 CLB EPR SELECT LOGJC 1 ● SELECT FOCCLBTL9 11 CLB-DEPR SELECT LOGIC 1 ■ SELECT FDCCLBPL9 10 CRSEPR SELECT LOGJC 1 = SELECT FDCCRSTL99 CON EPR SELECT LOGIC 1 ● SELECT FOCCONTL98 EPR DUAL/SPLIT 1. ~AL 0. sPLIT91 FLEXEPRSELECT LOGJC1 ● SELECT9,6-0

..- ------1‘-::Y:---------------”--l----l~~;;;;;;~---------------l~-~-~~~~;~---”-l;~-l;~-‘~~~;;~;~~;~---i-----------------i~~~~t”--”----l:: ;5i4 RSV - REL HEAOINGSELECT

SPARE 6562100- SYNC SELECT LOGIC I = SELECTII... ...... .........................l----i~~fi~;~~”~---”----------‘~l-~~~-fi~---------[;~-l~-&&~~~~~~--l-” ---------------1~~~~1”-------1

11 15-4 REL COURSE#l SELECT .11 3-1 SPARE 656411 0 SYNCSELECT LOBJC 1 = SELECTI l-------------"-----------l----l~~~;~;~~~~---------------l~-J-;~~-~~~---------1~-l;~~;~;~;~~~~~---l-----------------i~;;~i---------

:: :5;4 REL COURSE/2 SELECTSPARE 6E66

120- SYNC SELECTII

LOGIC1 = SELECT--- -....- ............-------------1----1 I I ~ 110,0~7B1 I-----------------l;;;; J--------l-------------------- ..... .................... --- .................:! :5;7 V15:j:~ SET (80 - 250) BJNARY 0-511 KNOTS . .

. 6560130 oISPLAYSELECT LOGIC 1 = SELECT-.. ------.........................l----l-------------------------l.- ~11 ~NoT~ f ~ 110,000781 1-----------------1;;;;1--------1----------------------- ................-14’15-7 ‘V25;~~[oSET (Bo - 250) BINARY . .146-1 656A140 DISPLAYSELECT

1LOGIC 1 ■ SELECT

15’15-7 Vrsj~f;OSET (Bo - 250)II--- ------.............------------.... -------------------------I;---;;;~;fifl-------l-;t;-~fi-~;;;------l-----------------I;;;;l---”1--”l

BINARY -;$ $-1

. .

DISPLAYSELECT LOGIC 1 ■ SELECT656C

--- ..---- -----------”-------------1----lB,my I-------------------------....................I-;I;-J;;;;;;”-----I-----------------l;;;;I--------l116’15-7‘Vrg~A:gEEDSET(BO- 250) 0-511 KNOTS166-1

. .656E

J6 o DISPLAYSELECT LOGIC 1 = SELECTII... ...... .........................l----l~;~”i----------”--------l~---~&~~~~i;~-------------------------l-----”-----------~~~~-l-------”I

;;:5i4 WR;LT SET (O - 2500) . .. 6570

170 OISPLAYSELECT LOGIC 1 = SELECTII--- .---.---;;’~------------------l I---- .............------------1--------------------l---l-----------------l-----------------I;;;I--------l

lB 15-0

I l~~~~ii~~~-~~~~~~~-~--~~j---li~~~~~~~~~~~---------1~;--i:~;-~~~--------1-~-1~-~j~-~~~~~--l~~--------"--......... 6572

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190 DISPLAYSELECT FOCNAOALI l-------------------------l----l-------------------------l--------------------l---l-------"---------l-----------------l----l--------l..- ....-.

i?o.

USP BIT BIT FUNCTION NOTE FORMAT SCALE RSB APPROXRESOL/LSB POS SENSE FTIU SIMVAR..- .-...-f20’15-7 L ~PR#D SET (0s85-2.0 11

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206-1 6576200 DISPLAVSELECT LOGIC 1 = SELECT

II11

FOCLEPRL... ...... I I ~ 100,,0 ~oo781 1-----------------1;;;1;;;;;;;-1---...-*---------------------........................-----------------------------------------21 15-7 R EPR W SET (OoB5-2.0 BINARV 0-5.11 EPR .*21 6-1 SPARE -

OISPLAYSELECT65?8

210 LOGIC 1 ● SELECT-------’”-----------”----1In,wy I

FDCREPRL..- ------ ---- -------------------------;---;ii-~~~;-------l";-li---~;;;;i;---l---"-"--------"--l;;;;l--------l;:’:5i7‘VS:P:;::DSET(80-250) . . .

657A220- oIsPLAYSELECT LOGIC1 ● SELECT/1..------- 11..--..-J----------------------.----*.---.-.---.-Q-----I I ~ I~ ,.,0007812 1-----------------1;;;1--------1

235-7---------------------...................

Vf;psS~~EDSET(80-250) BINANV 0-511 KNOTS236-1

.

OISPLAVSELECT657C

230 LOGIC1 = SELECTI l;;;-;;;~;;;~;~;--------iI..-..-..- -----------------------------I--------------------l---1-----------------1-----------------1;;;;1--------I

2424,15 FO C140BAR 1 ● CROSS-POINTERO ● CUE 657E24 14 CAS/HACHTOGGLE 1. TOOGLE;: :; l);)~NGPOINTERS ~.oN o.oFF

24 11 BAROFORMAT ::;~-llno=nl24 10 W&C UT24 9-o

II--- .----- 1 ITBO ‘------------------------- ---- -------------------------I--------------------l---l-----------------l------------”----lfi#----”--I25 15-0 NAINT TESTUORO

II 116580

..- ---------------------------------* -*---------------------------------------------- -----------------I II, I-----------<-----1;;;;1--------I26 15-0 SPARE

II...--.*-. II6582

----------------------------- -*-----------------------...................- --- ................-1 II I----------------+;;I--------J27 15-0 SPARE

II II6584

..- ..---- ......................... ---- ------------------------------------------------ -----------------1 II I28 15-0 CNECKSUN SUM OF MORDS

165461 I...............- .... ........

II I ICRC6586

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I215 TCAS

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--- l -------------------------------- ~------------------------------ ~.----- .-. -.- . . . . . ..-j.. - l ----------------- ~-------------------------------

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-.. [.-. --- {------------------------- :------------------------------ [---- .--. -.- . . . . . . . . . . . .. I ----------------- :... --.. - . . . ..-. -A l ------------- I

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DC-884 Display Control lerBlock Diagram

Figure 6-23 (Sheet 1) 22-14-00Page 198.5/198.6

Aug 15/91Use ordisclosure of Information on Ihm page K subject 10 the restrictions on the Itle page of this document

MAINTENANCE

Honeywell !Mti+th.

r ———— ———— ———— ——a—— ———— ———— ———— ——,—— —,1

J2 IIIIIIIIIIIIIIIIIIIII

PHOTO SENSOR 1{

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STATUS

OCUS G2FGC LEFT PRIORITY

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ARINC ILS 00

MLS 00

MLS 01

WINDSHEAR INSTALLED

PILOTICOPILOT

IRS TRIPLEXIDUAL

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DC-884 Display ControllerBlock Diagram


Aug 15/91on this page IS sublecl 10 the reslrtctlons on the title page of Ihm document.Use or disclosure 01 mlormahon

MAINTENANCE

Honeywell M%AM.


22-14-00Page 198.9Aug 15/91

Use or disclosure Of information on this page is subject to the restrictions on the title page of this document.

6. D. DA-884 Data Acquisition Unit (See Figures 6-24 and 6-25, and Tables 6-6and 6-7.)

AD-20325

DA-884 Data Acquisition UnitFigure 6-24


Length ...................................... 15.13 in. (384.3 MM)Width ........................................ 4.91 in. (124.7 mm)Height ....................................... 7.62 in. (193.5 mm)

Weight (approximate) .............................. 10.5 lb (21.3 kg)

Power Requirements .................................... 28V dc, 15W

Mating Connector J1 .............................. TR2P106P1O6PT-OOO1


DA-884 Data Acquisition UnitLeading Particulars

Table 6-6 22-14-00Page 198.10Aug 15/91

Use or disclosure of information on this page is subject to the restrititons on the title page of this document.

HoneywellThe DA-884 Data Acquisition Unitbased analog/discrete to digital


(DAU) is a multipurpose microprocessor-conversion unite The DAUS receive all

engine parameter signals, both analog and discrete, and miscellaneoussignals from aircraft sensors. Each DAU is paired with an engine andreceives signals from only that engine. In order to preclude the totalloss of engine sensor data, each DAU is configured in a dual manner. Theduality inherent in each unit includes the following:

. Dual/independent analog input buffers

. Dual/independentA/D converters

. Dual/independent data processing

c Dual/independent output ports

- ASCBA- ASCB B

. Dual/independent power supplies

With the exception of thermocouple and thermistor type sensors, the DAUis independent of other users of the engine or aft data which itreceives. The failure or removal of a DAU or the failure of a singlechannel in the DAU will not, with the above noted exceptions, affect anyother sensor users. This requirement is dictated by the dispatchcapabilities planned for the Gulfstream IV which require a stand-alonestandby engine instrument display.

Parameters transmitted out of the DAU will be via ASCB A or ASCB B, aslisted in Table 6-7. The data on these buses from a single DAU isidentical, since the A/D conversion and processing of the data are doneindependently but in the same manner.

The utilization of the avionics standard communications bus allows thedata to be received directly by several independent system LRUS. Theseinclude but are not confined to the following:

. Fault Warning Computers (FWC)

● Symbol Generators (SG)

● Navigation Computers (NZ)

● Performance Computers (PZ)

. Automatic Flight Guidance Computers (AFGC)

22-14-00Page 198.11Aug 15/91


l-kmco

\.w“ -Wfxi

USP 811 811 FUNCTION NOTE FORMAT SCALE RSIIAPPROXRESO1./LSBPOS SENSE

01

FTIU SIMVAR--- ------I-------------------------l----l~;;~;;-------------------l--------------------l---l-----------------l-----------------1----1--------1

IIOLCFLAGI l-------------------------l----l-------------------------l--------------------l---l-----------------l-----------------l--... ------1“ “DAUCONTROL ‘PACKEDL(MIC1 15 SELF TEST LOGIC 1 ■ TEST IN1 14 vALIO LOGIC 1 = VALI”i i3-n itATAIIEAOER1

00 ● ALTN DATAOol . NoT usEU

i 10 m NoT(JsE~1 11 m ALTNoATA11 11 SPARE1 10-8 COUNTER O-?HEX: v-o DAUAOORESS ?OH=lA 71H=

72 H*2A 7311=

. .

PROGRESS“65UO” -6600665066A0

II...------ 11......................... -----------------------------I--------------------l---l-----------------i-----------------1;--;1--------12 15-11 SPARE2 10-0 TURBINEGASTEPP BINARY(INT) O - 204?OEGC 11 1.0 LIEGC 6602ROATGIL

6652

II. . . . . . . . . II I66A2

1131062344,4, l-----------------l;;;; l;~-’;.-~---. . . . . . . . . . . . . . . . . . . . . . ---- -------------------- ----- -------------------- --- . . . . . . . . . . . . . . . . .3 15-3 RADIoALTITUDE TUO’S COMPLEMENT +/- 2553 FEET32-0 SPARE 0:077930FT 6604

665466A4

II-------------....-..--...-.---*---1---1 I------------------------- --------------------415-12 SPARE

l---l-----------------l-----------------i~;;~l--------l

4 11-0 APUEGTDAU11 BINAJIY(INT)4 11-0 SPARE OAU#2

O - 1019OEGC 12 0.24i380DEGC 6606ROAPUEGI6656

II -------------------------l----i-"--------------"-"------l66A6

..-.--...515-12 SPARE

II----------------------------------------1-----------------1;;;;1--------15 ]1-0 APURPUDAU#l BINARY(INT)5 11-0 SPARE OAU#2

O - 122.3% RPM 12 0.02913656X RPM 6608RDAPURPI6658

II..-......-------------------------1I-----------------------------6 15-10 SPARE

I--------------------l---l----”------------l-----------------1:;1--------1

69-0 STABPOS DAU#1 - STAB BINARV(lNT) O - 10.23DEG 10 0.01OEG STABUP6 15-9 SPARE

660ARDASTBP1

6 U-O665A

FLAPPOS OAU/2 - FLAPII... ...... ----------"-"-""---------l----l~!:fl:-~:::!-------------l~-:-::::-::~--------l-:-l:::-:::------.---l::!!-:::!---.----l:;7 15-0 A/OCALIURATION

660C665C

I IspME---...... .........................B 15

l----l-------------------------l--------------------l---1-----------------l.---..-....-...--l~l--.---.-l8 14-0 FUEL FLOU BINARY ( lNT) O - 10485 LB/llR 15 0.320 LB/llR 660E ROAFFL

665E

II.-. . . . . . . ------------------------- l----l ------------------------- l--------------------l--- l-----------------[ ----------------- 1::!:1--------1

oD

USP BIT BIT FUNCTION NOTE FORMAT SCALE RSB APPROX RESOL/1.S17 POS SENSE FTIU SIMVAR

I l-------------------------l----l-------------------------l--------------------l---l-----------------l-----------------l;..-------9 15-14 SPARE9 13-0 L(N4 PRESSURE TACH (Ml) BINARY (lNT) O - 163.83 % N1 14 0.01 % N1 6610 ROANIL

6660

I l-------------------------l----l-------------------------l--------------------l---l-----------------l-----------------I;6600

... ......10 15-14 SPARE10 13-0 HIGNPRESSURETACH (N2) BINARY(INT) O - 163.83Z N2 14 0.0] X N2 6612 RDAN2L

6662

II6682

... ------.........................1----1;;;;;;”------------------l;---;:~;;;”-l-----l---l---------------”-l-----------------1;;;;1;;;;;;;-11115-4 ~~;[E PRESSRATIO (EPR) .11 3-2

]2 0.0009760EPII

11 1 EPR/OAOCSOLNCE LOGIC 1 ● ON-SIDE66146664 FOAEDSL

110 EPR FLAG LOGIC 1 ● VALIDI l;;;;;--------------------l----l-------------------------l--------------------l---l-----------------l-

66U4 FDAEPRVL--- ------12 15-0

66166666

---l------l-------------------------l----l-------------------------l--------------------l---l-----------------l-----------------l----l--------[66U6

13 15-0 SPARE 65C066186668

II...---------------”----”----------l----l-------------------------l--------------------l---l-----------------l-----------------I:Al--------I14 15-0 SPARE

I661A666A

15 1315 1315 12

I l;;;&;;;;----------------l----l;fi-..------

:: 15 FUELFILTERFAIL LOG15 !4 COMBNYD SYS FAIL LOG15 14 FLT HYO SYS FAIL LOG

COMB HVO SYS NOT LDGFLT HYD SYS NoT LOGIGNITION2 LOG

15 11 CA8IN PRESSLOU LOG15 11 MAIN CABINODORS LOG15 10 RADIOALJ FAIL LOG159 COUL ANTI-ICEON LOG:: ; MINGA/l LOG

IGNITION LOG156 CALL155 OIL PRESSURELOU154 ENGINt153 PYLONNIJT:; :

I66UA

.. . . . . . . . . . . . . . . . . . ..- . . . . . . . . . . . . . . . . . . . .IEDLOGIC

l---l-----------------I-----------------1;;;;1--------Icl.C18 41- DAU

661CFOAFFFL

cl= !2 - DAU666CFOACNSFI

cl. 11 - DAU66BC FOAFIISF2

cl= 12 - lIAUFDACNSlll

cl-WAFNSH2FDAIGN2L

cl. #l - DAU FOACPL1cl= J2 - DAU FDACMCD2C 1 = VALID O = FAIL FDARAAFLcl.cl-

FDACAIOL

cl. FDAUAILFDAIGNILFDACALLL~nAOPLL.JAENIILFDAFLLLFllA~DI I

LOGIC i =LOGIC 1 =

HoT LOGIC i =,“cn

FUEL LOU LEVELFuEL PRESSURELOU

LOGiC i ●

LOGIC 1 =LOGlC 1 ●

,“..,,L,-

150FOANVUL

REVUHLOCK LOGIC1 ■

II----------------------------------l----l-------------------------l--------------------l---l-----------------l-

USP UIT Ill1 FUNCTION NOTE FORMAT SCALE NSIIAPPROXRESOL/LSD POS SENSE FTIU SIHVAR

II... ...... .........................l----l~&;;-i-:------------"---l--------------------l---1-----------------1-----------------l;;;;l;;;;~~l16 15 MING 10116 14 EHEROA1l OISCH LOGIC I =

STALLBARRIERFAIL661E FDAEIJDL

16 13 LOGIC 1 = VALID O = FAIL 666E FI)ASBFL16 Ii? DC PGUERFAIL LOGIC 1 - 66UE FDADCPFL16 11 AC POUERFAIL LOGIC 1 -

COOLING TIM INE NOTFDAACPFL

16 10 LOGIC I ● FIN)AClllL169 APU ALTEIWATUNHOT LOGIC 1 = FoAPUAN1169 UIHB SNEAHFAIL LOGIC 1 = FDAHNSF2168 COW ELEV HYO OFF LOGIC 1 = FDACENO1168 FLT ELEV NYD OFF LOGIC I = FDAFEII0216 7 CONVERTORFAN FAIL LOGIC 1 ● FUACFFL166 SMOKEOETECT LOGIC 1 = FUAWKD1166 FLAMEOETECT LOGIC 1 =

CONVERTORHOTFOAFLMD2

165 LOGIC 1 ■ FOACONNL164 ALTERNATOIIIIOT LOGIC 1 ● FDMLTNL163 Svo LOGIC 1 ■

BLEEOPRESSURENIGHFDASVOL

16 2 LOGIC 1 ■ FDAIIPNI.16 1 BLEEDNOT LOGIC 1 =

AILERONHYD SNUTOFFFDAULUNL

160 LOGIC 1 = FDMNSLI l;fi;~y;;;;~---------l----lLw,c ~ . ~F 0. ~H l,, - ~Au..- ------ I 1-----------------1........................-.................... ... -“---l;~ol-------l;~ol--------l

17 1517 IS17 14i7 1417 1317 1317 1217 1217 11II II17 10:; :

17917817 B11 717 7176176175:; :

17417 3;; ;

170

SPARE-STAB-FLAPGROUNDSPOILERFAIL51N6LERuDoiNLIHITRUODERLINITCOW RUDDERNYD OFFFLT WOOER HVD OFF::: g::l WI rcll

ISOLATIONVALVEENGINESYNCAUX IIVDNOTTRU IK)TLANDINGGEAR STATUSSPEEOURAKEEXTAFT E(jU[PNOIFUD IIADRACK NOTAPU FIREFIRE UELLMUTEAUX AC PDNERFAILAC EXT POUERCABINOXY UNCABINPRESSUREPANSTALLBARRRESERVEOALT FUELFAIL

Ii - DAU12- DAUt] - DAU42- UAU

#l - OAU42- DAU

.OGICI =

.OGICI =

.OGIC1 ●

.OGlc1 ●

.OGic 1 ●

.OGIC1 ●

.OGlc 1 ●

.OGlc 1 =

.OGIC1 ●

.OGIC1 =

.OGlc 1 =

.OGlc 1 =

.OGIC1 =

.OGlc 1 =

.Uolc 1 =LOGIC1 =M3GIC1 =L@lC 1 ●

LOGIC1 ●

LOGIC1 ■

I.OGIC1 =I.OGIC1 =I.OGIC1 =

6620#1 - DAU12- DAU

667066C0

11- DAU42- DAU#l - DAU#2 - OAU11 - DAU12- DAU11 - DAU;: - DAU

- DAU12- DAUt] - DAU82- DAU#l- DAU12- 7)AU11 - OAU42- DAU#l - DAU#2 - DAU#l - OAU12- DAU

LOGIC1 ●

FDAISOV1

FOALGS1FDASOE2

a 1g F-$ gfrJ a , ,~lN FUELFAIL LOGIC 1 =~ti

-1 I l-------------------------l--------------------l---l-----------------l-----------------l----l--------l... ...... ......................... ----

:$1o~:o

1

18 14~: ~:

18 121811181118 1018 10189188181186186Ill518518418418310318218218 118 1180180

IJSPBIT ElT FUNCTION NOTE fORfflT SCALE RSJIAPPROXRESOL/LSB POS SENSE

II

FllUSlttVAR... ...... ~fi~~T;~-"----------l--"-l~~~-;-~~&--;-~fifi-[;;-c-&;----------"-i---l--"-------"--"---l-----------------1~&l--------1,181518 15 ANTI-SKIDOFF LOGICI ● NORM O = FAIL #2 - OAU 6622FOMS02

ALIERNA1ORERGFAIL LOGIC1 ● 6672PITOTHEATFAIL LOGIC1 ●

66C2CAtllNOFRN(9.8) LOGIC1 ● #l - DAUCABINDFllN(9.6) LoGIC1 ■ J2 - DAUENGFIRELOOPALENT LoGIC1 = 11- DAUEM FAULTLOOP ALERT LoGIC 1 = 12- DAUFULL X-FLOU LoGIC 1 ■ #l - OAUFUEL INTERTANK LOGIC 1 =lIAUIDENT#l

12- I)AULOGIC 1 “

ENGINE FIRE -COULANTI-ICE OVERHEATSPARECLlUESLOPt OISCRETESPAREPULL-UPtU/N2 SELECTGROUNDPROXUTILITY HYD OFFSPAREEP14PPS FAILSERVICE DOORStRU FAILBAGGAGEDOORSSTALL BARROFF

p;: :.-

#1 - DAU /1 - IIAULoGIC 1 ● )2 - I)AU#l - DAU /1 - UAULOGIC 1 ■ 12- OAULOGIC1=N2 O=NI 41- DAULoGIC 1 = )2 - DAULoGIC 1 = /1 - lIAU

LOGIC 1 - 11 - OAULOGIC 1 = 82- DAULOGIC 1 = /1 - l)AULoGIC 1 = 12- I)AUL061C 1 ■ 11- DAU

TONEliENFAIL LOGIC 1 ■

II82- OAU

-.. ...... 11.....-*..-............... ---- ...............----------1 II*---------------------- .................I-----------------1;;;;1--------I1915-8 SPARE19 ? NUTCRACKERSNITCH LOGIC 1 =196

6624 FOANUTSLOATT CHGR FAIL 100IC 1 ■ VALIO O = FAIL 6674

195 OAU IDENT #2 LoGIC1 = 66C4194 W FAN FAIL LOGIC 1 ■

193 NAINTEHANCETESTENABLE LOGIC 1 =192 ST8Y PITOTHEATERFAIL LOGIC 1 ●

192 TAT PROIIEMT FAIL#l - DAU

LOGIC 1 ■ J2 - DAU FDATPNF219 1 APU MASTERUARN LOGIC I = ii - DAU19 1 APU All IIRG FAIL LOGIC 1 ● FAIL O = NORM )2 - DAU;; ; BRAKEFAIL LoGIC 1 = NoRM 0 = FAIL #l - DAU

ANTI-$KIO FAIL LOGIC 1 ■ NORM O = FAIL 42- DAUII--- . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1----1 ------------------------- l--------------------1 ---l -----------------1-----------------l;;;~l--------1

20 15-0 SPARE6626667666C6

I l-------------------------l----l-------------------------l--------------------l---l-----------------l-----..-.-.21 15-0 SPARE

66286678

II II66cll

-.. ...... ......................... ---- .........................I--------------------l---1 I I 1--------I................- ................. ....

+ww

m

m

<

UOROSEQUENCEDEPENDSUPON TNE VALUEOF 61 TS 13-12 IN USP 1

ALTERNAIE OATA O

USP BIT BIT FUNCTION NOTE FORMAT SCALE RSB APPROXRESOL/LSB POS SENSE FTIU Slt4VAR

II I l;~:;-~&~~~~"--""-"-l; ;:-;~~;-~;;-~--""-l;; -l~;;;-;;~~------"l;; ~--------------l----l~; fi~~;-[. . . . ----- . . . . . . . . . . . . . . . . . . . . . . . . . . . . .22 15-3 ENG OIL TEW . .

222-1 SPARE220 FLAG LOGIC 1 ● VALID FOAEOTVL

II II..- .-----..--------------------------- -------------------------I--------------------l---l-----------------l-----------------l----l--------I23 15-12 SPARE2311-0 TURD VIO INO - 1P BIRANY(INT) 0-5.096 IPS 12 OoOO12444IPS ROATVILLII... ------ 11 I-------------------------.... -------------------------....................l---l-----------------l-----------------l----1--------1

24 15-12 SPARE24 11-0 TURB VIO IND - HP BINARY(INT) 0-5.096 IPS 12 0.0012444IPS ROATVIHLII II..- ------..--..-..-.--..0........- ---- ......................... ............----------- .................I II I-----------------l----l--------I

25 15-12 SPARE25 1I-O FuELQUANTITV BINARY(INT) 0- 15306.5LBSII

12 3.737052LBS... ...... II I......................... .... ------------------------------------------------.................II I-----------------l----126 15-12 SPARE

::~::?:-1/

26 11-0 CO~ HYDPRESSUREDAU#l BINARY(lNT) 0 -4130.1Psl 12 1.00B573PSI RDACNP126 11-0 FLTNVllPRESSUREDAU42 RDAFNP2II---...... ...............----------II-0---------------------------1--------------------l---l-----------------l-----------------1----1--------1

27 15-12 SPARE27 11-0 UTILHVOPRESSUREDAU /1 BINARY( lNT) 0 -4130.1 PSI 12 1,00B573 PSI IROAUNPI21 11-0 AUX HVO PRESSURE DAU 12 RDAAIIP2II--- ...... II......................... .... -------------------------I-----”-”------------1I I--------------------.................1----1--------1

2815-12 SPARE281I-O EM BLEEDAIRPRESS BIMNY (lNT) O - 122.3 PSI 12 0.0298656PSI RDAEDAPL

II. . . . . . . . . 11. . . . . . . . . . . . . . . . . . . . . . . . . ---- . . . . . . . . . . . . . . . . . . . . . . ..- I --------------------l---l----------------- l-----------------l----l--------12915-12 SPAAE29 11-0 UNEELBRAKEPRESS BINARV(INT)

II0 -4130.1 Psl 12 1.008573Psl RDAHBPL

..- ------ 11......................... ---- -------------------------I;--;;;-;;;-;-”----1,610025,32* I--- ................- .................1“---l&fifl-l3015-0 COW NVD ~ANTITY OAU #l SINANY(lNT) -*3015-0 FLT NVD QJANTITV OAU 42

●

II IIRUAFIIQ2

. . . . . . . . . . . . . . . . . . . . . . . ----------- ---- ------------------------- I ;;”;;fi-~;------l;;-lo,2,0*o,25 tpo5 ~p531 15-4 BATTERYANPS

----------------- . . . . . . . . . . . . . . . . .TNO’S CDNPLEIKNT

I----l;;;;;;;il.

313-0 SPAREI l-------------------------l----l..- ------ I-------------------------....................l---l-----------------l-----------------1----1--------1

32 ]5-12 SPARE32 11-0 BAIlERYVOI.TS

IIBINARY(INT)

II0-37.367 VOLTSUC ]2 0.009125VDC RDABATVL

..- .....- ......................... ---- .----Q...-.....--.-.--..0.................... --- .................I II I-----------------l----l--------13315-12 SPARE33 11-0 AC VOITS

i l------”--------------”---lIB’WY““’)O - 167.527VAC 12 0.0409102VAC RDAACVL

. . . ------ 1---- . . . . . . . . . . . . . . . . . . . . ----- . . . . . . . . . . . . . . . . . . . . l---l-----------------l-----------------l----l--------134 15-12 SPARE34 ii-O iSS AC VOI.TSOAU #l BINARY(MT) O - 16?.527VAC 12 0.0409102VAC ROAEACV34 11-0 AUX AC VOLTSDAU 12I l;;;;--------------------l----l-------------------------l--------------------l---l-----------------l--

RDAAACV2.-. ......35I l;fi;-"------------------l----l-------------------------l--"-----------------l---l-----------------l---..------

36II. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . II 1 II“-.. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . I-----------------l----l--------1

ALTERNATEDATA 1

2u

(D

n

t-lo

s

WI BIT IllT FUNCTION NOTE FORMATI l-------------------------1----1i~;~-~~~i~--------------.. . . . . . .

22 15-3 ENG FUEL TMP22 2-1 SPARE22,0 FLAG LOGIC 1 ● VALID

--- ------ I i l-------------------------. . . . . . . . . . . . . . . . . . . . . . . . . ----

SCALE RS9 APPROXRESOL/LSB POS SENSE FTIU SIHVAR--------------------l---l-----------------l-----------------l--+/- 409.5 OEG C 13 0.1/0.0125 NOT

--------------------l---l-----------------1-----------------23‘15-3 ‘FUELTANK TENP

,“TUO’S COMPLEMENT “+/.409.5OEG C ‘13 ‘0.1/0.0125 ‘NOT

232-1 SPANE230 FLAG LOGIC 1 = VAL10I l-------------------------l----l-------------------------l--------------------l---l-----------------l----..----

24 15-12 SPARE24 11-0 ENG OIL PRESSURE BINARV(lNT) O - 76.59PSl 12 0.0187032PSII l-------------------------l----l-------------------------l--------------------l---l-----------------l-------------------- ------

25 15-12 SPARE25 11-0 ESS OC VOLTSDAU 41 BINARY(lNT) 0-51.047 VDC 12 0.0124657VDC

ROAEFTL

I---- --------ROAFTTL

II------------

ROAEOPLII.-.. . . . . . . . .

RDAEOCV125 11-0 AUXDC VOLTSDAU12

. .

---l------l-------------------------l----i-------------------------l--------------------l---l-----------------l-----------------1----1--------1RDAADCV2

26 15-12 SPARE26 11-0 OC VOLTS B[NARY(lNT) 0- 51.047VocII

12 0.0124657VDC... ...... .........................27 15-12 SPARE

l----l-------------------------i--------------------l---1-----------------l-----------------1----1:~~~:~:-l27 11-0 DC LOAD BINARY(lNT) O - 163.066S 12 0.0398208%II

ROADCLL... ...... .........................28 15-12 SPARE

l----l-------------------------l--------------------l---l-----------------l-----------------[----l--------[28 11-0 AUX DC LOAD

II.0. Q----- ------------------------- i----l~::fl:-!:~~~-------------l:-:-::::::-~-------l::-1=:=----1-----------------[----[!:!!:::!{29 15-12 SPARE29 11-0 AC LOADII

BINARY(lNT) O - 203.47BII 12 0.0496894%1----1

RDAACLL-.. ...... ......................... I......................... ....................30 15-12 SPARE

l---l-----------------l----”------------l----l--------130 11-0 AUXAC LOAOII-........ -------------------------l----l~!f!:~~:~--------------1~-:-:~::~::---------l::-1:::::::~:-:------1-----------------+l!?!!!!!:l

31 15-9 SPARE318-0 AC FREQUENCYII----....- -------------”-----------1-”--l~!~fl:-!!~:~-------------lf-:-:::-::----------l-:-1::~-!:----.-..--.l.---.-----..----.l.Jwq

32 15-9 SPARE32 8-u ESS AC FREQUENCYOAU #l BINANY(INT) 0- 511 Nz 9 1.0 Hz RDAEACF1328-0 AUX AC FREQUENCYDAU /2I l-------------------------l----i-------------------------l--------------------l---l----.---.---..---[.--.--.----------l----[!:!!!!!!--- ------

33 15-12 SPARE33 ii-o- DAU iioxTEHPI I;;;;--------------------I----IB’NMY‘1”1)

0-4095 OEG 12 1,0 OEG... ...... 1......................... ...........---------34 15-0

l---l-”---------------lv -----------------l----l--------1

Glw II---------"------------------------l----l-------------------------l--------------------l---l-----------------l-35 15-0 SPARE

0 N ---------1I -------------------------1----l-------------------------1--------------------l---l-----------------l-----------------[----1--------1%1 36 15-0 SPARE

.!WII--- ------

37;;;;;-;;--------------- 1---- 1;;;-;;-;;;;--” ---------- l--------------------l---1-----------------1-----------------1....1..-...--1

II.........3B

~;~fi-~~;;~---------”----1----I&---------------------l--------------------l---[-----------------[-----.-----------[--.-[--------I

II*.. ......39

~~i~-~~~~---------------1----l~~-”~----”-”------------1--------------------1---1-----------------1-----------------1‘---1--------1

II--- ------ -------------------------1----l------------------------- 1--------------------1 --- l----------------- {----.----.-------l ----1 -------- I

ANALOG ANDDISCRETEINPUTS

T_IIIIIIIII

IIII

-L

POWERI I *

SUPPLY I + I

+

I

I IBIJFFER

I I I

I I II BUFFER C%lv+

I I II I I

BUFFER

I I

—

—

+

—

+_

PROCESSC)R

Y

ASCB ASCB B

I I CHANNEL B1- .—. . L .—— ——— —— —. ——— .—

I II*

BUFFER I*

III

I

I + AID

IBUFFER I - CONV PROCESSOR

I I1 I

* +Ic BUFFER 1 *

I II I I

POWERSUPPLY 1- IE

CHANNEL AI 1

DA-884 Data Acquisition Unit Block DiagramFigure 6-25 22-14-00

Page 198.19/198.20Aug 15/91

INPUT lWIE QUANTITYSERIAL DIGITAL (429) 1DC SYNCHRO 1VARIABLE AC VOLTS 7VARIABLE DC VOLTS 12TACH 2VARIABLE RESISTANCE 3VARIABLE DC AMPS 3SPECIAL PHASE REF MODULATED 1THERMOCOUPLE 128 V DISCR=E 56GND DISCR~ 13TOTAL ANMOG 31

TOTAL DISCRETE 69

DAU INTERFACE SUMMARY

>ASCB A

AD-8974-RI

Use or disclosure of InformalIon on this page IS subjact to the restrictions on the tllle page of thm document.

This page intentionallyleft blank.

22-14-00Page 198.21Aug 15/91


Honeywell6. E. DP-884 Dimmer Panel (See Figures

MAINTENANCEMANUALGULFSTRE4MM

6-26 and 6-27, and Table 6-8.) “

DISPLAY BRIGHTNESS

Do

PFD<~ ND El<?= CAS ND+>PFD

DP-884 Dimmer PanelFigure 6-26


Length ..................................... 4.44 in. (112.73 mm)Width ...................................... 5.75 in. (146.05 mm)Height ...................................... 1.13 in. (28.58 mm)


Power Requirements:

Lighting ...................................... 5 Vdc, 2.5Wmax

Mating Connector J1 .................................. MS27473E1222S

Mounting ....................................... Unit Dzus Fasteners

DP-884 Dimmer PanelLeading Particulars

Table 6-8

The DP-884 Dimmer Panel provides the brightness control for the sixdisplay units. The controls consist of three dual concentric clutchedpotentiometers. One dual concentric potentiometer each controls thepilot’s PFD/ND pair, the EICAS DU pair, and the copilot’s PFD/ND pair.The panel contains no electronics. The actual brightness controlfunction takes place within the DUS. With the DP-884 Dimmer Panelremoved, the displays revert to a median value. The displays also go toa median value with potentiometers open or shorted.

22-14-00Page 198.22Aug 15/91

Use Or disclosure of information on this page is sutqect to the restrictions on the title page of this document

JI

PILOT PF5 H I <, ---- ‘lLoT ‘FD BRT

BRT CNTL W 2RIAIOK

DUI L 3 <i

PILOT ND H 4 ---+ ‘lLoT ‘DbRT

BRT CNTL W 5RIBIOK

Du2 L 6

IElWIT H 7 <!

t+ .3 / --* “ ‘R’

CNTL DU 3 W 8 <,w .2/ 4 R 2A

I,Cwl IOK

19L

‘1

CAS BR’ H IC <iH 3/ --- CA’ “T

CNTLDU4W II <;w -2 / 4 R2B

<!L

,Cwl 10 K

L 12I

ICOPILOT ND H 15 <,

H -3 / ---- cOpiLOT ‘o BRT

BRT CN’L W 1~ <’w .2/4 R3B

,Cwl IC)K

Du5 L 15 <iL

I

COPILOT PFCI + 16 <1H .3 / ---- cOplLOT ‘FD BRT

tSRT CNTL W 17 <:w .2/ R 3A

<;L

,Cwl 10K

Du6 L IE

I

o-5 VDC Is < !

DSiGND 2C <[

CHASSIS GND“ +“

SPARE 22 <-h

DP-884 Dimmer Panel SchematicFigure 6-27

22-14-00Page 198.23Aug 15/91

Use or disclosure of Information cm this page IS subjecf to the restrictions on the title page of this document.

6. F. FC-880 Fault Warning Computer (See Figures 6-28 and 6-29, and Tables 6-9and 6-10.)

AO-302S4

FC-880 Fault Warning ComputerFigure 6-28

22-14-00Page 198.24


MAINTENANCE

Honeywell MM%!h.


Length ...................................... 16.45 in. (417.8 mm)Width ......... ............. .. ......... .... .. . 4.91 in. (124.7 mm)Height ....................................... 7.62 in. (193.5 mm)


Power Requirements .................................... 28V dc, 45 W

Mating Connector:

J1 ................................... DPX2MA-A106P-A1O6P-33B-OOO3


FC-880 Fault Warning ComputerLeading Particulars

Table 6-9

The FC-880 Fault Warning Computer (FWC) is a digital computer primarilyresponsible for supplying data to the symbol generators for display ofwarnings, advisories, and aircraft system status information on the EICASdisplays. The FWC receives data directly from aircraft systems and overthe ASCB from the data acquisition units, symbol generators, displaycontroller, digital air data computers, priority flight guidancecomputer, and from the FMS. Specific discretes for cockpit annunciatorsor horns are outputs of the FWC.

Information displayed on the crew alerting system (CAS) display portionof the EICAS display is received, logically organized, placed throughsystem algorithms, and transmitted from the FWC to the symbol generatorsover the ASCB.

Both channels of each DA-884 Data Acquisition Unit are monitored by bothFWCS with errors between the two channels being annunciated on the CASdisplay. The DC-884 Display Controller contains the selectioncapabilities for the pilot to select the appropriate FWC to be utilizedby the display system.

The FWC contains a user-accessible checklist module. This module has thecapability to store 64k bytes of text data for the checklist andemergency checklist system pages. The actual contents of the module isuser-defined.

The FWC monitors various subsystem status and discretes to determine whenthe display and FGC can achieve a Category 11 approach. The logic iscontinuously monitored. Status is reflected on the PFD and EICAS asdetermined by the FWC.

22-14-00Page 198.25Aug 15/91

Use or disclosure of Information on this page is subject to the restrictions on the title page of this document.

MAINTENANCE

Honeywell M!!!%h.The FWC performs a trend and limit exceedance recording function forengine and APU data. This recording function is performed automaticallythroughout the flight and also upon request by a discrete input to theFMC. The FWC interfaces with the DL-800 or DL-900 Data Loader fordownloading trend and limit monitor data from FWC memory. The interfaceis over an RS-232 bus.

The FWC performs the ground maintenance test function. This includesaccepting test requests from the DC-884 Display Controller, verifyingthat the LRU selected for test is available and ready, and outputting thespecific LRU test commands. Test results are recorded (volatile memoryonly) and displayed.

Critical flight data displayed on the PFDs and engine display aremonitored and compared with the original source data. Additionally, theFWC assures the integrity of the other wrap-around buses by monitoringfor activity. The monitoring task is performed continuously through theflight as a means of assuring data integrity.

The FWC initiates the BC test discrete with every cold start power-up onthe ground. The logic required to generate the test request is includedin Section 4. The FWC digital outputs transmitted on the ASCB are listedin Table 6-10.

22-14-00Page 198.26Aug 15/91

Use or disclosure of information on this page is subject to the restrictions on the title pege of this document.

MSP BIT BIT FUNCTION NOTEFORMAT SCALE RSB APPROXRESOL/LSBPOSSENSE FTIU SIIIVAR

NN

.

I lfi[;;[fi--"-------------l----l;;-;~;-------------------l--------------------l---l-----------------l--------"--------l--... ......

--;l------;;;@-”;-@--”-----------lIPACKEOLWIC-----------------------------I--------------------l---l-----------------l-----------------l;;-;1--------1

1 15 TEST LOGIC1 = TEST 69201 14 VALID LOGIC1 ■ VALIO1 13 CHECKLlST/SYSTEFIS 1. CHECKL1sT0. sYsTEHDATA: ;g-~1 SPARE

COUNTER o-7 HEX1 1-; FUCADORESS 23 H - LEFT 27 H - RIGHTII... ...... .........................1----1-------------------------‘---”----”----------‘-- ‘-----~~~~-&i”~/;~j””--”---------l~~~~l--------l

2 NESSAGETEXT LINES 1,7,13.19 /HESSAGETEXT SENT IN FRAN S : LINES 1,2,3,4.5,II

2 CONTROLUORO 7,8,9,10,11,12 FRA14E2/3 69222 15-13 SPARE 13,14,15,16,17,18 FRAME4/52 19,20,21,22,23,24 FRAME6/72 12-8 MESSAGELINE NUHBER 00000 ● SPARE22

00001- 11000- LINES 1-2411001 -

27-511111 ● SPARE

SPARE; 4-2 LINECOLOR 000- BLACK 100= GREEN

001 = RED 101 = BLUE2 010 ● AHBER 110 ● MAGENTA2 011 ■ VELLOU 111 = UHITE21 FLASHENABLE I . FLAsH O . CoNTl~ous20 OISPLAYLINEENABLE 1 . V1sABLEO . BLANK

---l------l-------------------------l----l;;;;;--------------------l--------------------l---l-----------------l-----------------l=--;l-------.l3 15-8 CtlARACTEll237-0 CIIARACTEN1 ASCI1II.-.------.........................l----l~~~~i--------------------1--------------------1---1‘----------------1-----------------1~~1--------14 15-8 CHARACTER44 7-o CIIARACTER3 ASC11II-........ .........................I----IASC** I-------------------------....................5 15-8 CHARACTER6

l---l-----------------l----------------”I;*l--------l5 7-o CHARACTER5 ASCIII l-------------------------l----liiiii--------------------l--------------------l---l-------.---------l.---......

6 15-8 CHARACTER86 7-o CHAHACIER7

IIASCI1

--- ------ -------------------------l IASCI1. ..- . . . . . . . . . . . . . . . . . . -------7 15-8 CHARACTER10

I -------------------”l--- l-----------------l--”--”-----------l:l--------[7 1-0 CHARACTER9II

ASCII.-. ------.........................B 15-8 CHARACTER12

l----lfi~l~-------------------l‘-------------------1---1-----------------1-----------------1~~~1--------1

8.7-0 CHARACTER11 ASCII 692EII-.. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

9 15-8 CHARACTER14l----l~;~l~--------------------l‘-------------------1---1'----------------l-----------------l~~~~l--------I

I 9 7-o CHARACTER13 ASC11

II...------‘------------------------1---- lfi~ii--------------------l‘-------------------[---1‘----------------10 15-8 CHARACTER16107-0 CHARACTER15II

ASCII--- -------------------------------II----.........................I------------”-------l---l-----------------

6930----------------4 ;;;;1 --------1

6932“----------------l----l -------- I

-mmI

NN

I

hi-5

USPBIT BIT FUNCTION NOTE FORNAT SCALE RSB APPROXRESOL/LSUPOS SENSE FTIU SIHVAR... ...... ‘------------------------1‘---l~~ii-----------------”--l‘-------------------i---1-----------------1-----------------‘~~~~i--------l~~‘;5~8 ‘CHARACTER18

CNAIIACTER17 ASCII1- I II

6934... ------........................- ---- -------------------------I--------------------1I I

12168461 I--- . . . . . . . . . . . . . . ..- . . . . . . . . . . . . . . . . . ---- --------

MESSAGE TEXT LINE 2,8,14,206Q36

II p---------......................... -----------------------------I II----

-------------------- --- ----------------- I 16B5AI f................-------------22 MESSAGETEXT L [NE 3,9,15, 1

694A

I l~~~~~~-~~~~-[i~~-~-i~i~!~~--(-------------------------[--------------------t-"-1... ...... I-----

..................................1;-;1--------132 *

II

695E

I l-----*-------------------l..- ------ ------------------------- ---- II I-------------------- --- ................- .................42

l&l--------lNESSAGETEXTLINE5,11,17,23

6972II II.-.......-----------------------------------------------------------------------------.................I II 1-----------------1;;;[-------- I

52 MESSAGETEXTLINE6,12,18,240 6906ii NESSAGELINE 24 MILLUSE

UORO FOR TIIECAUTIONANO:;52 16-13 SPARE52 12-8 HE4E:~GELINE NUNGER52 1-652 s AMGERFIELOFLASHENABLE

THE FOLLOMIHGCONTROL----

HARNINGSTATUSLINE

11000 ■ 24

1 ■ FLAsH 0- cowTIMous52 i ~AIB~ f IELOOISPLAYENABLE 1 ● VISABLE O ■ BLANK523-252 1 BLUEFIELOFLASHENABLE520 BLUEFIELDDISPLAYENABLE

1 ● FLASH O = CONTINUOUSI ● vlslOLE 0. BLANK

I l-------------------------l----l----"--"-----------------l---*----------------l---l-----------------l--..-------62 15-0 HESSAGELIST CHECKSUM sunOFuomsFM ALL AcTIvERED tiEs3A6Es

II II699A

--- ...-0--..---.-.-.-..--.-0---------- -------------------------I II I----------------------- -----------------.................63 15-0 MAINTENANCETEST

l;;~l--”-----l

II II699C..- ..---- ------------------------- I-------------------------------------------------

64 15-0 MAINTENANCETESTl---l-----------------i-----------------1;;;;1--------I

II II689E

--- ----------------------------------- ---------------------------------------------65 lNPUTLMSCRETES

I l---l-----------------l-----------------1;;;;1--------165 15 L FUEL SHUTOFFOPEN LOGIC 1 ■ OPEN6S 14 R FUEL SHUTOFFOPEN LOGIC 1 ■ OPEN

6BA0

65 13 L FUEL SHUTOFFCLOSED LOGIC 1 = CLOSED65 12 R FUEL SNUTOFFCLOSED LOOIC 1 = CLOSED65 11 CONG HVD SHUTOFFOPEN LOGIC 1 = OPEN65 10 FLT HYO SHUTOFFOPEN LOGIC 1 = OPEN659 CONIIHYO SNOTOFFCLOSEO LOGIC 1 = CLOSED658 FLT HYO SHUTOFFCLOSEO LOGIC 1 = CLOSED657 GNO SPOILERUNARH LOGIC 1 ● SPOILERARtlEO656-1 SPARE650 INHIBITSELECT LOGIC 1 = INHIBITFUNCTIONENABLEII---..*---........................-l----l-------------------------l--------------------l---1-----------------/-----------------/----/-------j

tlsPBIT BIT FUNCTION NOTEFORNAT SCALE RSO APPROXRESOL/LSBPOS SENSE FTIU SI14VARI l;~;-;;~;;~~;;-~;;~-;;~;~----l-"-----------------------l--------------------l---l-----------------l----..-------

6666 15 TESTCOMPLETE LOGIC 1 = CONPLETE

TEST IN PROGRESS6ilA2

66 14 LOGIC 1 = IN PROGRESSTESTRESULTS

:: 13 BC1 STATUS LOGIC 1 ■ TEST PASSEO66 12 BC2 STATUS LOGIC 1 ■ TEST PASSEO66 11 UC3STATUS LOGIC 1 = TESTPASSEIJ66 10-5 SPARE664-3 CAT 1 & II LOGIC 00 = SPARE

ol . cAT it10 = cAT [11 - SPARE

662-0 SPAREII..- ...... .........................1----1 I II......................... .................... ... ................. .................

67 15-0 CNECKLISTCNECKSUMI

HEX1;;;;1--------160A4

..- . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1----1 ------------------------- l--------------------l---l-------”---------l-----------------1----1--------168’15-0 ‘SPARE(60-73)--- -----*ls?STtflS-PRGt-PRflh~tftR5-:lOlSP~Avrfl-nfltR-gSP-x-fllf-I3-1SlStT-lO-O-"----------l---lI I----------------- ----------------- 1----1--------1

II..-..----.........................i----l-------------------------i II...................-....................I-----------------l----l--------I74 SPARE(74-117)II--- -.--------;~;;;;i;-fi;;;;;;-:lfi;;~AyEo“,lEN~ ~,, 13 I II----..................... .................... --- ............-----1-----------------l----l--------l

IS SETTO 1I l;;;;;;i;;;;;[i;-i----l----l----------------"--------~-------------------l---l;;;~-;;;-;;;;;-"l-----------------l;;~;l---..--.-...

74 CHECKLISTTEXTSENTlNFR t4ES:LINES1.2.374

;:

;:747474

;:

;:

H7474

;:7474

;:7474

15-13

8-5

4-2

ii 114 a

CONTROLUOROSPARE

4:5;6 FRANE2/3 “ “7,0,9

60L!2FRAME4/5

10,11.12 FRANE6/7CHECKLISTL[NENUMBER 0000 = SPARE

0001 - 1100 ● LINES 1-121101- 1111 = SPARE

CHECKLISTCONTROL o-

;:

::

::

;:9-A-B-c-

LINECOLOR :0;001010011

FLASH ENABLE 1.DISPLAY LINE ENABLE 1=

. . . .SPARE-. . .CNECKL1STSPAREHYORAULICPAGEAC POWERPAGEOC POWERPAGEAPU/BLEEDPAGEA/c CONFIGPAGEFUEL SYSTEMPAGEENG EXCEEOANCEPAGESUW4ARYPAGEENG STARTPAGE!#W#NCE TEST

= BLACK 100 = GREEN= RED 101= BLUE- AMBER 110 ● MAGENTA= YELLOU 111 - UNITEFLASN O = CONTINUUUSVISABLE O ■ BLANK..-

1 l-------------------------l----l-------------------------l--------------------l---l-----------------l-----.--.-

HSP BIT BIT FUNCTION NOTEFORMAT SCALE RSllAPPROX RESOL/LSB POS SENSE FTIU SIklVAR

iD-5

u

NN

Io0

II I IASC,,-...-----......................... ....------------------------------------------------.................7515-8 CHARACTER2

I II I-----------------1;;;;1--------1

75 ?-o CHARACTER1 ASC11I l-------------------------1----l~;;ii--------------------1

68B4---...... II I...................----................. .................7615-8 CHARACTER4

1;;;;1--------1

767-0 CHARACTER3 ASCI1II I lA~c,*

6806... ...... ......................... ----------------------------------------------------.................77 15-8 CHARACTER6

I II I-----------------l;#-----”-I

77 ?-o CHARACTER6 ASCI1I l-------------------------lIASC**.-....... I II

6808--------------------------------------------------................... .................1

78 15-8 CHARACTER81--;;1--------1

707-0 CHARACTER7 ASC11I I-------------------------I----IA5C**

68UA---...... I Il.-------------------------...........---------... ................. .................,7915-8 CHARACTER10

I 1;;;;1--------1

791-0 CHARACTER9 AscltI l-------------------------llAsc*i..-..--.. I-------------------------------..*---------------

80 15-8 CHARACTER12l---l-----------------l-----------------1:;1--------1

807-0 CHARACTER11 ASC1[I I----------”--------------l----lAsC,,

68BE-..------ 1------------------------- --------------------81 15-8 CHARACTER14

l---l-----------------l -----------------l;;;f ------- I817-o CHARACTER13 ASC1[II I IASC**

68C0..-.-----......................... ------------------------------.--....--.-.----.0-i i---l-----------------182 15-8 -----------------1;;;;1--------ICHARACTER1682 7-o CHARACTER15II

AscllI Imcl,

68C2---------------------------------------------------------------I II----------------------------------------183 15-8 CHARACTER18 -----------------l&#------i

83 ?-O CHARACTER17 ASCI1II II

68C4... ...... ......................... ----......................... ...................-184 15-8 CHARACTER20 ASC11 l---l-----------------l-----------------1;;;;1--------184 ?-0 CHARACTER19II

ASCI1I IA$C*,

68C6... ...... ................--------- 1 I 1-----------------1...-------------------------------------------------85 16-8 CHARACTER22 -----------------l;;&------ I85 7-O CHANACTER21II... ---------”-----”------”----”---1----l:;:--------------------l-"----------------"-l---l------------.----l..-----.--.-.-.-.l:::l--------I

86 15-8 CHARACTER2486 ?-o CHARACTER23 ASCI[I l-------------------------lIASC,*

68CA... .----- 10----------------------------............--------87 15-8 CHARACTER26 l---i-----------------l-----------------1;;;;1--------I81 7-O CHARACTER25II

ASC11II

68CC..-....-.......................... ----------------------------------------------------188 15-0 CHECKLISTTEXTLINE2,5,8,11 I 1-----------------1-----------------1;;;;1--------I

I l------------------------lI.-.—

---...... I II68CE

---- ---------------------------------------------..- .................I102 15-0 CHECKLISTTEXTLINE 3,6,9,12 -----------------l;;;~l--------I

II.-.......;;;~~-;~~;;;-;~----1-”--1 I II68EA------------------------- . . . . . . . . . . . . . . . . . ..- --- . . . . . . . . . . . . -----

116 I -----------------1 ;;;;1-------- I

116 8EGINNING ATTRIBUTES116 15-11 LINENUIBER BINARY 0-15

68F2

11610-8 COLOR5

STANOAROCOLORCOOE116?-3 ~lll~CTERNU!SIER1162-1

BINARY O-63 5

116,0 80XENABLE 1. ENABLE... ......1 II......................... .... [......................... ....................[---1-----------------I-----------------l----l--------L

-1PIo-

(D

mI

w0

-1-101

(+m-5

zo

USP BIT BIT FUNCTION NOTE FORHATII

SCALE RSII APPROXRESOL/LSU POS SENSE..- ------

3;iiii[i;i-;iiiii-ii;---"-l----l-------------------------l--------------------l--"l------------"----l--~

117117 ENOING ATTRIBUTES117 15-11 LINE NUMBER

68F4BINARY 0-15 5

117 10-8 COLOR STANDARD COLORCODE 2117 1-3 CHARACTERNW4BER BINARY O-63 51172-0 SPAREI l---------"---------------l----l;;;-;-~~"------"-----l--------------------l---l--"--------------l-----..-------

118 CNECKSU14

II..-..-..- 1 ICRC........................------------------------------119 ERRORCHECK

II...------.........................l----l~;;~”;-------”-”--------120 NOLC FLAG

II... ...... --"---"------------------l----l------"---------"--*----=

68F6--------------------l---1-----------------l-----------------I;--;l--------l

m~~

:g~68F8 3C5

--------------------l---l-----------------l-----------------l----l--------lg~:--------------------l---l-----------------l-----------------l----l--------l~ =

$

(42s BIT A5DISCRETE All

I(13)+

(SAMEASA12)~

, BUFFER EUFFER* )

( CLOCK- GENERATOR

EEPROM2hx 16811

b

POWERSUPPLY

MONITOR

16BIT BIDIRECTIONALOATA GLOBAL BUS

g

WxMONITOR

DISCRETERECEIVERS

TIMINGMONITOR

DISCRETEA12

+ GNDIOPEN ADDRESSRECEIVER OECODER

*+ +28VIOPEN ~ OATA

RECEIVER MUX*

+ PROGRAMMABLE_RECEIVER

(W>

(W)-

—

r~ 1

+2W DRIVER DATA~TCH

--F==P16 BIT BIDIRECTIONALDATA LOCAL BUS

I

IlltRINC Al I ASCBA2

748— BUFFER

RAM (MAILBOX)6A X 16 BIT BuF=

=

4 EACH

DUALCHANNEL

42s

SERIAL10 FBuS

BUFFER J%DIRECT

MEMORYACCESS

SERIAL !/0CONT

(LOCAL)

(9)—

11 a 12

RECIEVERS

1=1

RBuFFER

SERIAL VO ~ MANCHESTER ~

CONT ASCB CODEJU ASCB 1

~ DECODER ~

MUX At40BUFFER

ARINC +DATAOUT +

QASCB2AAt+

RS-232 AD.14389 1 -R2

FC-880 Fault Warning ComputerBlock Diagram


Apr 15/93

Use or disclosureof informationon thispage issubjectto the restrictionson the titlepage of thisdocument.

L16 BIT BIDIRECTIONALDATA GLOBAL 8US

MAINTENANCE

Honeywell WJ!!L.

I I16 BIT DATA

BIDIRECTIONALLOCAL BUS

( BUFFER BuFFER

CONTROL

-1

CPU AlO

1

I 1- CLEAR

I 1-LOOI( AHEAD

Ik I

J

1 J

MEMORY AS

I POWERSUPPLYCONTROL CIRCUITS I I BUFFER

AODRESS

A6DECODER

r I

i

POWERSUPPLY+28Vcc TRANSFORMER CIRCUITS

A?

FC-880 Fault Warning ComputerBlock Diagram

Figure 6-29 (Sheet 2)

PCONTROL

1

CHECKLISTMODULE AS

AD.143w @ .R1

22-14-00Page 198.35Aug 15/91

Use or disclosure of information rm this page E sub)ecf to the restrictions on the title page of this document.

MAINTENANCE

Honeywell !!f&!%b.6. G. MD-880 Checklist Module (See Figure 6-30.)

1

D..——-~~mm4.n-

-. —.*.. . . . .— * . ..

1 J

—

MD-880 Checklist ModuleFigure 6-30

The G-IV FWC checklist program allows customer-defined pages of text tobe displayed on the system display page. There are six major types ofpages: normal index, normal checklist, abnormal index, abnormalchecklist, emergency index, and emergency checklist. An index page is alist of the normal, abnormal, or emergency procedures that a pilot maywish to review during flight operations. A checklist page is a detailedlisting of the normal, abnormal, or emergency items within a particularprocedure that a pilot may wish to check during flight operations.

The pilot has 10 pushbuttons which he uses to view the pages and indicatethat he has completed and/or checked the procedures or items. Thesebuttons are: normal, abnormal, emergency, enter, recall, page advance,page reverse, line advance, line reverse, and return. The information is

transmitted by the display controller over ASCB (WSP 5 Bits O-9). Pilotand copilot commands are accepted on a last entered basis.

22-14-00Page 198.36Aug 15/91


MAINTENANCE

Honeywell !i’ki$%kThere are six major modes of operation in the checklist program and twominor modes. Each of the modes is directly related to the type of pagebeing displayed. The modes are:

!hiQr

. Normal index mode● Normal checklist mode. Abnormal index mode. Abnormal checklist mode. Emergency index mode. Emergency checklist mode

Minor

. Disclaimer listing mode● Autocallup mode (This is a subset of emergency checklist mode)

The checklist module installs through the top of each FC-880 FaultWarning Computer in the A3 card slot.

Gulfstream Aerospace controls the software for the checklist module andattaches a GAC part number decal on the unit for reordering.

22-14-00Page 198.37Aug 15/91


~Mf;f~ANCE

Honeywell GULFSTREAMIV

7. DFZ-820 Dual Fliqht Guidance System

A. FZ-820 Flight Guidance Computer (See Figures 7-1 and 7-2, and Tables 7-1and 7-2.)

AO-20325

FZ-820 Flight Guidance ComputerFigure 7-1

22-14-00Page 198.38Aug 15/91

Use or disclosure of information on this page IS subject to the restrictions on the title page of this document

Honeywell !#!#ii.cE


Length .................................... 15.13 in. (384.3 mm)Width ...................................... 4.91 in. (124.7 mm)Height ..................................... 7.62 in. (193.5 mm)

Weight (approximate) ............................ 13.1 lb (6.04 kg)

Power Requirements .................................. 28Vdc, 40W

Mating Connector:

J1 ...................... Cannon Part No. DPX2-67S-106P-33B-OOO4J2 ...................... Cannon Part No. DPX2-67S-106P-33B-OO17

Mounting ..................... Tray, Honeywel1 Part No. 7003272-902

FZ-820 Flight Guidance ComputerLeading Particulars

Table 7-1

The FZ-820 Flight Guidance Computer (FGC) processes information aboutaircraft actual attitude versus a desired attitude as a function ofselected flight mode to produce autopilot pitch, roll, and yaw contro”outputs and flight director pitch and roll steering command outputs.addition to the modes selectable on the GP-820 Flight GuidanceController, the computer will produce pitch and roll control outputs ~any flight director mode except go-around.

the

In

or

The FGC has a dual processor architecture, each Drocessor Derforminadifferent control and computational functions. the A-processor per~ormsthe outer-loop flight control computations while the B-processor performsthe inner-loop flight control computations, as well as the autopilot andyaw damper servo loop closures. Each processor monitors the otherprocessor’s flight control functions as well as its own functions. Eachprocessor has dedicated program memory and scratch pad memory. The datagathered by the A-processor through the ASCB is simultaneously stored inthe bus data transfer RAM for access by the B-processor.

The status transfer RAM is used to exchange status information betweenthe A and B processors. This transfer occurs once per 25 ms real timecycle. In addition, a set of discretes is provided for immediateinteraction between the processors.

22-14-00Page 198.39Aug 15/91

Use or disclosure of information cm this page is subject to the restrictions on the title page of this document,

MAINTENANCE

Honeywell !!RN!#&.The A-processor has access to two presettable timers which it uses to setthe time base for the bus transactions (clock 1), and for the inner andouter computational cycle in both processors (clock 2). The B-processor,additionally, has a separate real time clock to time the servo loopclosure computational cycle (clock 3).

All the 1/0 is memory mapped. Each processor individually controls itsown analog and discrete input/output transfers with the exception of theserialized discretes. Discretes fall into two categories: direct andserialized. The latching of the serialized discrete inputs is under thecontrol of the A-processor only. Once the inputs are latched, however,each processor has independent access to them. The serialized discreteoutputs (to the control panel) are solely under the control of theA-processor.

The heartbeat monitor and power supply monitor interlocks ensuredisengagement of the FGC in case of a processor failure, a softwarefailure, a power supply failure or a power outage. The servo driveengage interlocks ensure that the flight control functions can beactivated only if all the monitors are satisfied. The flight controlsare output through the trim, A/P and Y/D servo drives.

The flight director interface outputs the validity annunciations computedby the A-processor.

The FGC digital outputs transmitted on the ASCB are listed in Table 7-2.

22-14-00Page 198.40Aug 15/91

Use or disclosure of information on this page is subject 10 the restrictions on the title page of thm document.

USP fIIT BIT FUNCTION NOTE FORMAT SCALE RSB APPROXRESOL/LSB POS SENSE FTIU SIMVAR

01..- ------ 1;fi;-;[;;-----------”----lI?E,,Ex I I l-----------------l-----------------l----l--------l----........................-.................... ---

,1 I II... ...... ......................... .... ......................... .......................I I l-----------------l---------"--"----l;;wl-----AFCS CUNTIN)L

1 15PACKEDLOGIC

TEST LOGIC 1 ● TEST 50801 14 VAL10 LOGIC 1 = VALID FASAPVLLI 13 MAIM START LOGIC 1 ● START~ ;; BANKLIMITHI/LO LOGIC 1 ■ HIGN

NOUE/ENMOE PRIORITY LOGIC 1 ● HI(NI FAPNEPRLI 10-8 COUlfTElt 0- ? NEXi Iio AFCS Al)iiiN3S iillii- LEFT 31 II- RIINIT

J-........1“”- 1 1-- ““.........................----.........................I II I 15”021-------”--------------------... ..................................----AFCS FUNCTIONSSTATUS

2 15-14 AUTOPILOTSIATUS2 13-12

OG s oFF/lNvALJo ON OFF (IF usp: #11.=ol~);Nl~AIWER STATUS ol , oFF/sTA~Y

51N12lAPSIATL

2 11-1029-8 PACN TRIM STATUS 11 . sELEcTEo/EffiAGEo2? FLltiNTOIRECTUNSTATUS ] s vAL~o o = FAIL FAPFDSTL26 EITHERCNANNELENGAGED LOGIC 1 = EITIIERCIIANNELENGAOEO25 FGC SENSONVALIOATILNICOMP LOGIC 1 ■ COMPLETE

COLD STAIIT LOGIC 1 ● START:: 01SENGAGEANNUNCLEAR CLEAN ● TRANSITIONFROM LOGIC 1 TO LNIC O22 AUNONNALDISCONNECT L061C 1 ● ABNORMALDISCONNECT21 RIGNTSEL/LEFTSEL20

1 . RloHT 0. LEFTTCS ACTIVE

FAFCSELL

3’

LOGIC 1 ● ACTIVEI;~~~;;;;;;-;;~;;-----l----l

FAPTcSAL--- ------ I......................... .................*.-l---1"----------------l-----------------t;;;;f--------

3 15-14 AllRS/lRSOAIA SELECTEO 00 ■ NEITHER3 13-12 OAOCOATA SELECTEO 0~ ● RIGHT 51N34

333 11-10

:339-83 1-63335-4

:

iu. LEFT11. AVERAGE

RAOIOAT SELECTEO 00 = OEFAULT;: : ~;~~T

11 . AvERAGElLS/NLSLAT SELECTEO 00. NEITIIERlLS/MLSVENT SELECTEO 01. RIGNT

PRIORNSI SEL STATUS11 . AvERAfjE

00 ● NOT USEO

N 3 ii ● ji(jT usEO33-1

N

NAV SOURCESELECTEO WO ● NONE001 = ILS

I: 010 ● NLs

n-- 3 011 ■ VOR100 ● RNAV101 ● VI.Fi10 ● INS111 = GPS

SPARE

8 97- 3$$%*a :e -1 3

q%= 30II. . . ------ --. ---... -.-0 ------------

ah=l----l-------------------------l--------------------[--.[------.-----.----l..---.?--

w=

USPUIT OITFUNCTION NOTEFORMAT SCALE RSB APPNOXRESOL/1.SOPOS SENSEII

FTIU SIMWIR... ...... .........................l----l;~~~~;ffiffi-------------1--------------------1---1-----------------1-----------------l;~~;l--------l

M(N)EACTIVEACTIVEIOENT: 15-11 PITCHNOOE [DENT INTEGER 0-31: yis

5006ROLLNODE IUENT INTEGER O-63 :

. TNNUSTMOUE IIIENT lNWGER

5’

0-31--- ------I~’-?&;~;-;:---l;-------l----l;;~;;~fiffi-------”----1--------------------1-:-1-----------------l-”---------------l;;;;l--------5 15-8 RSV-DAOCViiAVLOGIC 50885 J-4 SPAUE53-2 ROLLSUO-MOOE INTEGER o-3 25 1-0 PITCNSUO-MODE INTEGER

61---------I;’;;~-fi-;;--------l----Ifi;fiffi-----”--”-- 1:-:-:-”---”i---------i-:-l-----------------l-----”----””-----l---;l--------

: ;5;10 SP/U/E 5U8ALOUPRIORITYMOOE INTEGER 0-31

6 4~05

HIGNPRIOR[TVMODE INTEGER

71

0-31... ......I~;;fi;;;;;;-;;;;;;-----l----l-------------------------l--------------------l:-l-----------------l-----------------l;~;;l--------7 15-14 OFFSET IUENT @ . No oFFsET

01 = NEGATIVEWFSET500C

10. poslTlvEoFFsET11 ● SPARE

7 ROLLNODESARUEO IOENTSTRING1 13-6 SPARE1 5-o MOOE ARNEO 10ENT INTEGERII -------------------------l----1

o-53--- ------ 1---------------------------------------------815-0 SPARE

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239 SPARE FAPvAPAL238 BC ARM LOGIC 1 ● ‘ON’ FAPBCAL237 MC ARM LoGIC I = ‘ON’ FAPAPLAL236 AZ AIM LOGIC 1 ● ‘ON’235 AZ CAP LOGIC 1 ■ ‘ON’

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23 1 HOG CAP LOGIC 1 ● ‘ON’ FAPND(iCL230 VOR CAP LOGIC 1 = ‘ON’ FAPVORCLIi... ------ 1,1......................... ---- ......................... ....................

24 15-2 AIR OATACOWANO I l---l-----------------l-----------------1;;;;1--------24 (VALUEOEPENIISON NOOE ENOAGEOUSP 22 VERTMOOE)24 TMO’S COMPLEMENT

5BAE+/- 25600FPM 14 3.125/0.781 CLING

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305 iAOCHiSCONPARE LOGIC1 ■ ‘ON’364 ELEVMISTRIMNOSEUP LOGIC1 ■ ‘ON’ FAPENTUL383 ELEVHISTNINNOSEDN LOGIC 1 ● ‘ON’ FAPEMTDL38 Z TCSENGAGE LOGIC1 = ‘ON’

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38 1 TRIM I380 AP TRIMFAIL LOGIC 1 = ‘ON’

I l;;;~;;-----------------l----l-------------------------l--------------------l---i-----------------i--.-:------

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FZ-820 Flight GuidanceBlock DiagramFigure 7-2

AO.1SS32

Computer

22-14-00Page 198,48Aug 15/91

Use or disclosure of information on this pege is subjecf to the restrictions on the title page of this document.


22-14-00Page 198.49

Aug 15/91Use or disclosure of information cm this page is sub@cf to the restrictions on the title page of this document.

7. B. GP-820 Flight Guidance Controller (See Figures 7-3 and 7-4, and Table7-3.)

GP-820 Flight Guidance ControllerFigure 7-3


Length . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11.58 in. (294.13 mm)Width .................................... 13.78 in. (350.01 mm)Height ..................................... i?.7sin. (10.61 mm)


Power Requirements:

Primary ......................................... 28Vdc, 1.6ALighting ........................................ 28Vdc, 1.OA

Mating Connectors:

J1 ..........................J2 ..........................

Both connectors use strain rel

. . . . . . . . . . . . . . . . . . MS27473E20B35SB

.................. MS27473E20B35SD

ef MS27506-A20-2

Mounting ..................................... Unit Screw Fasteners

GP-820 Flight Guidance ControllerLeading Particulars

Table 7-3

22-14-00Page 198.50Apr 15/93


MAINTENANCE

Honeywell M$’gA&.

7. B.

The GP-820 Flight Guidance Controller is used to engage the autopilot,select the operating modes, select the pilot’s or copilot’s PFD, and armthe autothrottle system. It also contains controls for setting course,heading, indicated airspeed or Mach, vertical speed, and altitude. Thefunction of each switch or control is described in the followingparagraphs.

(1)

(2)

(3)

(4)

(5)

Autopilot (A/P) Engage Button - The A/P pushbutton engages theautopilot and yaw damper functions simultaneously, but disengagesonly the autopilot. The external yaw damper button engages anddisengages the yaw damper only. The A/P pushbutton and the PFDannunciate that the autopilot is engaged.

Primary Flight Display Command (PFD-CMD) Button - This pushbutton is

used to select which PFD is coupled to the flight director. Theflight director shall utilize the navigation source as displayed onthe selected PFD. Selecting the cross-side PFD shall clear allflight director NAV modes (VNAV, LNAV, APR, and BC). It does notdisengage the autopilot.

The system shall normally power-up to left (L) side selected, andsubsequent actuations of PFD-CMD shall alternate between right (R)and left (L) couple. The button annunciates L or R to indicatewhich PFD has been selected.

During a dual approach, both PFD-CMD indications shall illuminate tothat the system is averaging data received from both sides.

The flight guidance computer selections for attitude, heading, andDADC references are accomplished independentlyof PFD-CMD.

Autothrottle Arm (A/T ARM) Button - This pushbutton is used to armand disarm/disengage the autothrottle system. The power-up state isdisarm/disengage. The autothrottle system must be armed before anyautothrottle mode can be engaged.

Lateral Navigation (LNAV) Button - This pushbutton is used to selectand deselect the lateral navigation mode. The source of lateralnavigation shall be the source selected for display on the PFD towhich the flight director is coupled (FMS, VOR/ILS, or MLS).

Vertical Navigation (VNAV) Button - This pushbutton is used toselect and deselect vertical navigation from the FMS. VNAV willautomatically utilize the PZ-800 Performance Computer for speed,profile and fuel calculations. The pilot may override theseperformance values through utilization of the CD-81O CDU or theGP-820 Controller’s speed window.

22-14-00Page 198.51Aug 15/91

Use or disclosure of Information on this page is subject to the restrlcttons on the title page of this document.

MAINTENANCE

Honeywell M!%!htv7. B. (6) F1ight Level Change (FLCH) Button - This pushbutton is used to

select and deselect the flight level change mode. Upon theengagement of the FLCH mode, the SPEED window shall become active,displaying the speed value as commanded by the FMS. Manual speedselection can then be accomplished as described in paragraph7.B. (13).

Actuation of the FLCH mode shall cancel all other vertical flightdirector modes with the exception ofVNAV.

(7) Heading Select (HDG SEL) Button - This pushbutton is used to selectand deselect the heading select mode. The desired heading referenceis selected with the HDG SEL knob. This mode is a direct captureonly mode and will result in exiting any other captured lateralmode.

(8) Bank Limit Select (BANK) Button - This pushbutton allows manualselection of the bank angle limit in the HDG SEL mode. Alternateactivation of the BANK button causes alternate selection of a 28-degree high bank limit or 20-degree low bank limit. At power-up,high BANK is selected.

Automatic toggle to low bank status shall occur when climbingthrough 29,500 feet. Automatic toggle to high bank status shalloccur when descending through 28,500 feet. Manual selection of highbank status shall be allowed above 29,500 feet, and manual selectionof low bank status shall be allowed below 28,500 feet. When in lowbank status the button annunciates 1/2.

(9) Back Course (BC) Button - This pushbutton is used to select anddeselect the back course function and display for ILS back courseapproaches.

(10) Approach (APR) Button - Thisdeselect the approach mode.arm ILS or MLS vertical path

(11) Verticaldeselectthe VERTaircraftcan then

(12) Altitude

Speed (VS) Button -

pushbutton is used to select andThe approach mode must be selected tocaptures.

This pushbutton is used to select andthe vertical speed mode. -Upon engagement of the VS mode,SPEED window shall become active, displaying the currentvertical speed value. The desired vertical speed valuebe selected with the V/S knob.

Hold (ALT HLD) Button - This pushbutton is used to selectand deselect the altitude hold mode. The aircraft shall maintainthe altitude that the aircraft is at upon engagement of the ALTmode. This mode is a direct capture mode only and will result incanceling any other captured vertical mode.

22-14-00Page 198.52Aug 15/91


7. B. (13) IAS/MACH and SPD Button, and SPEED Window and Knob - The speed knoballows the pilot to manually input a speed target to the system.Speed values up to VH~M can be set. This knob is active anytime

I

Jothe SPD button annuncia es manual (MAN). The speed values are usedfor the VNAV or FLCH mode. If the autothrottles are engaged in aspeed hold mode, the speed knob is also used to set a speed valuefor autothrottles operation.

When MAN is not annunciated, the speed value is automaticallysupplied to the system from the FMS. The VNAV or FLCH modes thenuse this speed value. The autothrottles also use the same speedvalue for their functions. In manual speed, the IAS/MACH buttonchanges the SPEED window display from IAS to MACH. The PFDannunciates IAS or MACH and also displays the selected speed targetin the IAS or MACH target areas.

(14) Heading Select (HDG SEL) Knob, SYNC Button, and Heading Window -The HOG SEL knob sets the desired heading and positions the headingbug on the PFD. The selected heading is displayed in the HEADINGwindow and also on the PFD. A push-to-sync button, incorporated inthe knob, causes the bug to synchronize to the current aircraftheading.

(15) Vertical Speed (V/S) Button and VERT SPEED Window and Knob - Thevertical speed knob is used to input values of vertical speed tothe system. The VERT SPEED window and the V/S target area on thePFD display the selected value when the V/S mode button isannunciated ON. While TCS is engaged, the PFD displays dashes; theVERT SPEED window blanks.

(16) Altitude Preselect Knob and ALTITUDE Window - The altitudepreselect knob controls the preselect altitude displayed in theALTITUDE window and on the PFD. The system assumes that thisaltitude is a clearance altitude for all flight regimes exceptfinal approach. For this reason, the ALTITUDE window is alwaysactive.

(17) Course (CRS) Select Knob and SYNC Button - The CRS knob allowsselection of VOR, MLS, or ILS course by the positioning of thecourse select pointer on the PFD. The selected course is alsodisplayed as a digital readout on the PFD. A push-to-sync button,incorporated in the knob, causes the course select pointer to syncto the bearing (track, if applicable) of the selected NAV Source.If the applicable bearing or track information is not displayed onthe ND, the sync feature is disabled.

22-14-00Page 198.53/198.54

Apr 15/93Use or disclosureof informationon thispage issubject to the restrictionsonthe titlepage of thisdocument.

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5VCCNPPDWSSI FITNN02 L :m n% l--= “ 2mw””’ -’ IAD-202E6,4 ‘— }L——————— ——————— ——————— _____A


22-14-00Page 198.57Aug 15/91

Use or disclosure of information cm this page IS subject to the restrictions on the title page of this document.

MAINTENANCE

Honeywell %N!!b~7. c. PC-880 Turn Pitch Controller (See Figures 7-5 and 7-6, and Table 7-4.)

u JJ40 S676

PC-880 Turn Pitch ControllerFigure 7-5


Length ...................................... 4.82 in. (122.42 mm)Width ....................................... 5.75 in. (146.05 mm)Height ...................................... 2.63 in. (66.80 mm)


Power Requirements:

Potentiometer Excitation ........................ t15Vdc, 1.6vALighting .................................. 5 V ac or V dc, 5.0 VA

Mating Connector:

J1 ............................................... MS27473E12B35SA

Mounting ....................................... Unit Screw Fasteners

PC-880 Turn Pitch ControllerLeading Particulars

Table 7-4

22-14-00Page 198.58

Aug 15/91Use ordisclosureOfinformationon thispageissubjecttothe restrictions on the title page of this document,

The controller provides the means of manually controlling the autopilotthrough the TURN knob and PITCH wheel. The following paragraphs describeeach control.

7. C. (1) Pitch Wheel - Rotation of the pitch wheel results in a change ofpitch attitude proportional to the rotation of the wheel and in thedirection of wheel movement. The pitch wheel provides rate limitedpitch commands in pitch hold mode. The pitch thumb wheel providesdual tachometer outputs which are applied to both flight guidancecomputers to ensure fail-passive pitch wheel operation. Pitch wheeloperation is inhibited in GS, FLCH, VS, and VNAV modes.

(2) TURN Knob - Rotation of the TURN knob out of detent results in aroll command. The roll angle is proportional to and in thedirection of the TURN knob rotation. The TURN knob controls dualdetent switches and potentiometers to provide identical rollcommands to both flight guidance computers. The TURN knob must bein detent (center position) before the autopilot can be engaged.Rotation of the TURN knob cancels any lateral mode selected.

22-14-00Page 198.59/198.60


MAINTENANCE

Honeywell W!!&.

JI JI

I————— —_______ ___,

0T05vDC OR VAC(-901 AND-902) 9~––––––-–––1 77 El

I I

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CENTER TAp, SIGNAL GROUNO

+15 EXCITATION

-15 EXCITATION

CENTER TAP, SIGNAL GROLND

+15 EXCITATION

(SYSTEM I)TURN CONTROL DETENT SWITCH

(SY5TEM 2)TWfN CONTROL OETENT SWITCH

(SY3TEM t)

6YS.TEMI)

(SYSTEM0

(SYSTEM 2)

(SYSTEM 2)

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1

CHASSIS GROUND

TURN CONTROL WIPER OUTPUT (SYSTEM 1)

TURN CONTROL WIPER OUTPUT (SYSTEM 2)

i I

I I I) 10(+)

I+

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(SYSTEM I)TURN CONTROL OETENT SWITCH

TURN CONTROL DETENT SWITCH

(SYSTEM 2)TU7N CONTROL DETENT SWITCH

TURN CONTROL OETENT SWITCH

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I ——.. — II II, I >19(-)I I

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+-

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Ii

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II I

I I I ) 22(-)I

I I—————————- -——— -—— —-—— _- —__ --- —-————— -——— ——- ————

PC-880 Turn Pitch Controller SchematicFigure 7-6 22-14-00

(NORMALLY CLOSED IN DETENT)

(CLOSEO ou7 OF DETENT)

(NORMALLY CLOSED IN DETENT)

(CLOSED WT OF DETENT)

PITCH[PITCH

PITCH(pITcH

WHEEL GENERATOR OUTPUT

UP POLARITY SHOWN)

WHEEL GENERATOR OUTPUT

UP FOLARITY SHOWN)

AD-12103-Rl

Page 198.61/198.62Aug 15/91

Use or disclosure of mform~tlon on this page IS subject to the reslnchons on the hlle page of Ihm document.

Honeywell !!#!%r.cE


22-14-00Page 198.63

Aug 15191Use or disclosure of reformation cm this page is subject to the restrictions on the title page of this document.

Honeywell7. D. SM-600 Dual Servo/SB-600 Bracket

Bracket (See Figures 7-7 through

MAINTENANCEMANUALGULFSTRE4MIV

and TM-260 Dual Trim Servo/TB-2617-11, and Table 7-5.)

SM-600 Dual Servo and SB-600 BracketFigure 7-7

TM-260 Dual Trim Servo and TB-261 BracketFigure 7-8

22-14-00Page 198.64

Aug 15/91Use or disclosure of information on this page is subject to the restrictions on the tnle page of this document,

MAINTENANCE

Honeywell HY!#b

Dimensions, including bracket (maximum):

Length (TM-260) ............................. 11.28 in. (286.5 mm)Length (SM-600) ............................. 11.75 in. (298.5 mm)Width ....................................... 8.08 in. (205.2 mm)Height ...................................... 4.17 in. (105.9 mm)

Weight including bracket (maximum) ................ 10.2 lb (4.63 kg)

Power Requirements:

Clutch ........................................... 28 V dc, 1.9 VASynchro .................................... 26 V, 400 Hz, 10.4 VA

Mating Connector:

J1 (TM-260, SM-600) ............................... MS27473E12A35SJ2 (TM-260, SM-600) .............................. MS27473E12A35SAJ1 (TB-261) ....................................... MS27473E1OA35SJ2 (TB-261) ...................................... MS27473E1OA35SA

Mounting ................................................. Hard Mount

SM-600 and TM-260 Dual ServoLeading Particulars

Table 7-5

The dual servo translates electrical inputs into a brake controlledrotational mechanical output. The tachometer output is fed back to theflight guidance computer servo amplifier to null the command signal. TheSM-600 synchro position output is not used in the system.

Each servo of the dual servo is identical and is connected to a separateoutput connector. Servo connector J1 is connected to the pi1ot’s flightguidance computer, and connector J2 is connected to the copilot’s side.The active flight guidance computer provides a brake output to disablethe other servo.

22-14-00Page 198.65/198.66

Aug 15/91Use or disclosure Of information on this page is sub]ecl to the restrictions on the title page of this document.

MAINTENANCE

Honeywell K?Pi!%h~-

JI(II d BRM -r---- -r

L-l I I

ORN I(+VOLTACE) 1s1-12

2CW 12 ~ ~CLUTCH I I

CXCIIATION ,3 ● I ‘0 ICtll

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SM-600 Dual Servo SchematicFigure 7-9 22-14-00

14 W14T/BLK/014N J2

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13 5$5

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AD.458.R2

Page 198.67/198.68Aug 15/91

Use ordisclosureofInfonnallonon this page IS subject 10 the restrictions on the title page of Ihm document.

JI

SHIELD GMD IS -KZOmI

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AD-12753

Page 198.69/198.70Aug 15/91

Use or disclosure of mformallon on this page IS subject to the restrictions on the title page of this document.

MAINTENANCE

Honeywell !!!W#h~

~–––––___–__––.-. ———— ———— —— ———1

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LIMIT SWITCH

LIMIT SWITCH

ES

AD.656

TB-261 Bracket SchematicFigure 7-11

22-14-00Page 198.71

Aug 15/91Use ordisclosureofIntonationon thispageIssubjectto the restrictions on the title page of this document.

Honeywell8. PRIMUS@ 800 Weather Radar System

A. WR-800 Receiver Transmitter (See

MAINTENANCEMANUALGULFSTRE4M IV

Figures 8-1 and 8-2, and Table 8-1.)

The WR-800 Receiver Transmitter (RT) generates and receives X-band radiofrequency energy for the purposes of weather detection and groundmapping. The 9345 *3O MHz transmitted signals are sent to the antennavia a waveguide transmission line. Echo signals received by the antennaare routed through the waveguide to the RT and then to the receiver. Thereceiver processes the signals, encoding them into one of four levels,depending on their intensity, scan converts them, and enters them intothe random access memory. This memory is read in raster scan format, andits output is delivered to the indicator and electronic flightinstruments.

bQ 5320

WR-800 Receiver TransmitterFigure 8-1

22-14-00Page 198.72Aug 15/91

Use or disclosure of information cm this page is subject to the restrictions on the title page of this document.

MAINTENANCE

Honeywell !!K$!%$k.


Length ....................................Width ......................................Height .....................................

14.68 in. (372.87 mm)7.50 in. (190.50 mm)7.63 in. (193.80 mm)

Weight .......................................... 20.5 lbs (9.35 kg)

Prime Power .................................. 27.5 V dc, 6A nominal115 V ac, 0.5A nominal

RF Power, Peak (At R/T F1ange) ....................... 1.2 kW Nominal

Frequency ............................................ 9345 f 30 MHz

Noise Figure ..................................... 9 dB at RT flange

Pulse Width and Pulse Repetition Frequency:

Range WEATHERScale (NM) ~ w

300 180 15

200 180 15

100 440 6

50 600 3.5

25 600 3.5

10 600 3.5

MAPu

15

15

6

2

2

2

STC: Digital Control

WX: 6 dB/Octave at ranges shorter than

Beam-Filling

9 dB/Octave at ranges greater than

Beam-Filling

MAP: 6 dB/Octave

WR-800 Receiver TransmitterLeadi~~b~;r~i:ulars

22-14-00Page 198.73

Aug 15/91Use ordisclosureofinformationon thispageissubjecttotherestrictionson thetitlepageofIh(sdocument.

Transmitter: Positive Anode Magnetron

Local Oscillator: Gunn Diode

TR: TR Limiter (gated)

Stabilization: Microprocessor Controlled

Pitch Input:ARINC 429, Label 324

Roll Input: ARINC 429, Label 325

Tilt: t15 deqrees maximum (controlled from

Elevation Angle:

Signal Processing:

Display Memory:

indicator)

Tilt angle annunciated

*3O degrees maximum

Digital correlation inplanes.

128k (2 - planes)

on indicator

RHo Theta, X,Y

Data Correlation:

Antenna Turn-Around:

256 vertical x 256 horizontal

(236 vertical x 240 horizontal displayed)

32k - (REACT) (3rd plane)

Not inhibited in any mode.

Approximately 1 degree from scan endantenna goes to half and then quarterspeed before turn-around. After turn-around, antenna goes from quarter speedto half speed and reaches full speedapproximately 1 degree after turn-around.

Mating Connectors: Cannon KPSE-06F22-55S and KPSE-06F22-55SW

WR-800 Receiver TransmitterLeading ParticularsTable 8-1 (cent)

22-14-00Page 198.74AUG 15/91


MAINTENANCE

Honeywell w&IJ&iv

----- 9 -------- 99 -------- 9-------------9- - 9 ---- 9 ---9-9--- - 9------99-99--9--,--0-9 --,-m

# REcElvER/TRj$NsMITTER

ANTENNA-ANTENNA DRIVE

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FROM AIRCRAFT ~---m, mwaw=, mm=== ===== ===== m======= == c_ ======_=_ ============ m-----a------- -,===GYRO AD-11326-R 1

WR-800 Receiver TransmitterBlock DiaqramFigure ~-2 22-14-00

Page 198.75/198.76Aug 15/91

Use ordisclosureofInlormellonon this page IS sublect to the reslnctlons on the title page of thm document

Honeywell !!$ii~.c’


22-14-00Page 198.77Aug 15/91

Use or dkclosure Of information on this page is subject to the restrictions on the title page of this document.

MAINTENANCE

Honeywell !!iiP’#kiv8. B. WC-81O Weather Radar Controller (See Figures 8-3 and 8-4, and Table

8-2.)

AD-1 2761

WC-81O Weather Radar ControllerFigure 8-3


Length (from rear of bezel) ................... 7.0 in. (177.8 mm)Width ........................................ 5.75 in. (146.1 mm)Height ........................................ 1.87 in. (47.5 mm)

Weight ............................................. 1.9 lb (0.86 kg)

Power Requirements:

Primary ..................................... 28 V dc, 0.06 AmpereLighting ..................................... 5 V dc, 0.92 Ampere

Mating Connector:

J1 .............................................. MS27473E14A-18SA

Mounting ....................................... Unit Screw Fasteners

WC-81O Weather Radar ControllerLeading Particulars

Table 8-2

22-14-00Page 198.78Aug 15/91

Use or disclosure of information on this page is subjecf to the restrictions on the title page of this document.

Honeywell !&!!!!.=’The WC-81O Weather Radar Controller contains all the controls required tooperate the PRIMUS@ 800 Weather Radar System. The controller has threesnap-action pushbuttons, two rotary switches, and three potentiometers toselect and control the operating modes of the radar system. Thefollowing paragraphs describe each of the PRIMUS@ 800 modes and featurescontrolled by the WC-81O.

8. B. (1) Target (TGT) Alert - The TGT button allows the pilot to enable/disable the target alert mode at the radar system. In the targetalert mode, an alert is flashed on the weather display when a red orheavy precipitation area is detected in a 15-degree arc between 60and 120 nautical miles ahead of the aircraft.

(2)

(3)

Ground Clutter Reduction (GCR) Mode - The scintillation frequency ofground radar returns is lower than that of rainfall radar returns.A digital frequency filter is used to separate ground returns fromrainfall returns, and only the rainfall returns are displayed whenthe GCR mode is selected. Since some of the rainfall returns fallinto the same spectrum as the ground returns, there is some loss ofweather return in the GCR mode. As a result, the weatherpresentation in this mode cannot be considered calibrated. However,the GCRmode gives the pilot a dramatically improved look at weatherin terminal areas or mountainous terrain where it may be necessaryto tilt the antenna toward the ground to see weather ahead. The GCRannunciator above the switch lights to indicate to the pilot that heis in the GCR mode.

Rain Echo Attenuation Compensation Technique (RCT) Mode - Rain echoattenuation compensation techniques automatically increase thereceiver gain in real time as a function of attenuation due topenetrated rainfall to maintain system calibration. This circuit isimplemented digitally in the PRIMUS@ 800 and, in conjunction with adigitally generated sensitivity time control (STC), a signal isgenerated when the receiver gain has reached its maximum availablevalue. At this time, a blue color is displayed for the remainder ofthe displayed range. This gives the pilot an unmistakable warningthat attenuation is hiding possible severe weather areas that cannotbe accurately detected. The RCT annunciator above the switch liqhtsto indicate

(4) GAIN Adjustgain. AUTOthe weatheranalysis or

to the pilot that he is in the RCT mode.

- The GAIN potentiometer selects the radar receiverGAIN calibrates the system for optimum performance inmode. Variable gain is used for additional weatherground mapping.

(5) MODE Control - The rotary MODE control is used to select one of thefollowing positions:

● OFF - This position turns the weather radar system functionallyoff.

22-14-00Page 198.79Aug 15/91


MAINTENANCE

Honeywell MPJIAL.

8. B. (6)

(7)

(8)

(9)

● STBY -

● TEST -

.wx -

● GMAP -

Used to select standby mode. Standby is useful forkeeping radar in ready state while taxiing, loading, etc.In standby, antenna does not scan, transmitter isdisabled, display memory is erased, and tilt remainsactive. STBY is displayed in mode field.

Used to select a special test pattern to allowverification of system operation; 100-mile range isautomatically selected; TEST is displayed in mode field.Transmitter outrwt Dower is radiated in TEST mode.

Used to select weather detection operation. If selectedprior to end of warmup period, WAIT will be displayeduntil RT warms up (approximately50 seconds). WX isdisplayed in mode field. Transmitter outDut Dower isradiated in WX mode.

Used to select ground mapping display; MAP is displayed inmode field. Transmitter”out6ut Dower is radiated”in-MAPmode.

Slave (SLV) Annunciator - In a dual-controller setup, thisannunciator will light on the controller whose MODE control is inthe OFF position when the MODE control on the other controller is inan operating mode (mode other than OFF). If the MODE controls on

both controllers are in the OFF position, the SLV annunciator onboth controllers will extinguish.

RANGE Select - Operating ranges from 10 to 300 nautical milesprovide up to 94,156 square nautical miles of weather detection.Selection of FPLN causes the radar information to be blanked fromthe navigation display. The display will also be set to a rangewhich is pin-programmable at the rear connector (J2) of thecontroller. The radar continues to transmit in the mode selected.

TILT Adjust - The TILT potentiometer adjusts the antenna tiltfrom full down (-15 degrees) to full up (+15 degrees).

Forced Standby (No WC-81O control) - With weight-on-wheels a groundis applied to J1-P which forces the system into the STBY mode asdescribed in paragraph 8.B.(5).

(10) Brightness (BRT) Control - The BRT control is used to control radar(raster) information brightness (intensity) of the navigationdisplay (ND).

22-14-00Page 198.80Aug 15/91


Honeywell !f!!%b.c’

MODEWIPER

x!&{$pj--~;;4-BIT

SWITCHES6.144 MHZ

i

COUNTERS OscA3U27, A3U28. t 1 A2U 4, A2Y1

I

FSBY (N)

*OFF

%MODE

STBY ENCODER /3 BUFFERTEST PROM /

RANGE Wx + A2U12 1 T* A2U16

CONTROL GMAP. . . . _

1025

D

lTL

50 LOGIC

100 CONVERTER

200 A2ul

300 A2U3

4FLIGHT PLANRANGE CODE

P RNG 2“PRNG2’P RNG 2’P RNG 2’

!3i?RANGEENCODER

5 PROMA2U7

‘E4(N’nOE l(N) AID

CONVERTERpf+zm=CONTROL

---

>’57

*

m ToTE:—REFERENCE DESIGNATIONS ARE AS FOLLOWS:I I I Al = POWER SUPPLY CCA

f

A

t

A2 = LOGC 1 CCAA3 = LOGIC 2 CCA

[ CONTROL I I LABEL I

EPROMA3U6

OE1

OE2

OE3

OE4OE5

OE6 TPROMA3U8

(N)

(N)

(N)

(N)

(N)

(N)

J4

LINE*SCI (P)

UART +

:}

SERIAL

A3U12DRIVER sc, (N) i DATA

A3U7 TO R-T

I 1

DATA BUS TBO - TB6

~

STOPS COUNTERS WHEN MAX LOGICOR MIN TILT IS REACHED GATES

A3U3, A3U5A3U13, A3U19

I

UPIDOWN DIGITAL0E5 (N)* TRI-STATE

COUNTERS COUNT-

BUFFERSP A3U4 A3U14

A3U1O OE6 (N)- A3U17A3U15

b tANALOG

DIA

A3U9

l-lCOUNTERS TO CORRECT ERROR

cAMPLIFIER

t

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A3U18- NETWORK DOWN TO CORRECT ERROR

m -T

A3U21 ANALOG FEEDBACK COMPARATORA3U18

WC-81O Weather Radar ControllerBlock DiagramFigure 8-4 22-14-00

Ao-11S39.R5

Page 198.81/198.82Aug 15/91

Use ordisclosureo!Inforrnallonnn thl~page IS subjecltotherestrictionson thetitlepage01thisdocument

MAINTENANCE

Honeywell !#&!#!h


22-14-00Page 198.83Aug 15/91

Use or disclosure of information on thispage is subject to the restrictions on the title page of this document.

8. C. WA-800 Antenna Pedestal and FP-900 24-Inch Flat-Plate Radiator (SeeFigures 8-5 and 8-6, and Table 8-3.)

D111//];/(,, NOTE: THE FLAT-PLATE RAOIATOR

‘111111,, ,IS SUPPLIED AS A SEPARATE

Ill,,PART AND ASSEMBLED TOTHF

,I; ;:! II IJ,}] ANTENNA PEDESTAL BY iH”i i.kER,

i~!ij~ \ANTENNA PEDESTAL

M15S5354u‘W’ll,l,,,,,“1] .,,, !

ll’’1l’11,},,“1‘4,41,01, 11, ,‘h 11,I‘1 AD.9425

WA-800 Antenna Pedestal and FP-900Flat-Plate Radiator

Figure 8-5

Physical Characteristics:

Weight ........● ............................... 15.25 lb (6.93 kg)

Mating Connector .. .. .. ... ... .. .. ... .. .. .... Cannon KPTM6F20-39S07

“X” Band Quick Disconnect ...... .. .. .. .. .. .. .. ...... Airtron 63906

Operating Characteristics:

Power ................................... 27.5 V dc, 1.3 A nominal115 V ac, 50 mA nominal

Scan ......................................... 60 and 120 degreesScan Rate .................. 14 looks per minute (120-degree scan)

28 looks per minute ( 60-degree scan)Polarization .......................................... HorizontalStabilization ..... Line of sight, +30° pitch or roll with 0° tiltTilt Excursion ................. 15 degrees up and 15 degrees downElevation Excursion ............ 30 degrees up and 30 degrees downBeam Width (3 dB points) ............................. 4.2 degrees

WA-800 Antenna Pedestal and FP-900 Flat-Plate RadiatorLeading Particulars

Table 8-3

22-14-00Page 198.84Aug 15/91

Use or disclosure of information cm this page is subpecf to the restrictions on the title page of thm document,

HoneywellThe RT generates a high level RFantenna. The antenna flat-Dlate


pulse to be transmitted from theradiator transmits the RF enerqy in

narrow beam. By turning the antenna radiator t60 degrees from ~ieforward direction of the aircraft, the narrow beam can be made to scan avolume in the forward direction in order to locate weather targets. If aweather target is encountered, the RF energy is reflected from the targetand received by the antenna during the time between transmission pulses.This received RF energy is routed to the receiver in the RT where it isprocessed for display on the indicator.

The azimuth stepper motor drives the flat-plate radiator left and right60 degrees for a total scan of 120 degrees in accordance with the logicdrive signals from the RT. The sector scan rate is 14 looks per minutefor the 120-degree scan. The antenna pedestal elevation assembly permitsthe radar beam to be tilted up or down 30 degrees in each direction for atotal of 60 degrees when actuated by a combination of signals external tothe antenna pedestal and the aircraft IRS.

ANTENNA

ELEVATION AZIMUTHIOTARYY31NT ROIARYJOINT

)

R-1

mELEVATIONAXIS

m

A21MWHAX15AZlMUWt40T0R CONTINUOUS

* 30 OEG. MAx CONTROLUXIC liOCEG. SCAN

FROM R-T14KloKsfMlN

? Ik’’?!ii?--

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i

I

I

I

I

I

II

/

+1 iI

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CON7RIL

AMPLIFIERS

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DISABLE STEPPER

MOTOR

B303

L-— — — — ——— ———— — — — — -- — — ——

ClSWIEP SENSE

RESOLVERlRb

JBXI I

R-T STABILIZATIONASSY

---;----m~

AD-1 7256

WA-800 Antenna PedestalBlock DiagramFigure 8-6 22-14-00

Page 198.85Aug 15/91

Use or disclosure of mformatlon on this page IS subject to the restrictions on the title page of this document.

MAINTENANCE

Honeywell &w&.WARNING: HEATING AND RADIATION EFFECTS OF WEATHER RADAR CAN BE HAZARDOUS

TO LIFE. PERSONNEL SHOULD REMAIN AT A DISTANCE GREATER THAN“R” FROM THE RADIATING ANTENNA IN ORDER TO BE OUTSIDE OF THEENVELOPE IN WHICH RADIATION EXPOSURE LEVELS EQUAL OR EXCEED 10mW/cm2, THE LIMIT RECOMMENDED IN FAA ADVISORY CIRCULAR AC NO.20-68B, AUGUST 8, 1980, SUBJECT: ‘RECOMMENDED RADIATION SAFETYPRECAUTIONS FOR GROUND OPERATION OF AIRBORNE WEATHER RADAR.”THE RADIUS, R, TO THE MAXIMUM PERMISSIBLE EXPOSURE LEVELBOUNDARY IS CALCULATED FOR THE RADAR SYSTEM ON THE BASIS OFRADIATOR DIAMETER, RATED PEAK-POWER OUTPUT, AND DUTY CYCLE.THE GREATER OF THE DISTANCES CALCULATED FOR EITHER THEFAR-FIELD OR NEAR-FIELD IS BASED ON THE RECOMMENDATIONSOUTLINED INAC NO. 20-68B.

The American National Standards Institute, in their document ANSIC95.1-1982, recommends an exposure level of no more than 5 mW/cm2.

Honeywell Inc. recommends that operators follow the 5 mW/cm2 standard.Figure 8-7 shows the MPEL for the 24-inch antenna and PRIMUS@ 800 radarpower.

—

+2 RADIuSR

\

\/AIRCRAFTLUBBER LINE—-— —

~$)

/y -

270°

/

./MPEL

BOUNDARY

PRIMUS@ 800 MPEL BoundaryFigure 8-7

DIAMETER

II5 mW/CM2

OF FLAT- RADIUS RPLATE OF MPELRADIATOR BOUNDARY

AD-31304

22-14-00Page 198.86Aug 15/91

Use or disclosure of reformation on this page is subject 10 the restrictions on the title page of this document,

~:Ml:~JANCE

Honeywell .ULFSTREAMIV


22-14-00Page 198.87Aug 15/91

Use or dkclosure of information on this page is subject to the restrictions on the title page of this document.

8.1 PRIMUS@ 870 Weather Radar System

A. WU-870 Antenna and Receiver Transmitter Unit (See Figures 8-8 and 8-9,and Table 8-4.)

The WU-870 Antenna and Receiver Transmitter Unit generates and receivesX-band radio frequency energy for the purposes of weather detection andground mapping. The WU-870 incorporates an 18-inch flat-plate antennaand contains all the circuitry required for transmitting, receiving,signal processing, scan conversion, serial data, and control interface

I to the other system components as well as the EDS.

WU-870 Antenna and Receiver Transmitter UnitFigure 8-8

22-14-00Page 198.88Apr 15/93

Useordisclos.rreof information on this page issubject to the restrictions onthetitle page of this document.

MAINTENANCE

Honeywell KN!Rb


Base Diameter ................................ 10.04 in. (255.02 mm)Height (Antenna flat) ....................... 10.06 in. (254.51 mm)

Weight ............................................... 16.0 lb (7.3 kg)

Prime Power ........................ +22 to +32 V dc, 110 watts, maximum26 or 115 V ac, 400 Hz, 0.12 VA maximum

Antenna

Size ..................................... 18-inch flat-plate radiatorStabilization ............................. Line-of-sight, 130 degreesTilt ..................................................... t15 degreesScan ........................................ 120 degrees (*60 degrees)Scan Rate ...................... 13.5 t 1 looks/reinat 120 degrees scanGain ............................................................ 35 dBBeam Width ................................................ 5.6 degrees

Transmitter

Frequency ............................................... 9345 t30 MHzPower ..................................... 1.3 KW, nominal, magnetronPulse Widths ....................... 1.2, 1.5, 2.4, 4.8, 9, and 18us,

determined by selected rangePRF ....................................... 120, 240, 360, and 480 Hz,

determined by selected range

Receiver

Noise Figure ......................................... 8.5 dB, typicalIF Frequency ................................................. 35 MHzIF Bandwidth ......................................... 0.8 MHz, nominalVideo Bandwidth ................. Commensurate with selected pulse widthSTC ............................................. Present in all modesMDS ................................. -112 dBm nominal, on 300 NM range

Dist)layedRanqes

WX/MAP ............................. 5, 10, 25, 50, 100, 200 and 300 NMfull scale with five concentric range rings

(white for WX, green for MAP)

Flight Plan ........... 5, 10, 25, 50, 100, 200, 300, 500, and 1000 NMfull scale with five concentric range rings

Turbulence/GCR Mode .............................. 5, 10, 25, and 50 NM

Mating Connector ........................................ MS3126F22-55S

WU-870 Antenna and Receiver TransmitterLeading Particulars

Table 8-4

22-14-00Page 198.89Aug 15/91

Use or disclosure of Information on this page is subpscf to the restrictions on the title page of this document,

[ MODULATOR I I 117ANTENNA

I

t

I

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CIFICULATOR

‘ -11’?m

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MAGNE~ONI 1 I ~ AGC ~ 1

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CONTROLBuS

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iAAS SELECT (NO

42s ADC (NO) DISCRETE

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PARALLELALTITUDE ‘SELECT (NO)

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—

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RT EFIS

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~LTEF’s~RTEF’s—

AD.14317.R3

WU-870 Antenna and Receiver TransmitterUnit Block Diagram

Figure 8-9 22-14-00Page 198.90Aug 15/91

Use or disclosure of reformation on this page is subject to the restrictions on the title page of this document.

WARNING: HEATING AND RADIATION EFFECTS OF WEATHER RADAR CAN BE HAZARDOUSTO LIFE. PERSONNEL SHOULD REMAIN AT A DISTANCE GREATER THAN“R” FROM THE RADIATING ANTENNA IN ORDER TO BE OUTSIDE OF THE

ENVELOPE IN WHICH RADIATION EXPOSURE LEVELS EQUAL OR EXCEED 10mW/cm2, THE LIMIT RECOMMENDED IN FAA ADVISORY CIRCULAR AC NO.20-68B, AUGUST 8, 1980, SUBJECT: “RECOMMENDED RADIATION SAFETYPRECAUTIONS FOR GROUND OPERATION OF AIRBORNE WEATHER RADAR.”THE RADIUS, R, TO THE MAXIMUM PERMISSIBLE EXPOSURE LEVELBOUNDARY IS CALCULATED FOR THE RADAR SYSTEM ON THE BASIS OFRADIATOR DIAMETER, RATED PEAK-POWER OUTPUT, AND DUTY CYCLE.THE GREATEROF THE DISTANCES CALCULATED FOR EITHER THEFAR-FIELD OR NEAR-FIELD IS BASED ON THE RECOMMENDATIONSOUTLINED INAC NO. 20-68B.

The American National Standards Institute, in theirC95.1-1982, recommends an exposure level of no more

Honeywell Inc. recommends that operators follow theFigure 8-9.1 shows the MPEL for the 18-inch antennapower.

—

+2 RADIUSR

\

\/AIRCRAFTLUBBER LINE—- —+—

/

$$)270°

/

./MPEL

BOUNDARY

document ANSIthan 5 mW/cm2.

5 mW/cm2 standard.and PRIMUS@ 870 radar

IIDIAMETER 5 mW/CM2OF FLAT- RADIUS RPLATE OF MPELRADIATOR BOUNDARY

18 IN.

I

12 “H45.7 CM 388 CM

AD-17728-RI

PRIMUS@ 870 MPEL BoundaryFigure 8-9.1

22-14-00Pacie198.91iug 15/91

Use or disclosure Of reformation on this page is subject to the restrictions on the title page of this document.

8.1 B. WC-874 Weather Radar Controller (See Figures 8-10 and 8-11, and Tables8-5 and 8-6.)

AD-21363

WC-874 Weather Radar ControllerFigure 8-10


Length (from rear of bezel) .................. 7.0 in. (177.8 mm)Width ....................................... 5.75 in. (146.1 mm)Height ...................................... 1.87 in. (47.5 mm)

Weight ........................................... 1.9 lb (.86 kg)

Power Requirements:

Primary ................... +22 to +32 V dc, 8.5 watts,Lighting ........................... 5 V ac/dc at 1.0 A,

Mating Connectors:

maximumnominal

J1 ............................................. MS27473E14A-18SJ2 ............................................ MS27473E14A-18SA

Both connectors use strain relief MS27506-B14-2

Mounting ...................................... Unit Dzus Fasteners

WC-874 Weather Radar ControllerLeading Particulars

Table 8-5

22-14-00Page 198.92Aug 15/91


8.1 B.

The radar system consists of twoWC-874 Controller sets the radar

MAINTENANCEMANUALGULFSTRE4MIV

WC-874 Weather Radar Controllers. Eachmodes, range, antenna tilt, and weather

display brightness on its respective navigation display. If one WC-874Controller is off, the other controller will set the radar modes, rangeand antenna tilt, and display a slave (SLV) annunciation on the OFFcontroller indicating that it is slaved to the settings on the ONcontroller. The controller has four snap-action pushbuttons, two rotaryswitches, and three potentiometersto select and control the operationmodes of the radar system. The following paragraphs describe each ofthe PRIMUS@ 870 modes and features controlled by the WC-874.

(1) Target Alert (TGT) - The TGT button allows the pilot to select ordeselect the target alert mode of the radar system. Target alertis selectable in any WX range except 300 NM. The target alertcircuit monitors for red level or greater targets within t7.5degrees of dead ahead. Also the target must have the followingdepth and range characteristics:

(2)

Selected Target TargetRange (NM) Depth (NM) Range (NM)

5

;:

1::200300

FP (Flight Plan)

2 5-552 - 604 ;: - 754 - 1004 1%- - 1506 200 - 250

Inactive2 5-55

It should be noted that while target alert is functional at theabove ranges, it is improbable that a realistic target would bestrong enough to be detected if its range exceeds five times thedisplayed range. Also, note that the target alert is inactivewithin the displayed range. Selecting target alert preventsvariable gain from being selected.

Ground Clutter Reduction (GCR) Mode - The scintillation frequencyof the ground radar returns is lower than that of rainfall radarreturns. A digital frequency filter is used to separate groundreturns from the rainfall returns, and only the rainfall returnsare displayed when the GCR mode is selected. Since some of therainfall returns fall into the same spectrum as the ground returns,there is some loss of weather return in the GCR mode. As a result,the weather presentation in this mode cannot be consideredcalibrated. However, the GCRmode gives the pilot a dramaticallyimproved look at weather in terminal areas or mountainous terrainwhere it may be necessary to tilt the antenna toward the ground tosee weather ahead. GCR is operational in WX mode and selectedranges of 50 NM or less.

22-14-00Page 198.93Aug 15/91


8.1 B. (3) Rain Echo Attenuation Compensation Technique (RCT) Mode - The rainecho attenuation compensation technique (REACT) function permitsthe radar receiver to adjust its own sensitivity automatically tocompensate for attenuation losses as the radar pulse passes throughweather targets on its way to illuminate other targets. This isdone by measuring the intensity of signals, and deducing from themthe density, and therefore, the attenuation of the target, and thenusing this information to adjust the sensitivity. This is donecontinuously on each radar azimuth radial. There is a maximumvalue to which sensitivity may be set due to the receivergenerating noise, and would fill the display with noise if it weretoo high. When this maximum value is reached, a blue color isdisplayed for the remainder of the displayed range. This gives thepilot an unmistakable warning that attenuation is hiding possiblesevere weather areas that cannot be accurately detected. REACT isalways selected in TEST mode. REACT is available in all modesexcept GMAP.

(4) Turbulence (TRB) - When the turbulence submode is selected, theradar processes return signals in order to determine if aturbulence signature is present. Areas of potentially hazardousturbulence are displayed as gray white. The high power of thePRIMUS@ 870 permits detection of hazardous turbulence in areas ofotherwise weakly reflective rainfall. Any areas shown asturbulence should be avoided. TRB may only be engaged in the WXmode and in selected ranges of 50 NM or less.

(5) Brightness (BRT) Control - The BRT control is used to control radar(raster) information brightness (intensity) of the navigationdisplay (ND).

(6) GAIN - A single-turn rotary control which varies the RTA receivergain. A pull for variable gain position is provided. Selection ofRCT overrides the GAIN control setting and forces preset gain.Rotation of knob CCW from the 12 o’clock (straiclhtup) Dositionreduces receiver gain. Rotation CW from the 12-O’C1;reduces STC so that receiver gain is increased untilmaximum when GAIN control is at the full CW position

(7) Mode Control - The rotary MODE control is used to se”following positions:

OFF - This position turns the weather radar systemoff.

ik”positionit is at

ect one of the

functionally

STBY - Standby is useful for keeping radar in ready state whiletaxiing, loading, etc. In standby, the antenna does notscan, the transmitter is disabled, the display memory iserased, and the antenna is stowed in a tilt-up position.

NOTE: In dual control mode, if one controller is in STBY mode andthe other controller is in an operating mode, (WX, MAP,etc.) the radar is not in STANDBY mode.

22-14-00Page 198.94Apr 15/93

Use or disclosure of informationon this page is subject to the restrictions onthe title page of this document,

MAINTENANCE

Honeywell %%!+%.

8.1 B. (8)

TEST -

Wx -

GMAP -

Standby should be selected anytime it is desired to keeppower on the system without transmitting.

The PRIMUS@ 870 contains a forced standby function (FSBY).This permits an external system, such as the squat switch,to force the radar into standby automatically. If thisoccurs, the mode FSBY will be displayed. If this functionis wired, and should it operate, the user may override itby selecting another mode. In systems with dualcontrollers, this must be repeated on the second controller(-413/414 Controllers only).

Used to select a special test pattern to allow verificationof system operation; 100-mile range is automaticallyselected; TEST is displayed in mode field. Transmitteroutrmt Rower is radiated in TEST mode. Any faults presentwill be displayed when selecting TEST mode. See Table 8-6.

Used to select weather detection operation. If selectedprior to end of the warmup period, WAIT will be displayeduntil the transmitter warms up (approximately 50 seconds).WX is displayed in mode field. Transmitter outcmt Dower isradiated in the WX mode. In the WX mode, fourprecipitation levels are displayed as green, yellow, red,and magenta.

The cn”ound-maDDinaoDeration is selected bv settina themode-control to G~P; The TILT control is-turned downuntil the desired amount of terrain is displayed. Thedegree of down-tilt will depend upon the aircraft altitudeand the selected range. The receiver STC characteristicsare altered to provide equalization of ground-targetreflection versus range. As a result, the selection ofpreset GAIN will generally provide the desired mappingdisplay. However, the pilot may desire to decrease thegain manually by selecting manual gain and rotating theGAIN control. Transmitter outDut Dower is radiated in MAPmode.

Range - Rotary control used to select one of six ranges (10 to 300nautical miles). The control also has a seventh position FPLN.Selection of FPLN causes the radar information to be blanked fromthe navigation display. The display will also be set to a rangewhich is pin-programmable at the J2 connector of the controller.The radar continues to transmit in the mode selected. The EFISwill use the SCI bus from the controllers for range selections ifavailable, otherwise the EFIS will use the range discretes from thecontroller.

(9) TILT - Single-turn rotary control which varies antenna tilt between15 degrees up and 15 degrees down. The range between +5 and -5degrees is expanded for ease of setability.

22-14-00Page 198.95Aug 15/91

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8.1 B. (10) AUTO TILT - Places elevation control under auto tilt which adjustsantenna tilt in relation to altitude and selected range. Tiltknob can be used for fixed offset corrections of up to t2.Odegrees.

(11) COLORCAL” - The COLORCALW feature of the PRIMUS@ 870 COLORADANmore accurately displays the precipitation levels of distantstorms. With the COLORCAL’”feature, distant storms that wouldnormally be displayed as low intensity are shown as heavier, moreaccurate intensity. The radar system cannot detect light rainbeyond a certain range. Beyond this range there is not enoughrain to send back a detectable signal. For the radar to detect atarget at this range and beyond, the target must be at leastmoderate rain. The COLORCALN feature will not let any target atthis range or beyond be displayed as a green (weak) signal. Alltargets in this range will be displayed in yellow, red, ormagenta.

There is also a range where moderate rainfall cannot be detected.At this range and beyond, only heavy and extremely heavy rain canbe detected. The COLORCALN feature will not allow green or yellowtargets at this range and will only show targets as red andmagenta.

NOTE. The boundaries at which these COLORCAL’”changes occur are—.affected by the REACT compensation system if it is active.These boundaries are irregular in REACT. COLORCAL’”isdisabled when in variable GAIN.

(12) Hidden Modes

(a) Roll Offset - This is an in-flight adjustment that can bemade when stabilization errors are detected. To enter thismode, select WX mode and VAR GAIN. Select RCT four times

within 4 seconds. VAR and RCT will not be annunciated.Adjust the GAIN control until the ground returns aresymmetrical. To exit roll offset mode, select RCT four timeswithin 4 seconds. VARwill be displayed.

(b) Stabilization Off - Select TGT four times within 4 seconds.STAB will be annunciated. To enable stabilization, selectTGT four more times within 4 seconds.

(13) SLV - Indicates that displayed data is controlled by the oppositeside controller.

22-14-00Page 198.96Aug 15/91

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MAINTENANCE

Honeywell YJW&.

EFIS Fault Codes-413/414 -415/416

Cent ro 1ler Contro 1ler Fault Description

None None No fault

01 21 Azimuth scanning incorrectly (> 2.5 degrees for > 2 seconds)01 31 Antenna Elevation error (> 2 degrees for > 2 seconds)

02 03 Analog to Digital Converter Failure02 22 STAB reference (< 1/2 A/D scale for > 2 seconds)02 32 NAV Computer High Speed ARINC429 Failure

03 13 +15 Volts Failure (> *1.5 Volts)03 23 Automatic Gain Control Failure (c03 33

-1 V or > 9.73 V for 8 seconds)-15 Volts Failure (> *1.5 Volts)

04 16 Magnetron Voltage Failure (< 1500 Volts min or > 2700 Volts max)04 24 Mixer Current Failure

05 25 AFC Lock Failure05 35 AFC Sweep Failure

06 26 Fan Voltage Abnormal

070707070707070707

040714172734363730

Digital Air Oata FailurePulse Pair Processor FailureParallel Altitude FailureEPROMTest FailureVLSI Test Failure-Loss of Video Ready InterruptDADC Altitude FailureAnalog Altitude Failure - If input is > 60,000 feetRAM Test FailureNonvolatile !krnory Failure

~: When reporting fault codes, it is essential that the type of controller (e. g., -413) be reported.

Fault Display FormatTable 8-6

22-14-00Page 198.97/198.98

Aug 15/91Use ordisclosureOfInformationon thispagek subject10therestrictionson thetitlepageofthisdocument.

OE1 (N)

h

TGT\

,2 BUFFER

/ A2U1OTBO – TB6

/GAIN SW

FSB1

%

COMPA2ul

0E2 (N)

/4 BUFFER/ A2U1O

96 KHZ

t

PROGRAM

CTRA3U19 AND A3U16

1

-El--@A2U5

BINCTRWU21TBO - TB6

GCR

GAIN WIPER

OE4 (N)3 ,5 1 ,5

+’

/

i

CONTROL LABEL

PROM PROM

A3U13 A3U8

I

1

AD

*CONVGAIN

TBO - TB6

A3U3D-A3U1AOP AMPL

*

A3U1 BOP AMPL

I192KHZ

LUART

A3U4 uLINEDRIVER SclA3U2

TO RTAINTERNAL BUS

AIDCONVTILTA3U5

TILT WIPER

TBO - TB6

BOE1 (N)

DECODER OE2 (N)

MUX OE3 (N)

10F8 0E4 (N)

A3U14 0E5 (N)

OE6 (N)

OE7 (N)

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Use or dmclosure of Intormallon On this

Figure 8-11 22-14-00Page 198.99/198.100

Aug 15/91page IS subject 10 the restrictions on the title page of Ihls document

MAINTENANCE

Honeywell 8!!W%L.


22-14-00Page 198.101

Aug 15/91Use or disclosure of information on this page is subject to the restrictions on the title page of thts document,

9. FMZ-800 Fliqht Management System (FMS)

I ~. NZ-920 Navigation Computer (See Figures 9-1 and 9-2, and Tables 9-1, 9-2,and 9-3.)

Ao-lle41

NZ-920 Navigation ComputerFigure 9-1

22-14-00Page 198.102



Length ...................................... 17.03 in. (432.6 mm)Width ........................................ 4.91 in. (124.7 mm)Height ....................................... 7.62 in. (193.5 mm)

Weight (approximate) .............................. 14.81b (6.71 kg)

Power Requirements .................................... 28V dc, 65W

Mating Connector:

J1 ........................ Cannon Part No. DPX2-67S-106P-33B-OO89

Mounting ....................... Tray, Honeywell Part No. 7003272-903

NZ-920 Navigation ComputerLeading Particulars

Table 9-1

The NZ-920 Navigation Computer provides many varied navigation functions;however, the primary function is to provide high-accuracy long rangelateral and vertical navigation. To accomplish this function, thenavigation computer connects to a variety of sensors.

The prime radio navigation inputs are VOR bearing and DME distance. Thismay be broken down into two categories -- VOR/DME and multi-DME. SomeDMEs can be commanded to scan two or three channels, while others supplydistance from a single channel (station). The navigation normallyoptimizes the present position calculation accuracy by utilizing DMEdistance data from at least two stations versus VOR bearing and DMEdistance data from a single station. The VOR bearing input is subject toerror due to bends in the beam, multipath effects, station misalignment,etc. Using multi-DME, the bearing information from a station may becalculated instead of simply measured.

Two types of radio (VOR/DME) configurationsare supported. First the FMSwill support a system with single channel DMEs. In this system, both theon-side VOR/DME and the cross-side DME are brought into the navigationcomputer. The second system configuration supported by the FMS is adirected scanning DME. In this configuration,multiple distances can bereceived from a single DME receiver. In this configuration, distancesfrom two stations can be utilized for a more precise position fix.

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The navigation computer also has input capability for three long rangesensors through ARINC 429 interfaces. The navigation computer can accepttwo more long range sensors through the ASCB interface for a totalcapability of five sensor inputs. These sensors will typically be threeIRS sensors, though Omega or GPS sensors can be substituted in place ofIRS units. These long range sensors are utilized in addition toVOR/DME/DME inputs for overland flight. The IRS, GPS, and Omega/VLFinputs are the only navigation source inputs utilized when VOR/DMEsignals are not receivable. The navigation computer will automaticallychoose the best navigation combination (VOR/DME, IRS, Omega, GPS) basedon predefined priority. When using VOR/DME inputs, a blending of theseinputs and IRS and GPS information occurs. This blending is done viacomplementary filtering. Filtering lessens the effects of error and noisein both the VOR/DME, IRS, and GPS inputs and, thus, provides a smooth andaccurate position derivation.

NOTE“ GPS blending was activated in the -976 NAV computer and is for‘“ display only in the -963, -964, and -978 Navigation Computers.

The navigation computer provides automatic tuning of the aircraft VOR andDME receivers. Calculation of aircraft present position from VOR/DMEinformation requires input of bearing and distance and knowledge of thestation coordinates. The database is periodically used by the navigationcomputer to find the coordinates and frequency of the high and lowaltitude VORTAC and VOR/DME stations in the aircraft vicinity. When thedesired VOR and DME stations are chosen, the frequency is output to thenavigation receivers. Automatic receiver tuning is operationallytransparent to the pilot other than a periodic change in the receiver’sfrequency display and the RMI pointer. Provision is included for remotetuning of receivers via the CDU or manual tuning through the radiocontrol heads. For remote tuning via the CDU, the pilot can choose toenter the station identifier or enter the frequency. The frequency of theentered station found in the database is output to the navigation controlheads, then to the receivers. For manual tuning via the radio controlhead, the navigation computer will input the frequency code from thereceiver and compare it to the frequencies of stations in the aircraftvicinity. The frequency comparison will allow the navigation computer todeduce what station is being tuned. A comparison of calculated bearingand distance to input bearing and distance will resolve the remotepossibility of two or more stations in the aircraft vicinity transmittingon the same frequency.

An important part of the navigation computer is the nonvolatile memoryarea or database which contains information on navaids, airports, andairways. The database is integral to the navigation computer to allowquick access of the stored information. The database is reprogrammableand is loaded with worldwide data. The worldwide data includes VOR,VORTAC, VOR/DME, airport reference points, runway thresholds, and highaltitude airway intersectionsplus airway routes and SIDs/STARsprocedures. The navigation data requires updating on a 28-day cycle.Updates are required due to changes in the data associated with eachstation or airport. The database memory is also used to store pilot-defined waypoints and pilot-defined routes; however these are notupdated every 28 days. They reside in memory until changed by theoperator.

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The navigation computer has a database capacity of 1.28 megabytes whichenables loading of data to allow international operation without changingdatabases. Also, the computer contains an internal battery that keepsthe clock and calendar running when power is removed.

The navigation computer provides a lateral steering signal to the DAFCS.The DAFCS will cause the aircraft to bank according to the lateralsteering signal. The lateral steering signal is proportional to thecalculated distance and angle deviations from the desired lateral course.Lateral steering mode is a desired track mode.

In addition to providing a lateral steering signal, the navigationcomputer also provides vertical navigation (VNAV). The VNAV modes areVNAV altitude (VALT), VNAV altitude preselect (VASEL), VNAV flight levelchange (VFLCH), and VNAV vertical path guidance (VPATH). The verticalcommand is proportional to the calculated distance from the desiredvertical path. VNAV allows the pilot to define waypoint altitudes anddescent angles to waypoints and command the DAFCS to pitch the aircraftand fly the desired vertical path.

The navigation computer also provides guidance and map information to thepilot via the navigation display. The navigation computer will output tothe Nav display the positions of 51 waypoints (lateral and vertical), 10closest Navaids, and 9 closest airport reference points as well asholding pattern information. The display will include the course linebetween the waypoints. The waypoint and course line display will givethe pilot a pictorial view of the waypoints with respect to the weather.

The navigation computer has a multi-microprocessor architecture (Figure9-2). The primary 16-bit processor has a coprocessor for performingfloating point mathematics. The primary processor (pair) handles all thenavigation and guidance calculations as well as the database manipu-lations and CDU input/output. The remaining two 16-bit microprocessorsare dedicated to handling the sensor inputs and navigation outputs to theEFIS and autopilot. The two 1/0 processors communicate to the mainprocessor through shared random access memory.

The navigation computer digital outputs transmitted over the ASCB arelisted in Table 9-2 for basic data and Table 9-3 for background data.

22-14-00Page 198.105

Apr 15/93Useor disclosure of information on this page issubject to the restrictions onthetitle page of this document.

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51 a LTS #2 LONGVALID l=VALID/O=INVALID51 7-o LTS #2 SENSORLAT (msb)

---1------1-------------------------l----l-------------------------l--------------------l---l-----------------l-----------------l----l--------J52 1S-0 LTS #2 SENSORLONG(msw)

-;;k;--l-------------------------l----l---- ---------------------1 ;-------------------1 ---1- ----------------l~--;;------------l----l--------lLTS #3 SENSORLAT (lSW) SQ41CIRCLE - 960 DE N-S

53 023 0.000043 DSC

LTS #3 LAT VALID

‘-- ------ liii-;i-iiiii-tii-ii:ij- l---- l~:~fi:!::fl:::--------lI ----;;;-;ffi-;-;-----l ---l -----------------l;;;-------------l----l--------l54 15-9 SRiICIRCLE o-54 8

23 0.000043 DIXLTS #3 LONGVALID l=VALID/O=INVALID

54 7-o

-i;l;j:i--l;:-::-:::-:i:::;-l----l-------------------------l--------------"-----l---l-----------------l -----------------l----l--------1-;;l;;:;--1;;~--------------------l----l-------------------------l--------------------l---l-----------------l-----------------l----1--------l

-;;l;;:;--l;;~--------------------l----l-------------------------l--------------------l---l-----------------l-----"-----------l----l--------l

-;;l;;:~--l;;~--------------------l----l-------------------------l--------------------l---l-----------------l-----------------l----l--------l

---l------ l-------------------------l----l-------------------------l-------------------- l---l-----------------l-----------------l----l--------l

0 PNCS BACSOROIINDCODE0001 - HrsO DATA - VURSION1.0

-,(n0?c0’-,(8~

Du)c1Wz

-IT’-5U)

+5%mmU3Z-1 A.*me<

WSP BIT BIT PUNCTION NOTEFORMAT

-;;l------l-------------------------l----l-------------------------l:----------------l~~l-~:-------------l~-~~~--------l~~~l~~fl~--lNAVOPERATIONFODE

59 1s-8 STATE INTEGER C(W859 7-0 SUB-STATE INTEGER C@E

-;;l;;:;--l-------------------------l----l-------------------------l--------------------l---l-----------------l-----------------l----l--------lCUSTONDATA CNECKSffl

-;;l;;-;--l-------------------------l----l-------------------------l--------------------l---l-----------------l-----------------l----l--------lT~ONARY MTA CtlECRSUN

I - l;;~--------------------l----l-------------------------l--------------------l---l--... ------ ---------------l-----------------l----l--------l62 15-562 4 ACTIVE NAVCYCLE L~IC 1 = NSU62 3 SINCLE SYSTEN BIT LOGIC 1 = sINGLE62 2 NASTER/SLAVESTATUS 1 - MASTER O = SLAVE62 1 CUSTOMCNECKSUNVALID LOCIC 1 - VALID62 0 TE14PORARYCNECKSUNVALID

IL(P31C 1 = VALID

‘-- ------l-------------------------l----l-------------------------l--------------------l---l-----------------1-----------------l----l--------\63”15-0 “SPARE (USP 63 - 76) - ‘

-;;l------l~;;;-;~----------------l----l;~-;;-;;;---"--------- l--------------------l---l-----------------

-;;1------ l~;-J;;;--------------l ---- l;;;---------------------- l-------------------- l--- l -----------------

----------------- l----

I----.-.,---------- ----

I‘-- ------l-------------------------l----l-------------------------l--------------------l---l-----------------l-----------------l--79 NDLC FLAG

I l-------------------------l----l::-~---------`---------l------------"-------l---l-----------------l-----------------l------- ---.--

-------- I-------- I-------- I-------- I

Pt4CS BACSGROUNOCODE0002 TNRU000C - WAYPOINTS- VERBION1.0

WSP BIT BIT FUNCTION NOTEKSWAT

I l;;i-;;;----------------l----l;;-~-------------------l:----------------l~~l-:::-------------l::-:::::--------l::~l:flfl--l--- ------

--:l------l-------------------------l----l;~;-~;~-------------l--------------------l---l-----------------l-----------------l----l--------lF14CSCONTROL

1 15 TEST LCMC 1 - TEST1 14 vALID LffiIC 1 = VALID1 13 DATASASETRANSFER LOGIC 1 - TRANSFER1 12-11 SPARE1 lo-s COUNTER o-7Rmc

D 1 7-o F14CSADORSSS 12u - LSPT 16 E - RIGNT(/Jn ‘-- ------l -------------------------l----l-------------------------l--------------------l---l-----------------l-----------------l----l--"-----lIm= 2 15-0 SSADERcoos 0002 B = ORIGIlt,DESTINATION

+72

-m 2

+s:2

W(n 2U32Z-l A.gJ 2

(Drt~ 2

m(DtO- 2I aflJ 2

wm ~. 2

-@o 2003or 2=U2m+%0

-03 --:1 ------cmSCn+ 3 15-8

0?3 7-o

al I--- --.-.-

NN

“*w

4 15-84 7-o

---1 ------5 15-s5 7-o

---1------6“6 156 146 136 126 116 106968

ALTMTE DESTINATIONTOP OF CLLMS,-1 UAYPOINT0003 a - 0 - 4 UAYPOINTS0004 E - 5 - 9 WAYPOINTS0005 U - 10 - 14 WAYPOINTS0006 B = 15 - 19 WAYPOISTS0007 E = 20 - 24 NAYPOISTS000S B -25 - 29 WAYPOINTS0009 u - 90 - S4 WAYPOINTS000A II = 95 - S9 WAYPOINTS000B E = 40 - 44 NAYPOINTS000C ii = 45 - 49 HAYPOINTS

-------------------------l----l-------------------------l--------------------l---WAYPOINT#() IDENT

CNARACTSR2 ASCIICNARACTER1 ASCII

--~;;~-;------------l ----l-------------------------l--------------------l---ASCII

CNARACTSR3 ASCII

--m;;:-;------------l----l-------------------------l--------------------l---ASCII

CNANACTER5 ASCII

----------------- -----------------l----l--------1I

----------------- -----------------l ----l --------1I

----------------- -----------------l----l--------1I

-------------------------l----l-------------------------l--------------------l---l-----------------l-----------------l----l--------llJAYPOINT#() DISCRBTSS #l

VALID WAYPOINTPRIHARYWAYPOINTNOLOAT UAYPOINTDRAWCNBVRDNNONDISPLAYABLEWPTWAYPOINTDISCONTINUITYRADIALRSSERVSD

1 - UPT DATA IS VALIO1 = WAYPDINTIS PART OF PRIMARYFLIGHT PLAN1 = HOLDAT WAYPOINT1 = DRAWCNSVRON1 - DO NOTDISPLAY ID2NT1 = DIS~NTINUITY1 = RADIALDRAWOUTOF WPTPAD ZSRO

FHCS BACKfWNINDCOl)K 0002 TlllUJ 000C - UAYPOINTS - VERSION 1.0

NN

USP BIT Brr FUNCTION NOTE FORUM

--il;-----l~;;&-----------------i----l-------------------------l:----------------l~:l-~:-------------l~-~~~~"-------l~~~l~flfl--lPAO ZERO

66 RssERvm PAD ZSRO65 RSSERVHI PAD Zmo6 4-0 SPARE

I l;;&--------------------l----l-------------------------l--------------------t---l-----------------l-----------------l----l--------i--- ------7 15-0

--;l;;-;--l-------------------------l----l-------------------------l--------------------l---l-----------------l;---------------l----l--------lUM#()AL?coImNAItn TWS c@lPL~Z +1- 163840 FEET 15 10.0/s.0 FSET

FLIGNT LSVEL DISCNETE

--;l;;-i--l-------------------------l----l:::::-::!:::=---------l--------------------l---l-----------------l;---------------l----l--------lUPT #() ALTSRNATEMT TUOoS CC@lPL2TiENt +/- 163840 FEET 15 10 .0f5. O FEET

90 FLIGBT LSVEL DIS(XETE 1-FLT LVL/O=FEET

‘-- ------l -------------------------l----l-----"-------------------l;-:-;;;-;-;---------l---l-----------------l:;;;------------l----l--------lI10 15-0 UATPOINT#() LA? (lSW)

II

SMICIIWLE--- ------ -------------------------l----l-------------------------l;-:-;;;-;:;---------l::-l::::::::-::-----l---------"-------l----l--------l

11 15-s WAYFOINT1() LONW2UIWlsb) SBIICIRCLE11 7-o

24 0.000021 DEG MSTLATITUDE (msb)

---l------l-------------------------l----l-------------------------l--------------------l---l-----------------l-----------------l----l--------l12 15-0 lDNGITUDE(msw)

-;;l;;-;--l:~;;-;;;-;;;;;-;;;;;-l----l-------------------------l;-------------------l-`-l-----------------l-----------------l----l--------l

I32 0. 000000083/ D~ EAST

--- --~---l-------------------------l----l~~~~~~u~---------------l--~-~~~-~~---------l---l-----------------l----14 15-0 C0URS6 INTO UPTS( ) (-v)

I l:~;:-;;;;fi;;-;;;;;--l----l-------------------------l:;-;;;;;-~---------l-`-l-----------------l-----------------l----l--------l.-.. ------

15 15-0 TUO’S CDMPLRfR!T

-iilii-;--l-------------------------l----l-------------------------l--:-----------------l::-l::::::~:-:-----l:-:::::~~------l----l"-------l-i;l;;:;--l:~:*-~flfl::-~~--l---- l------------------------- l-----”-------------- l---l -----------------1---------”-------l ---- l--------lNSNTUAYPOINT DATA( 17-30

I - l-------------------------l----i-------------------------(--------------------l---l-----------------l-------------=---t----l--------i--- ------

31 15-0 NEXT UAYPOINT DATA ( 31-44)

I l-------------------------l----l-------------------------l--------------------l---l-----------------l-----------------l----l--------l--- ------45 15-0 NENTUAYK)INTDATA (45-5S)I l-------------------------l----l-------------------------l--------------------l---l-----------------l-----------------l----l--------l--- ------

59 15-0 NEXT UA~INT DATA(59-72)

I l-------------------------l----l-------------------------l--- -.---- --------------------l ---l -----------------l-----------------l----l--------l12 15-0 SPARE (72-76)

-;;l------l;;~;l-;~----------------l ----l;:-;;-~--;------------- l-------------------- l--- l----------------- l----"------------l ----1--------1

II--- ------ ---;;---;-------------- I---- l;;;---------------------- l-----”--------------l ---1-----------------I ---------”------- l----1--------1-~~1------1----------"-------------- l---- l-------: ----------------- l--------------------l ---l -----------------l -----------------l ---- l--------l

-::l------l~~~-~~~----------------l----l~~-~-------------------l--------------------l-`-l-----------------l-----------------l----l--------l

FMCS BACKGROUNDCODEOOOD- NAVAIDS - VERSION 1.0

o

-.mmcu-.(vc)

USP BIT BIT FUNCTION NOTEFORMAT SCALE R2B LSB POS SENSE FTIU SINVAR

---l ------l~;;-;;;----------------/----l;;-~-------------------l--------------------l---l-----------------l-----------------t----l--------l--:l------l-------------------------l----l-------------------------l--------------------l---l-----------------l-----------------l----l--------l

1“ “FMCSCONTROL PACKSDLOCIC1 15 TEST LOCIC 1 = TEST1 14 VALID LOGIC 1 = VALID1 13 DATABASETRANSFER LOGIC 1 = TRANSFER

D1 12-11 SPARS

m 1 10-8 cOUNTER o -7H~nm=

1 7-o PMCSADDRESS 12 H - LEFT 16 H - RICHT

4?--;l;;-;--l~;~-;;;;--------------l----l;;;;-~-----------------l--------------`-----l---

-3- --;1;;;;--1-------------------------l----l-------------------------l;-:-;;;-;-;---------l;;-4!?2? NAVAID #1 LATITUDE(lSW) S~ICIRCLE!%Zz I--- ------l-------------------------l----l-------------------------l---------~----------l---

-------------- ---

------------------0.000021 D=-----------------

0.000021 DEG

-----------------l----l--------l

-----------------l----l--------1NORTR

-----------------l----l--------1

EAST

-<.-mrt~ 4’7-0 ‘ LATITUDE (msb)

4 15-8 NAVAIDLONGITUDE(1sb) SQ41CIRCLEmma”I am --; I;;-;-- I-------------------------I---- l-------------------------l~-:-~~~-~~~---------l~~-l-----------------1-----------------1----1--------1u LONGITUDE(maw)

w ~.-Wo --+-~---l ------------~------------l---- l-------------------------l--------------------l---l-----------------l ----------------- l---- l--------l00=$ 6 15-5 NAVAID#1 ELEVATION TkO’S CCNIPLRfENT +1- 20480 FEET 11 20.0/0.625 FT POS MT

2s0 6 4-O SPARE*-O

-03 --;1------1-------------------------1----1--- ---------------------- l--------------------l ---l ----------------- l----------------- l---- l--------lc-o NAVAIDTYPE

2:7 15 NAVAIDVALID LCGIC 1 - VALID7 14 DME LOGIC 1 - DFIE

05w

7 13 VOR LOGIC 1 = VDR

* 7 12 ILS LOGIC 1 = ILSw 7 11 nLs LfxlC 1 = ULS

7 10 NOB L@31C 1 =79 TUNED LCCIC 1 =7a DISPLAYBDLEFT LOGIC 1 -77 DISPLAYEDRIGHT LCGIC 1 =76 SPARE75 VOR/DNECOLOCATION LOGIC 1 =7 4-o SPANS

---l------l-------------------------l----l----------

NDBTUNEDBY THIS NAVCC+IPUTERLEFT VORDISPLAY STATION (NAVAID #l&#2 ONLY)RIGHT VORDISLAY STATION (NAVAID#lL#2 ONLY)

$NON-COLDCATED‘

I I--------------- -------------------- ---

8 NAVAID 41 IDENT8 15-8 CNARACTER 2 ASCII

8 7-O CNARACTER 1 ASCII

‘-- ------l-------------------------l----l-------------------------l--------------------l---I

-----------------l-----------------l----l--------l

-----------------l-----------------l----l--------l

Slfcs BAWMUNJNDCODE000D - NAVAIDS - v~NSION 1.0

0

-.w

o

i+w

USP BIT BIT PUNCTIDN NOTEFORMAT

I l--i=ii=-i------------l ---- l-------------------------l:----------------l~l-:::-------------l::-::::~--------l~::l::flfl--l--- ------9 15-8 ASCII9 7-0 CHARACTER3 ASCII

I l-------------------------l----l-------------------------l--------------------l---l-----------------l--- ------ -----------------l---- l--------l10 NAVAID #2 (10-16)

I--- ------ l-------------------------l----l-------------------------i--------------------l---l-----------------l-----------------l----l--------l“NAVAID #3 (17-23)

. .

-;:l------liiiiii-ii-iii-ji;--------l----l-------------------------l--------------------

II--- ------;~;;;;-;;-;;;:;; ;--------l ----l-------------------------l--------------------

-:i------l-------------: -----------i ---- l-------------------------l--------------------NAVAID #6 (38-44)

-;;l----"-l-------------------------l----l-------------------------l----"---------------

I---------l~~~-!~-~f~~f~~--------l----l-------------------------l--------------------52 NAVAID48 (52-58)

---l-----------------1 -----------------1----1--------1

---l-----------------l-----------------l----l--------l-------------------- ----------------- l---- l--------1I I

I I--- ----------------- ------------------

--- -----------------l -----------------I

----1--------1

II---- --------

-;;1------1-------:--~-----~--------l----l-------------------------l--------------------l---l ----------------- -----------------i ---- l--------1IHAVAID 49 (59-65)

-;;1------1-------------------------l----l-------------------------l--------------------l---l-----------------l -----------------l ----l --------1NAVAID#10 (66-72)

-;;1------1-------------------------l----l-------------------------l--------------------l---l-----------------l-----------------l----l--------lSPARB (73-76)

II--- ----.- -------------------------l----l;~-;;-:;;-------------l--------------------l ---l-----------------l-----------------l---- i--------l77 CHECNSOM

II--- ------ -------------------------l----l;;;----------------------l--------------------l---l -----------------l-----------------l ----1--------178 SRRORCHECK

II--- ------ -------------------------l----l;;-&-------------------i--------------------l ---l -----------------l-----------------l ----1--------1-~~l------l~~:-~~~----------------l----l-------------------------l--------------------l ---l -----------------l ----------------- l---- l--------l

—.——- ...-—.---- -.——_.-— -------- .— ---- . . -

-aQJ

NN

USP BIT BIT PUNCTION NOTEFORMAT

I l;~-;;;----------------l----l--------------------"----l:~---------------l~~l-~~~-------------lfl~-~~~~~--------l~~~l~flfl--l--- -.----

--fl------l;;;-~;~-------------l----l::-~-------------------l--------------------l---l"----------------l---"-------------l----l--------l1 PACSEDLOGIC1 15 TEST LOGIC 1 - TEXT1 14 VALID LoGIC 1 = VALID1 13 MTASASE TRANSFER LOGIC 1 = TNANSFES1 12-11 SPARS1 10-8 COONTER o- 7 Em1 7-0 E?ICS ADDRESS

II

12 H- LEPT 16 H - RIGET--- ------

-------------------------l----l;;;;------------------- l-------------------- l---l -----------------l-----------------l----l--------l2 15-0 E2ADERCODE

I l-------------------------l----l;;;;;;;---------------l;-:-;;;-;";---------l---l-----------------l~;;;------------l----l--------l--- ------3 15-0 AIRP~T #1 LATITUDE(law)

--:l;;-;--l~;:;-;;-;;;;~;;;;;;;l----l;:;;;i;---------------l;-:-;;;-;:;---------l:-l;::~-~-----l~;;-------------l----l------"-l

4 7-o LATITGDE(sub)

i l;~;;;;-;:;;;----" -----l----l------" ------------------l--------------------l---l-----------------l-----------------l----l--------l--- -------5 15-0

I l-------------------------l----l-------------------------l--------------------l---l-----------------l----------6“ ‘AZRPORT#1 RSVATION “ “6 15-5 ELEVATION TkN3’S CCt4PLR4SNT +/- 20460 FEET6 k-O SPARS

11 20.0/0.625 FT

I l-------------------------l----l-------------------------l--------------------l---l--------------------------

P@ ALT

7’ ,. .----------------l ----l--------l‘AIRPORT #1 USE

. . .

LOGIC 1 - VALIDPAD ZEROPAD ZEROPAD ZEROPAD Zmo

7 15 AIRPORTVALIO7 14 RESERVED7 13 RESERVED7 12 RESERVEO7 11 RMERVEO7 lo-o SPARE

I l---------------------------- ------a 15-o AIRPORT#l RUNNAYCOURSEI l---------------------------- ------

9 AIRPORT#l IDENT9 15-8 CHARACTER29 7-o CHARACTER1

---- ------------------------- ;-:-;;; -;:---------l---l-----------------i;;;----------̀--J----l--------lI IS~ICIRCLE

I

16 0.00549---- ----------- -------------- I-------------------- l---l -----------------l -----------------l ----1--------1

ASCIIASCII

------------------ I-----------------l ----l--------l-J;#- -----------------------1---- l-------------------------l--------------------l ---CHARACTER4 ASCII

10 7-o CNARACTER3 ASCII

--- ------ l-------------------------l----l-------------------------l--------------------l--- -----------------l -----------------l ---- l--------lI

-~~l------l~~~~~-!~-~~~~~~~-------1----l-------------------------l--------------------l ---l ----------------- l-----------------l ----1--------1

o.

0

1+w

FM(M BACKGROUNDC411M000E - AIRPOltTS - Vm810M 1.0

USP BIT BIT FUNCTION NOTEFORMAT

I l-------------------------l----l-------------------------l~----------------l~~l-~~-------------l~-~~~~~--------l~~~l~flfi--l--- ------19 AIRPORT#3 (19-26)

I l-------------------------l----l-------------------------I--------------------l---l-----------------l-----------------l----l--------l--- ------27 AIRPORT #4 (27-34)

I--- ------l-------------------------l----l-------------------------l--------------------l---l-----------------l-----------------l----l--------l-::l------l:::::-!~-f :~::::-------1----l-------------------------l--------------------l---l-----------------l-----------------l----l--------l

43” ‘AIRPORT #6 (43-50) “

I l-------------------------l----l-------------------------l----------------------------- 1---51 AIRPORT #7 (51-58)

I l-------------------------l----l-------------------------l--------------------l------------59 AIRPORT 48 (59-66)

--- ------l-------------------------l----l-------------------------l--------------------l---I63 AIRPORT #9 (67-74)

----------------- -----------------l----l--------1I

----------------- -----------------l ---- l--------1I

----------------- -----------------l----l--------1I

-;; I------I -------------------------1----l-------------------------l--------------------l---l-----------------l-----------------l----l--------lSPARE (ltSP 75-76)

-;;l------l;;;;;-;;----------------l----l-------------------------l--------------------l---l-----------------l-----------------l---- 1--------1SUMOF UOROS

-;;l------l-------------------------l----l;;;---------------:------l--------------------l---l-----------------l -----------------l ----l --------1mROR CNsm

I l;;;-;;;----------------l----l;;-~-------------------l--------------------l---l-----------------l-----------------l----l--------l--- ------79

I l-------------------------l----l-------------------------l--------------------l---l-----------------l---------

ix

FMCS BACKGROUHOCOOE000F - FFPF - PRIVATE FMCSDATA - VERSION 1.0

WSP BIT BIT FUNCTION NOTE FORMAT

--j------ l-------------------------1---- l;i-=------------------- l:----------------l~::l -~~~----------”--l~=~~--------lw~~%HOLCFLAG

--j------l -------------------------l----l-------------------------l --------------------l---l-----------------l-----------------l----l--------lFMCS CONTROL PACKSDLMIC

1 15 TEST LOGIC 1 = TEST1 14 VALID L@21C 1 - VALID1 13 DATABASETRANSFER LOGIC 1 = TRANSFER1 12-11 SPARE1 10-8 COUNTER o- 7 HEK1 7-o 3?ICS ADDRESS 12 Ii - LSFT, 16 H - RIGHT

--;1;;:;--1-------------------------1---- l-------------------------l--------------------l ---l -----------------1-----------------1‘--- 1--------1NEAOERCOOE 000FH - SPARE

001111 - DUAL NAV PRIVATE FLTPLN MSC DATA001211 - DUAL NAV PRIVATE PRIGIN, DESTINATION,

ALTERNATEDESTINATION, TOP OF CLINB,-1 WAYPOINT

001311 - DUALNAVPRIVATE O - 4 UAYPOINTS001 4tl = DUAL NAV PRIVATE 5 - 9 WAYFOINTS0015tl - DUALNAVPRIVATE10 - 14 UAYPOINTS00161J - DUAL NAV PRIVATE 15 - 19 WAYPOINTS001 7H - DUALNAVPRIVATE 20 - 24 WAYPOINTS0018H = DUALNAVPRIVATE 25 - 29 WAYPOINTS0019H = DUALNAVPRIVATE 30 - 34 WAYPOINTS00 IAN - DUALNAVPRIVATE 95 - 39 UAYPOINTS00IBH = DUAL NAVPRIVATE 40 - 44 WAYPOINTS00ICt! = DUALNAVPRIVATE 45 - 49 WATPOINTS00IDH THRU 5555tl SPARE

--;l------l-------------------------l----l-------------------------l--------------------l---l-----------------l-----------------l----l--------lPRIVATE F?4CSDATA

--~1---1:==:l=------------l ---- l-------------------------l--------------------l---l -----------------l-----------------i----l--------l-1-----l=-:l--”-------------l ----1~~-==------------ I--------------------l---l-----------------l-----------------l----l--------l

N -::&-.- .,. ERRORCHECK “CRC

-------------------------l----l;~-~-------------------l --------------------l---l-----------------l-----------------l----l--------lNOLCFLAG

-------------------------l----l-------------------------l --------------------l---l-----------------1- ----------------l----l--------1

RS232 RCVR

RETURN

RS232 XMTR

RS232 RCVR

RETURN

RS232 XMTR

RS232 RCVR

RETURN

RS232 XMTR

RS422 RCVR

{

(H)

CDU DATA (L)

{RS422 XMTR (H)CDU DATA (L)

RS422 CDU

{

(H)

CNTL RCVR (L)

RS422 XMTR

{

(H)

CLK CDU (L)

RS422 CDU

{

(H)

CNTL XMTR (L)

RS422 RCVR

{

(H)DATA LOADER (L)

RS422 XMTR

{

(H)

DATA LOADER (L)

RS422 XMTR

{

(H)

DATA LOADER (CLK) (L)

JIA JIB

55565556

28

29

52

53

1011

12

13

14

15

17

18

19

7

8

35

36

32

33

20

21

D——— ———— ———— ———— ———— ———— — 1

=1

i

RS232INTERFACE

1

, 1==1 DATA

LOADER

m=+%14ARED 1/0 cRAM PROCESSOR

LEN

BUSINTERFACE

CKTS

1011

35

X16

17

18

19

23

24

PROCESSOR A SHARED 1/0 B

CO-PROCESSORRAM PROCESSOR

*TRANSMITTER *

h—TO BUS 1 TO BUS 2 ~(SH2) t

i(SH2)

sL ———— ———— ———— ———— ———— ———— 4

NZ-920 Navigation ComputerBlock Diaaram

Figure 9-2 (~heet 1) 22-14-00

262732334344

45465051—

19

2831

22235758

AIRCRAFTBATTERY +28 V DCAIRCRAFTBAITERYRETURN

[H)}PRIMARYASCB BUS

[L)[H)

}SECONDARY ASCB BUS

[L)[H)

}AFIS/ACRSRX BUS

[L)

[H)}MLWLS PRIMARY

[L)[H)

}DME PRIMARY

[L)[H)

}LTS NO.2

(L)(H)

}LTS NO.1

(L)(H)

}NAV PRIMARY

(L)(H)

}NAV SECONDARY

(L)(H)

}DME SECONDARY

(L)(H)(L)}

LTS NO.3

(H)}GEN BUS SECONDARY

(L)(H)

}GEN BUS PRIMARY

(L)

AD-30257,SH1#

Page 198.123/198.124Apr 15/93

USe or disclosure of information on this page is subject to the restrictions on the title page of this document.

.—— ——— ——— ——— — ——— —— ——— ——— — 1Cnnti tJIAJIBIle

I03

I04

, r . . . . .

BUS 2(SH1)

1- X20

21

RS422 30

RECEIVER 31

+28 V DC POWER

POWER RETURN

CHASSIS GND

SIGNAL GND

[H)(L)}

DME PRIMARY

(H)

}NAV PRIMARY

(L)

(H)

}DME SECONDARY

(L)

24

25

9

16

37

38

39

40

41

42

43

44

45

46

47

49

53

54

83

84

85

66

87

86

89

90

91

92

93

94

95

S6

97

96

S9

10(

10

10!

[H)}

GENERAL BUS NO. 3[L)

TAG SYNC

CDU SYNC

TRUE/MAG SELECT

DNSIDE TUNING CNTL

REMOTE TUNING CNTL

LAT WPT ALERT

VERT WPT ALERT

DEAD RECKONING

OFF SET ALERT

APPR SENSITIVITY

INDEP OP

CDU MSG

DEGRADE ACCURACY

NAV COMP VALID

VERTICAL TRACK AURAL ALERT

CROSS-SIDE TUNING CONTROL

FROMBus 1(SH 1)

4OVERSPEEDPROTECTION

RS422OFFSIDEVOR CONNECTEDNAV/DMEMANUALTUNESECNAV/DMEMANUALTUNEPRI

DISCCNTLINPUT-NOCLOCKASCEHIGHILOWSPEEDBUS-LTSNO.1HIGH/LOWSPEEDBUS-LTSNO.2HIGH/LOWSPEEDBUS-LTSNO.3

TAG SYNCCDU SYNC

LTSNO.1 NUMBER BITNO.1LTS NO. 1 NUMBER BITNO.2LTS NO. 2 NUMBER BIT NO. 1

LTS NO. 2 NUMBER BIT NO. 2



SDI NO. 3

CROSS FILL ENABLE

VER B ASCB

FUEL FLOW CONFIG IDO

FUEL FLOW CONFIG ID1

FUEL FLOW CONFIG ID2

OPERATIONAL MODE IDO

wow

PERF COMP INSTALLEI

LTS NO. 1 CONFI(

LTS NO. 1 CONFI(

LTS NO. 1 CONFK

LTS NO. 2 CONFI(

LTS NO. 2 CONFI(

LTS NO. 2 CONFK

-N OUTPUTDISCRETES

105I06

34

48

59

60

61

62

DL CONNECTED

RADIO CONFIG IDO

RADIO CONFIG ID1

RADIO CONFIG ID2

MAINT TEST ENABLE

lLS/MLS SELECT

LTS NO. 3 CONFIG

LTS NO. 3 CONFIG

LTS NO. 3 CONFIG

OPERATIONAL MODE ID1

INITIATED XMIT

INITIATED REC

DME SCAN TYPE

RADIO BUS TYPE

SINGLE ASCB

SDI NO. 1 = LEFT

SDI NO. 2 = RIGHT

CDU VALID

TRUE REF SELECTED

AFIS ENABLE

AD-30257, SH2#

63

d INPUTDISCRETE:

64

65

66

67

66

69

70

71

72

73

74

75

76

77

7e7~

4<4

— J-—

.—— ——— ——— —— ——— ——— ——— ——— ——

NZ-920 Navigation ComputerBlock Diagram




22-14-00Page 198.127

‘Aug 15/91Use or disclosure of information on this page is subject to the restri~lons on the title page Of this document,

9. B. CD-81O Control Display Unit (See Figures 9-3 and 9-4, and Table 9-4. )

PHOTO PHOTOSENSOR ANNUNCIATORS SENSOR

/ / I

CRTDISPLAY~

LEFfLINESELECT\KEYS

SCRATCHPA~

MODEKEYS—

ALPHA-NuMERIcKEYS

Elm MLF

B 020” S3.ONMSLC J q\610.L ~

09 10Z 160/4920

DESTKSLC c1

El

HB 348 ‘ 4. ONMKSLC 0s t tZ

J!

a4560

%-(-’ ‘LTN

ARGIVr41-D ‘–’

II \

-“ d/ fm“fmmmi=mairiim rm m

RIGHTLINE

~SELECTKEYS

FUNCTION~KEYS—MODE KEYS

\~:e.R;::ss

AD-11942-RI

CD-81O Control Display UnitFigure 9-3

22-14-00Page 198.128


MAINTENANCE

Honeywell !!N$!h.


Length (from rear of bezel) ................ 10.00 in. (254.0 mm)Width ....................................... 5.75 in. (146.1 mm)Height ...................................... 7.50 in. (190.5 mm)

Weight ........................................... 12.7 lb (5.76 kg)

Power Requirements:

Primary ........................................... 28 V dc, 40 WLighting ................................................. 5Vac

Mating Connector:

J1 ............................................... MS3126F22-55SX

Mounting ...................................... Unit Screw Fasteners

CD-81O Control Display UnitLeading Particulars

Table 9-4

The CD-81O Control Display Unit (CDU) is the pilot interface with theFMS. The CRT displays relative flight information to the pilot. Thepilot enters alphanumeric data into the system via the full alphanumerickeyboard. This data appears in the scratchpad to be line selected to theappropriate position on the CRT display. The CDU function controls aredescribed in the following paragraphs:

9. B. (1) CRT Display - A color CRT is used to display data on the CDU. Thedisplays consist of 9 lines, each line containing 24 characters.The first 1 ine is a title 1 ine and the ninth 1ine is the scratchpad.The intermediate lines and scratchpad are available for data displayand entry.

(2) Photosensors and Brightness Control - CRT brightness control isprovided in order to maintain readability under dim light as well asdirect sunlight. This is accomplished in two ways:

(a) Manually by the brightness knob - The brightness knob isprovided to manually vary the intensity of the CRT display fora given ambient light level.

(b) Automatically by the photosensors - The photosensors sense theambient light and adjust the CRT brightness automatically tomaintain the relative brightness set by the brightness knob.

22-14-00Page 198.129

Aug 15/91Useor disclosure of information on this page is subject to the restrictions on the title page of this document.

MAINTENANCE

Honeywell !MEJ+AAW9. B. (3) Annunciators - The six annunciators located at the top of the CDU

keyboard panel operate independently from the CRT and keyboard.Lighting of the annunciators is initiated by the navigation computeror performance computer via the RS422 serial data link. The twocolors used for annunciations are white and amber. White indicatesan advisory type annunciation, and amber indicates an alerting typeannunciation. The following paragraphs describe each annunciator:

(a) Display (DSPLY) Annunciator - The DSPLY annunciator is anadvisory type (white) that lights when the CDU is displaying apage that is not relative to the current aircraft lateral orvertical flightpath. The DSPLY annunciator will light under

the following conditions:

● When displaying a flight plan page other than page 1.

. When displaying a stored flight plan page.

. When displaying any of the review pages for SIDS and STARS.

● When displaying the ‘CHANGE ACTIVE LEG’ message.

. When defining the ‘Intercept’ waypoint on the active leg.

(b) Dead Reckoning (DR) Annunciator - The DR annunciator is analerting type (amber) that lights when the FMS is navigatingvia the DR mode which is defined to be the loss of radioupdating and the loss of all position sensors. The DRannunciator will light under the following conditions:

● When the FMS has been operating in the DR mode for longerthan 3 minutes.

. When the APRCH annunciator is illuminated and positionupdating from all sources is lost for more than 30 secondsand the radio is not procedure tuned.

s When the APRCH annunciator is illuminated and positionupdating from all sources is lost for more than 5 secondsand the radio is procedure tuned.

(c) Degraded (DGRAD) Annunciator - The DGRAD annunciator is anadvisory type (white) that lights when the FMS has entered adegraded navigation mode. The definition of degraded is whenthe FMS cannot guarantee the required accuracy for the presentsegment of flight. The DGRAD annunciator will light under thefollowing conditions:

22-14-00Page 198.130

Aug 15/91Use or disclosure of reformation on this page is subject to the restrictions on the title page of this document,

● When the aircraft is within the terminal area, 50.8 NM, andthe required accuracy cannot be guaranteed with theavailable position sensors.

● When the aircraft is within the approach area, 10 NM, andthe required accuracy cannot be guaranteed with theavailable position sensors.

I● Any time the FMS loses all sensors except the GPS. Present

use of GPS requires it to be compared with other approvedsensors. lf this cannot be done, the DGRAO annunciator isturned on.

. If the DEGRAD annunciator is on when the DR annunciator is

turned on, the DEGRAD annunciator will be turned off.

9. B. (3) (d) Message (MSG) Annunciator - The MSG annunciator is an advisorytype (white) that lights when the FMS is displaying a messagein the scratchpad to the flightcrew. The annunciator shallextinguish after the message(s) have been cleared from thescratchpad.

(e) OFFSET Annunciator - The OFFSET annunciator is an advisory type(white) that lights when a laterally offset path has beenentered into the FMS using the progress page. The annunciatorturns off when the offset has been removed. If there is anoffset when the APRCH annunciator is lighted, the offset willbe removed and the annunciator turned off.

(f) Approach (APRCH) Annunciator - The APRCH annunciator is anadvisory type (white) that lights when in approach mode. TheNZ-920 Nav Computer output sensitivity of lateral deviation tothe PFD will be ramped to a higher sensitivity when theannunciator is lighted. The APRCH annunciator will light underthe following conditions:

. If the destination elevation is specified, distance todestination is less than 15 NM, altitude is less than 2500feet above the destination elevation, and the speed is lessthan 200 knots.

. If the destination elevation is not specified, distance todestination is less than 15 NM, and speed is less than 200knots.

● If the APRCH annunciator is illumflies out of the above conditionswill be extinguished and the sens

● When the pilot selects APP on thecontroller.

nated and the aircraftthe APRCH annunciatortivity ramped to normal.

flight guidance

22-14-00Page 198:131


Honeywell !!!!!!!!f~wc’9. B. (4) Line Select Keys - There are four line select keys on each side of

the CRT display. For reference, the keys are defined one throughfour from top to bottom on either side of the CRT. A left and rightis also assigned to define the side of the CRT on which the key islocated. For example, line select key ‘IL’ is the top left lineselect key.

In the case of an index display, the line select keys can be used toselect submodes within the major modes. In displays other thanindex, the line select keys 4L and 4R are primarily used for directaccess to other modes in the FMS. Data can be copied to thescratchpad through the use of a line select key, and once data hasbeen entered into the scratchpad, either via line selection ormanual keyboard entry, it may be selected to any of the allowableline select fields on a given page. This can be accomplished simplyby depressing the key adjacent to the line in which you wish thescratchpad data to be inserted.

(5) Function Keys - There are four function keys, and the function ofeach is described in the following paragraphs:

(a) Previous (PREV) and NEXT Page Keys - The number of pages in aparticular mode or menu display are shown in the upper right-hand corner of the display. The format is ‘AA/BB.’ ‘AA’signifies the number of the current page that is displayed.‘BB’ signifies the total number of pages that are available forpilot viewing/modification. Page changes shall be done byselecting the PREV and NEXT keys. When in the PLAN mode, thesekeys will increment or decrement the map center waypoint.

(b) Clear (CLR) Key - The CLR key performs the following functions:

. When a message is present in the scratchpad, pressing theCLR key shall delete that message.

. When an alphanumeric entry resides in the scratchpad, onecharacter shall be cleared from the scratchpad (from rightto left) for each time the button is depressed.

c When an alphanumeric entry resides in the scratchpad and theCLR key is held down, the first character is cleared within100 ms. After 400 ms have elapsed, characters will becleared at 100-ms intervals for as long as the key is helddown.

(c) Delete (DEL) Key - When there is no message in the scratchpadand the DEL key is depressed, a “*DELETE*” will appear in thescratchpad. This may now be line selected to delete waypointsand other items displayed on the CDU. When there is a messagedisplayed, the delete operation will be inhibited.

22-14-00Page 198=132

Aug 15/91Use or disclosure of information cm thts page is subject to the restrictions on the title page of this document,

MAINTENANCE

Honeywell !!!t!!!+AAiv9. B. (6) Mode Keys - There are five mode keys, and

described in the following paragraphs:the function of each is

(a) Performance (PERF) Mode Key - Pressing the PERF mode key shallenable the pilot to access the performance page(s). The pilotmay select or enter the applicable information through the useof the line select keys.

(b) Navigation (NAV) Mode Key - Pressing the NAV function key shallenable the pilot to access the NAV index page. The pilot mayselect any of the submodes by depressing the line select key.

(c) Flight Plan (FPL) Mode Key - Pressing the FPL shall display thefirst page of the flight plan. If there is no flight plancurrently entered, the pilot may manually enter a flight plan,select a stored flight plan, or create a stored flight plan.

(d) Progress (PROG) Mode Key - Pressing the PROG key shall displaythe first page of the progress pages. The purpose of this modeis to show the current status of the flight. This firstprogress page shall display the ‘to’ waypoint, the destination,the navaids that are currently tuned for radio updating, andthe update status of each navigation computer.

(e) Direct To/Intercept (DIR) Mode Key - Pressing the DIR mode keyshall display the active flight plan with the DIRECT andINTERCEPT prompts.

(7) Alphanumeric Keys - The control display unit provides a fullalphanumeric keyboard to enable pilot inputs to the scratchpad ofthe CDU. A key is provided for each letter of the alphabet as wellas each number, decimal, minus sign and slash.

(8) Scratchpad - The scratchpad is used for data entry, messages, anddata transfer using the line select keys. The scratchpad isindependent of the page being displayed. Any item in the scratchpad

remains when selecting other pages. Pilot entries to the scratchpadare made using the alphanumeric keys. If the scratchpad is empty,line select keys can be used to copy many items from the respectiveline to the scratchpad. Messages have display priority in thescratchpad. Entries existing or made in the scratchpad when amessage is displayed remain for display when the message is cleared.The editing mode of the scratchpad is activated by ending the entrywith a “-”. In the editing mode, the “PREV” and “NEXT” keys move acursor, displayed in the inverse video, in the scratchpad. Thecharacter in the inverse video field can be removed with the “CLR”key or a new character inserted by entry. “DEL” will delete theentire scratchpad entry when in the edit mode.

22-14-00Page 198.133/198.134

Aug 15/91Use or disclosure Of Information on this page is subpct to the restrictions on the title page of this document.

NAVCOMPUTER

PERFORMANCECOMPUTER

5VLIGHTING JKEYBOARD

LIGHTING COMMON D< 1 DLIGHTING

DATABUSOUT

OATA TERMINALREADY BUS OUT

{

M

N

{

P

RmI (H)

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RS422(H) OUT

(c)

{

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DATA BUS INT

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(

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INPUT

(

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(c)

(

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CLEAR TO SEND BUS INRS422

I(c) INPUT

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CLK BUS INh

(c)

DATABUSOUT

DATA TERMINALREADYBUSOUT

BRIGHT/DIM DISCRETE

TEST ENABLE

28VDCPOWER

{

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LAMP TEST

28VANNPOWER

{

ANN INTENSITY CONTROL

CHASSISGND

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A

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+

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~VIDEO

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I I

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CD-81O Control Display UnitBlock DiagramFigure 9-4 22-14-00

Paae 198.135/198.136

R PHOTO SENSOR OUT

L PHOTO SENSOR OUT

CDU VALID

AD-11970.RI

Aug 15/91Use or disclosure 01 mformallon on this page E subject to the restrictions on the btle page of this document

MAINTENANCE

Honeywell !M%thl.


22-14-00Page 198.137

Aug 15/91Useor disclosure Of reformation on this page IS subject to the restrictions on the title page of this document.

9. c. DL-800 or DL-900 Data Loader (See Figures 9-5, 9-6 and 9-7, and Tables9-5 and 9-6.)

DL-800 Data LoaderFigure 9-5

I


Length ...................................... 8.30 in. (210.82 mm)

Width ....................................... 5.75 in. (146.05 mm)

Height . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4.50 in. (114.30 mm)

Weight (maximum) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5.6 lb (2.54 kg)

Power Requirements:

Primary ........................................ 28Vdc, 22Wmax

Mating Connector:

J1 ................................................. MS3126F16-26S

Mounting ........................................ Unit Dzus Fasteners

DL-800 Data LoaderLeading Particulars

Table 9-5

22-14-00Page 198.138


Honeywell !&!fr.c’

AO-29793

DL-900 Data Loader(Access Door in Open Position)

Figure 9-6


Length . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ...”-” 8“OOin” (203s20Width . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ...””” 5.75 in. (146.05

Height ........................................ 2.24 in. (56.9

Weight (maximum) ................................... s“Olb(l.sG

Power Requirements:

Primary ....... ........... ......... ............”. zBVdct ‘w

mm)mm)mm)

kg)

max

Mating Connector:

J1 ................................................. MS3126F16-26S

Mounting ........................................ Unit Dzus Fasteners

DL-900 Data LoaderLeading Particulars

Table 9-6

22-14-00Page 198.139

Apr 15/93


The DL-800\900 Data Loader is used to transfer navigation-related data tothe NAV computer. The DL-800/900 has the capacity ~o transfer in excessof 600k bytes stored on a 3-1/2-inch minidiskette. It has an RS-422interface with the NAV computer. The DL-800/900 is also used to downloadtrend and limit exceedance data from the FC-880 Fault Warning Computer(FWC) through an RS-232 interface with the FWC.

9. c. (1) Navigation Database Loading - The DL-800/900 Data Loader providestransfer of data derived from the Jeppesen database from a 3-l/2-inch floppy disk to the NAV computer local EEPROM memory. This dataincludes navaids, waypoints, airports, airport runways, airportprocedures, and jet routes from the Jeppesen data source. Thedatabase is updated every 28 days. The data transfer rate is 312kbaud. The total time required to load a full database isapproximately 8 minutes. The LEFT/RIGHT/AUX switch on the front ofthe data loader is used to select which RS-422 port will be used tooutput data.

(2) Flight Plan Loading - The DL-800/900 Data Loader also has thecapability of interfacing with a ground-based Lockheed JetPlanComputer or equivalent. It is capable of transferring an optimizedflight plan from the ground-based computer to the navigationcomputer via a 3-1/2-inch floppy disk.

(3) Trend and Limit Exceedance - The DL-800/900 Data Loader interfaceswith the FWC through an RS-232 bus. This single RS-232 bus is relayswitched between the two FWCS. Control of this relay is by a switchlocated in the right avionics bay. Data download is independent ofthe LEFT/RIGHT/AUX switch.

(4) AFIS Flight Plan Loading - The DL-800/900 Data Loader also has thecapability of transferring a flight plan in AFIS format to thenavigation computer via a 3-1/2 inch floppy disk.

22-14-00Page 198~140


MAINTENANCE

Honeywell W!+fb

“ (H)RS232 DATA IN a <

RS232 SIG GNLI V< !(c) (H) ‘1

n > Z RS232 DATA OUT-m

●CENTRAL PROCESSING UNIT (CPU) I

{:: ~

(c) -280 CPUDATA BUS IN G

I-8 KBYTESRAM I

FROM NAVCOMPUTER H (H) -32 K BYTES EPROM

- RS422 INTERFACE(c) I ~

- RS232 INTERFACE I:}

DATA BUS OUT

{~

(c) (H) T TO NAV COMPUTERCLK BUS IN JFROM NAVCOMPUTER K (H)

ICHASSIS GND b I

P

I

{:;

B(L)

28 V DC POWER(H)

c I (R)● POWER SUPPLY CARD

{ :,

RIGHT D -5VDC

LOADER-+12VDC

INSTALLED (L) -DISCRETES FOR CPU

DISCRETE FROMLE17 E

NAV COMPUTERAUX F

(A)

El-EQ- AD-l19694?2@

DL-800/900 Data LoaderBlock DiagramFigure 9-7

22-14-00Page 198.141

Aug 15/91Use ordisclosureofinformationon thispageissubjecftotherestrictionson thetitlepageofthisdocument.

Honeywell f&!!![r.cE9. D. PZ-800 Performance Computer (See Figures 9-8 and 9-9, and Tables 9-7 and

9-8. )

AD-1 1S41

PZ-800 Performance ComputerFigure 9-8

22-14-00Page 198.142

Aug 15/91Use ordisclosureofinformationOn thispagek subjecttotherestrictionsOn thetinepageofthisdocument.

MAINTENANCE

Honeywell !!WA&.


Length ...................................... 17.03 in. (432.6 mm)

Width ..... ... ....... ...... ..... .. ... ... . .. .. . 4.91 in. (124.7 mm)

Height .... .......... ..... . ..... .. ... .... .. .. . 7.62 in. (193.5 mm)


Power Requirements ......... .. ..... ..... ... . .. ... .... .. 28 V dc, 65 W

Mating Connector:

J1 ..................... Cannon Part No. DPX2MA-106P- 106 P-33 B-OOO1


PZ-800 Performance ComputerLeading Particulars

Table 9-7

The purpose of the PZ-800 Performance Computer is to provide the pilotwith detailed performance information and automatic control of thethrottles. Previously, detailed performance information was onlyavailable by spending a great deal of time in the aircraft performancemanuals. With the performance functions of the PZ-800, the pilot now hasaccess to a computerized performance manual. The pilot is supplied withflight-planning information prior to takeoff such as fuel and timerequired. In flight, the system provides the pilot with real-timeinformation based on current aircraft and atmospheric conditions. Incase of changes to the flight plan, the pilot can be updated immediatelyon the consequences of the changes. This capability extends the pilotsability to conduct the flight safely and economically.

The performance system is enhanced by the autothrottle. Aircraftparameters calculated by the performance functions can be coupled to theautothrottle and flight guidance computer for effortless tracking ofthese commands. Alternatively, the autothrottles can be set to maintainpilot-entered values. In addition to the PZ-800 Performance Computer,the autothrottle function requires one SM-81O Autothrottle Servo-ClutchAssembly for each engine controlled. Inputs from the throttles, engines,flight guidance controller, and display controller are also required.

The PZ-800 Performance Computer digital outputs transmitted over the ASCBare listed in Table 9-8.

22-14-00Page 198.143

Aug 15/91Use or disclosure of information on this page is subpscf to the restrictions on the title page of this document.

NN

WI1!%H

-0N/0o0

USP BIT BIT FUNCTION NOTE FORMAT

01

SCALE RSB APPROXRESOL/LSBPOS SENSEI... ...... ........................-l----i;;-;~;-------------------l--------------------l---l-----------------l-------------=---l::!!l:!::!:=-IHOLC FLAG

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65A0

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2 1S-0 MAINTENANCETESTUNEN OAT HCADER= OATA O, OTHERUISE● PARE 6692

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3433323

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---i . CAPTURED1. ~MEo1. CAPTUREO1 . ARMEO1 = CAPTUREO1 . HoLD] ■ OVERRIDE1. ENGAGEo1 . L~MITJNG

31 hiiOTftROtfLEARM SM1lCM 1. ~~o30 TRANSITIONII

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! :: W40 LII!ITEO I . vALID 65A6MO LIMITED 1. vALIo

4 13 POHERLIMITEO 1 = VALID4 12 WEPLACARO LIMITEO 1. VALIO4 11-94B LANDINGGEAR LIHITEO 1 . vAL~D41 ENGAGEINHIBIT 1 = INHIBITi 6-o SPARE

61I II--------------------------------------*-----------------------------------------------.----------------I Ii i-----”----”------1;-;1--”-----1CHECKSUM SUNOF UDRDS

65A8

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NN

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FTIU SIMVAR--- ------ II--------------------... .................I-----------------l----l--------I

I--;1 I II I------------------------------------------------------------...................-P14C-PCONTROL PACKEDLOGIC

l---l-----------------l-----------------i~i~;l--------i

1 15 TEST LOGIC 1 ■ TEST 61C01 14 VAL101 13-12

LOGIC 1 = VALIDSPARE

1 11 HODE/ENGAGEPRIORITY LOGIC 1 ■ HIGH (SELECTED)co~~;S~CTEDPZ)

1 10-8 O-7 HEX1 ?-o PM-P ADDRESSII

18 H - LEFT, lC H - RIGHT--- ------ 11......................... ---- -------------------------I 1151100,5,0 ~Bs IAL”AYSPOSIT*VE 16,921 I....................---.................-.....-..-..-*--------.....-.2 15-1 GR~;~l~EIGHT THO’SCONPLEllENT +1-163840LBS20

●

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REDUCEDTHRUSTTYPE0:000078125

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.

83 CONFIGURATIONNISNATCH ~ . HIS~TCH61CE

82-1 CURRENTSPEED 00= V1;01=V2;

I :010. VR i 11. SpARE

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USP BIT 611 FUNCTION NOTE FORMAT SCALE RSII APPROXRESOL/LSO POS SENSE FTIU SIMVAR

Q,

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A+/- 512 KNOTS . .

94 SPEEOOISPL Y 1 - OISPLAYVref 61000 = DISPLAYVspeed

93 SPARE92-1 CURRENTSPEED DO= VSe; Ol=Vfs;

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10 15-5 EPR LIMIT/CHD0:000078125 6102

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11 15-0

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13 15-0 SPARE

I l-------"-----"---""------l----lfi~;-;~~~i~~~~---------l--- ------6108

;;-”;;-;;:---------l-;-;;-l;;~--;;-----l-----------------l;i;~l--------I14 15-9 BANKANGLEBUFFETLIHIT .:: ~1

. .SPARE 61DAVALID 1 - vAL[o

II... ...... .........................l----l-------------------------l II...................-....................I15 15-0 SPARE

--------”--------1 ;;;;1--------1

II -------------------------l----lfi;;;-;;fii~~;~-------l--------------------lI61OC

. . . . . . . . .16 16-4 CA:~~CH TARGET

I--- . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . l;i~;l--------l16 +/. 512 KNoTS 12 0. 25/0.015625 FORMARLI 610E16 MACH +/. 2.0 MACH 12 0.00098/0.000061 FONWARD163-2 SELECT 00. CL~MO

01 = CRUISE10 = DES11 - SPARE

16 1 CAS/HACH i-= HACH, O ● CAS16 D VALID

II1 - VALIO

--- ------ -------------------------1 I I----......................... ....................1---1-----------------117 CNECKSUN

-----------------l----l--------ISUMOF UOROS

II----------------------------------l----l;;;----------------------l--------------------l---l-----------------l -----------------1;;;;1--------118 ERROR CHECK

6iE0I l;i;-;i~:---------------l----l;~-;;~-------------------l--------------------l---l-----------------l-----------------l----l---... ......

19II-........ .........................l----l-------------------------l--------------------l---l-----------------l-----------------l----1--------1_. —-— .— .

MSP BIT BIT FUNCTION NOTE FORNAT SCALE RSB APPROXRESOL/LSBPOS SENSEII II

FTIU SIMVAR-.. ...... .............---------------- -------------------------I--------------------l---1 I................. .............----0 NOLCFLAG 7E HEX

1----1--------1

II... ------.........................II I II---- .-...-.-.-?------------------------------------- -----------------..---------------1 1----1--------1. ..‘PNC-PCONTROL ‘PACKEOLOGIC

: 15 TEST LOGIC 1 ● TEST1 14 VALIO LOGIC 1 ● VALID1 13-11 SPARE1 10-8 COUNTER 0- 7 NEX1 7-o Pl&P AODRESS 1A H - LEFT lE H - RIGHT

2“15-6 ‘SPARE. .

26-0 MAYPOINTNUHBER BINARY(INTEGER) 0-121I l--------------------"---"l----l~&;~-;;fi~;~~~~--"----l;;:-;;;;-;~:"-------.. ......

3 15-0 ALTITUOE

I I..-..................................1----1--------17 1 MAYPOINTI I Ill.-.................. .................----........

16 .2.0 FT. .UPII-........;fi~;;~;;;[;;-----”-llwo,~CO”PLEWNT I-------------------------------------------------l---l-----------------I-----------------l----l--------I4 15-5 +/-1023KNOTS 11 1.0/0.0312 KNOTS ALHAYSPOSITIVEi i:o- SPAiE-

11-.. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1----1 I. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ------ 11011,.,0,01562 IAL”AYSPOSITIVE I I I5 15-6 OILIBRATiOAIRSPEED

--- . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ---- . . . . . . . .THO’SCO14PLEHEN1 +/- 511 KNOTS

55-0 SPAREII II-.. . . . . . . . . . . . . . . . . . . . ------------ -.. d ------------------------- . . . . . . . . . . . . . . . . . . . . --- . . . . . . . . . . . . . . ..- . . . . . . . . . . . . . . . . .I 111132,10FT,“lN I

6 15-5 VEN#L SPEEO TUO’SCOllPLEtlENT +/-327066FTo/HIN .*. UP64-0

I l-------------------------l----lT”o,~CMPLEMNT-.. ...... 1......................... ....................l;;l;;;:;i;;;------l;;;;;~;----------7 15-6 UINOSPEEO +/- 511 KNOTS7 5-o SPARE

II..- . . ---- . . . . . . . . . . . . . . . . . . . . . . . . . 1----1 I I,. 10035,,0* 00549 i~;f------------”------------------------- -------------------- --- . . . . . . . . . . . . . . . . .815-6 Ul~OD~RECTION THO’ S COMPLEMENT +/- 180 OEGREESa 5-0---11-”- 11-.................................---------------------------------------------------------------------I I~ 106,0001953 l~:”;--;--------

9 15-8 OELTAISATEMPERATURE TWO’SCO14PLEHENT +/-63.5DEGC9 7-o SPARE

● .

II I lmo,~CWPLEHENT-..------......................... -----.*-------------------------------------------...................1 116120 LBS1015-0 FUEL REMAINING

I.................+1-65534LBS. ● ALUAYSPOSITIVE

II------------

11-----.-.....

II..-.........

II-----......-

1-----.-.....I

1i~15-o ~NExTUAYPOINTDATA (11-19I--- ------........................- ---- ---------------------------------------------I l---l-----------------l Ill.........------------........

I--------------------------------------.........................I II...................----.................I20 15-0 NEXTNAYPOINTOATA (20-28

-----------------l----l--------I-.. ------------------------------- I I---- ---------------------------------------------1---1 I29115-0 1NEXTUAYPOINTDATA (29-37

................. .................1----1--------[

I... ...... ......................... ---- ------------------------- ...................-I l---l-----------------l38115-0 1NEXTUAYPOINTDATA (38-46 -----------------l----l--------1

II-.. ------ 1......................... ---- -------------------------147 15-0 SPARE

--------------------l---l-----------------1-----------------l----l--------1

i l-------------------------lI... ------ ---- ........................-4815-0 SPARE

I--------------------l---l-----------------1-----------------l----l--------1

II-.. ...... .........................49

l--”-l&&-~;;&-----””-----l-------------”------l---1 ICHECKSUM................. .................i----l--------1

II--- ------....................*..-.I----lcRc I......................... ....................i---l-----------------150 ERRORCHECK -----------------l----1--------III II---...............................----.........................

51I

HOLCFLAG--------------------l---l--------:--------l-----------------l----l--------l

II7ENEX

-..------....................-----11----.....................-..”I--------------------l---l-----------------l-----------------l----1--------1

FLT TESTONLY

CDUINTERFACE

CHASSIS QN[

CHASSIS GND

SIGNAL GND

SIGNAL GND

SERVO POWER (H)

sERvo POWER (H)

CLUTCH POWER (H)

CLUTCH PowER (H)

COMPUTER POWER (H)

COMPUTER POWER (H)

SERVO POWER (L)

SERVO POWER (L)

CLUTCH POWER (L)

CLUTCH POWER (L)

COMPUTER POWER (L)

COMPUTER POWER (L)

ASCB PRIMARY PORT{

(H)(L)

ASCB SECONDARYPORT{

(H)(L)

RS232 RCVRRS232 RTN

RS232 XMTR

RS232 RCVRRS232 RTN

RS232 XMTR

{RS422 XMTR (F71 GLK) U

RS422 xMTR (ml CNTL){

(H)(L)

RS422 XMTR (F71 DATA){

(H)(L)

RS422 RCVR (Ffl DATA){

(H)(L)

{(H)

‘W22 ‘CVR ‘H’ CNTL) (L)

RS422 XMTR (CDUCLK){

(H)

(L)

{

(H)

‘s422 ‘MTR ‘CDU CNTL) (L)

1S422 XMTR (CDU DATA){

(H)

(L)

RS422 RCVR (CDU OATA){

(H)

(L)

1RS422 RCVR (COU CNTL){

(H)

(L)

RS232 RCVR

RS232 RTN

RS232 XMTF!

1A

1

2

3

4

5

6

7

8

9

10

11

12

13

14

33

M

39

!0

41

45

46

47

B4

05

e6

B7

3s

09

90

91

92

)3

17

)6

)s

Ml

01

02

03

M

05

M

12

43

b4

1=1

2

E

POWERSUPPLY

~

4

5 d ASCBCPU

lB~— ——— — ‘————— ‘————— ‘————— 1 JIA JIE

23

24

48

49

50

54

59

~ 60

~ 62

~ 63

~ 6s

MONITORS 72

73

74

75

m

67

M69

ao

91

92/ xx

Ell R: “’’”( ‘N”mI‘L--J’JJ

\

PERFORMANCE \CPU I I

~

IAUTOTHROITLE

CPU

-L—————— ——————— —

Performance ComputerBlock DiagramFigure 9-9

22-14-00

17

18

IIIIIII

—

———— ———— J

SERVO CLUTCH ORIVE NO. 1

SERVO CLUTCH ORIVE NO. 2

CROSS SIOE NT ENGAGE

AA ENGAGE

MAINT TEST ENABLE

GEAR DOWN

LEFf/RIGHT SELECT

ASCS VEFl NB

LEFf BLEEO SRC ON

RIGHT BLEED SRC ON

ASCB SINGLE/OUAL

FLAPS IN MOTION

LE17 AC PACK ON/OFF

RIGHT AC PACK ON/OFF

wow

NT ENGAGEIOISENGAGE

INTERLOCK NO. 1

INTERLOCK NO, 2

INTERLOCK NO. 3

INTERLOCK NO. 4

MT DISCONNECT

AIT ENGAGE GNO

PERF COMP INSTALLEO

(c) PLA 1 Pos(c) PLA 2 Pos

(L)

}SERVO NO. 1 TACH

(H)

(L)

(H) }SERVO NO. 2 TACH

(L)

}PLA 1 Pos

(H)

(L)

}PLA 2 Pos

(H)

(L)

}PLA REF

(H)

(H))

SERVO NO. 1 ORIVE(L)

(H)

(L) }SERVO NO. 2 DRIVE

AO-30256#

Page 198.149/198.150Aug 15/91

UseordisclosureofInlormellonon lhl~page IS subject10theresfrldlorlson the title page Of this document.


22-14-00Page 198.151


MAINTENANCE

Honeywell ?$N!!+A&.9. E. SM-81O Servo-Autothrottle (See Figures 9-10 and 9-11, and Table 9-9.)

The SM-81O Servo translates electrical inputs into a clutched mechanicalrotational output to automatically control the throttle setting. Aspline output on the clutch drives a cable drum. The servo is connected

to the throttle control rigging by the cable drum drive.

#12765

SM-81O Servo (Autothrottle)Figure 9-10

22-14-00Page 198;152


MAINTENANCE

Honeywell W!$$h.


Length .................................... 11.10 in. (281.94 mm)Width ...................................... 4.05 in. (102.79 mm)Height ...................................... 3.90 in. (99.06 mm)


Power Requirements:

Clutch Excitation ............................. 28 V dc, 24Wmax

Stall Torque ................................... 39.9 to 47.8 lb-in.

Mating Connector:

J1 ............................................ MS3106A-16S-1S(C)

Mounting ................................................ Hard Mount

SM-81O ServoLeading Particulars

Table 9-9

rP

MOTOR INPUTPOLAFilTY CAUSES {Ccw SERVO

OUTPUT ROTATION,‘ND ‘h’ ‘1”

AS VIEWED FROM I

OUTPUT END I

I

JI+B BLK 2 L2 ~ I R2 12

1=~ ‘~:= ‘“ ~:: ~“~ ‘“J

ELK

III 50

R4 - - C3SEL

‘ ‘1

III

IIIN5614 TM” ‘;! ‘B II

-c 4 1, I

I Ss IwHT If% 1[

9. L3 6- -10 RED II

1

SS2

IIGRAY

(-+E< iYEL S BLK1

/+ h

II

{ :!i BRN f>

13+A

f\ WHT-BLK \

TACH 14 iiIr I

-DWHT-RED II

YSS5 VSS3<)

/$#RYk

&’RY

SM-81O Servo SchematicFigure 9-11

AD-12766

22-14-00Page 198.153

Aug 15/91Use or disclosure Of information on this page is subject to the restrictions on the trtle page of this document.

MAINTENANCE

Honeywell !W%h.10. Enqine Pressure Ratio (EPR) System

Engine Pressure Ratio Transmitter (EPRT) (See Figures 10-1 and 10-2, andTables 10-1 and 10-2. )

CAUTION: THE EPRT CAN BE DAMAGED BY AMBIENT TEMPERATURES OF GREATER THAN+125 “C (+257 “F) OR BY INPUT PRESSURE OF GREATER THAN101.80 INHG.

Engine Pressure Ratio TransmitterFigure 10-1

Dimensions:

Height ........................................ 2.35 in. (59.7 MM)

Width . . .. ... ... .. ... . .. .... .... ....... . ...... 5.48 in. (139.2 mm)Length ....................................... 7.82 in. (198.6 mm)

Weight ............................................... 3.1 lb. (1.4 kg)

Power Requirements ............................. 115 V, 400 Hz, 7 watts

Mating Connector ....................................... MS27404T16B35S

Mounting ................................................... Hard Mount

Engine Pressure Ratio TransmitterLeading Particulars

Table 10-1

22-14-00Page 198.154


10. A. An aircraft engine pressure ratio (EPR) system contains two solid-state,microprocessor-controlled EPRT, one on the captain’s side of theaircraft and one on the first officer’s side. Each EPRT receives fanduct pressure (FDP) from its associated (on-side) aircraft engine andtwo Aeronautical Radio Incorporated (ARINC) 429 data signals. BothARINC 429 data signals provide data for total pressure (P~) andcalibrated airspeed (CAS). CAS data is used for built-in test equipment(BITE) only and does not affect calculation of EPR. One ARINC 429 data

signal is input from the aircraft on-side air data computer (ADC), and

one is input from the aircraft off-side ADC. During normal operation,

data from the on-side ADC is used. If the data from the on-side ADC isnot valid, data from the off-side ADC is used.

B. The two EPRT calculate the ratio of FDP to PT. Discrete trimplugsignals (TRIMPLUG O through TRIMPLUG 5) adjust this ratio. Two EPRTtransmit the adjusted ratio on the ARINC 429 bus.

c. The EPRT has three modes of operation:

● Initialization mode.

● Normal operation mode.

c Memory access mode.

D. The EPRT enters the initialization mode after power-up, or after a powertransient of longer than 50 ms, or after watchdog timer (WDT) time out.The initialization mode lasts 750 ms maximum. During the initializationmode, the EPRT checks internal performance BITE, reads the discreteinput signals (TRIMPLUG O through TRIt4PLUG 5, SDI 1, and SDI 2), andinitializes software. When the initialization mode is complete, theEPRT enters normal operation mode.

(1) The discrete trimplug signals (TRIMPLUG O through TRIMPLUG 5)provide aircraft engine compensation values used during EPRcalculation.

(2) The discrete source/destination identifier (SDI) signals (SDI 1 andSDI 2) identify the location of the EPRT on the aircraft.

E. The EPRT enters the normal operation mode when the initialization modeis complete. In the normal operation mode, the EPRT:

. Receives P~ and CAS data from the aircraft ADC.

. Computes PT.

. Measures FDP from its associated aircraft engine.

. Outputs the calculated FDP on ARINC 429 bus.

22-14-00Page 198.155


MAINTENANCE

Honeywell %$i!+k.. Calculates EPR using a ratio of FDP to PT. If valid P~ is not

available from the on-side ADC, the EPRT uses the P~ data from theoff side ADC.

● Adjusts the calculated EPR according to the discrete trimplug signalinputs so that the aircraft engines are tuned for the same thrust.

. Transmits the calculated EPR on ARINC 429 bus to aircraft EPRindicator.

c Monitors internal EPRT performance with BITE.

. Outputs BITE results (maintenance data) on ARINC 429 bus.

10. F. The memory access mode is used during testing and calibration only.

22-14-00Page 198;156

Aug 15/91Use or disclosure of information on this page IS subject to the restrictions on the title page of this document.

pig

J1-16(H)J1-9 (L)

J1-53(H)J1-49(L)

J1-55(H)J1-44(L)

J1-11J1-47

J1-17J1-26J1-25J1-34J1-32J1-33

J1-19

MAINTENANCE

Honeywell W&!.Ak.Name

115 V ac InputPower

On-side DADCARINC 429 input

Cross-side DADCARINC 429 input

Left SDIRight SDI

TriTriTriTriTriTri

mplug O (LSB)mplug 1mplug 2mplug 3mplug 4 (MSB)mplug 5

Remarks

104 to 122 V ac, 360 to 440 Hz15 Watts (max)

The EPRT receives label 242 (totalpressure) and label 206 (airspeed)normally from the on-side DADC.The EPRT will use the cross-sideDADC when the on-side is invalid.

Left EPRT: J1-11 = open, J1-47 =gndRight EPRT: J1-11 = grid, J1-47= open

The trimplug discretes adjust theoutputed (trimmed) EPR by the followingequation:Trimmed EPR = ((EPR - 1) * MF) + 1Multiplication factor (MF) isdetermined by the following table:

(parity -- odd)Trimplug common

Pin 33100

:110011010010110100110010110

3200000000000000001111111111111111

NOTE:

3400

0

0

0

0

0

0

111111110000000011111111

o=

Engine Pressure Ratio TransmitterInput/Output Information

Table 10-2

2500001111000011110000111100001111

26001

:011001100110011001100110011

17010101010

:101010101010101010101

MF0.9750.9770.9780.9800.9820.9830.9850.9870.9880.9900.9920.9930.9950.9970.9981.0001.0021,0031.0051.0071.0081.0101.0121.0131.0151.0171.0181.0201.0221.0231.0251.027

shorted to J1-19open

22-14-00Page 198.157


~

J1-2 BITE

HoneywellName

Fault

J1-12 NonvolatileMemory Reset

J1-31(H) #1 EPR ARINC 429J1-39(L) output

J1-37(H) #2 EPRARINC 429J1-52(L) output


Remarks

5 V dc (1OK in series) = No recordedfaults. O V dc = Faults recorded.

Ground clears all fault data recordedin memory.

Both identical outputs transmit thefollowing data:Label 340 - engine pressure ratioLabel 350 - flight line fault codeLabel 352 - fan duct pressureLabel 353 - shop fault code

Engine Pressure Ratio TransmitterInput/Output information

Table 10-2 (cent)

22-14-00Page 198.158


CHASSIS GND ~

ON-SIDEDADC

{n”

II

ARINC 429II

‘?’”

CROSSIDEDADC

{n”

II

ARINC 429 ‘ I

c

lPARITY+

PROGRAM PINCOMMON —

SHIELD GND—

~

16

9

?4

53

49

4s

55

44

w

11

47

17

26

25

34

32

33

19

18

Honeywell !!!!!gb.c’

KEl=3 G

H

L

3 c

-1

FAN DUCTPRESSURE

J

II .

1- —————————..

J1

]31

39

WI

37

52

45

7

1

3

12

2

EPROUTPUTARINC 42S(12.8 KHZ)

MONITOR) INPUT

RS 232 1TESTONLY

NON-VOLATILE~ MEMORY

RESET

~ BITE FAULT

AD-213E6

Engine Pressure Ratio TransmitterBlock DiagramFigure 10-2

22-14-00Page 198.159

Aug 15/91Use or disclosure of Information on this page is subject to the restrictions on the title page of this document.

11. Optional VLF/Omeqa System

A. OZ-800 Receiver Processor Unit (RPU)Table 11-1.)

See Figures 11-1 and

AD-15616

11-2, and

OZ-800 ReceiverFigure

Processor Unit11-1

22-14-00Page 198.160



Length........................................ 14.58 in. (370.33 mm)Width ........................................... 2.25 in. (57.15 mm)Height ......................................... 7.60 in. (193.04 mm)

Weight (maximum). ...................................... 6.5 lb (2.95 kg)

Power Requirements.................................... 28 V dc, 40 Wmax

Frequency .............................................. 10.2 to 13.6 kHz

Mating Connector..................... Cannon Part No. DPXBMA-57-33S-OO01

Mounting........................................................... Tray

OZ-800 Receiver Processor UnitLeading Particulars

Table 11-1

The 0Z-800 RPU receives and processes data from the ground-based OMEGA/VLFstations to provide updated position and velocity information to the FMSnavigation computer. The RPU receives initialization data from the FMS. TheRPU receives the amplified antenna signals and converts them into positioninformation. The RPU also supplies the antenna with *12 volts dc power.

The RPU receives the following signals over an ARINC 429 low-speed bus:

Label Parameter Name

210 True Airspeed (TAS)314 Heading041, 042 Initialization of Set Position125 Greenwich Mean Time (GMT)260 Date272 Station Deselect

22-14-00Page 198.161


Honeywell f~~~f~~vc’

The RPU provides the following outputs over the 429computer:

Label

310311366367270

271

272

125260

bus to the FMS navigation

Parameter Name Min. Required Rate (1 Second)

LatitudeLongitudeN-S VelocityE-W VelocityDiscrete Word 1(Refer to Table 11-2)Discrete Word 2(Refer to Table 11-3)Discrete Word 3(Refer to Table 11-3)GMTDate

2

2

22

BIT(S) FUNCTION SET

1-89 -1315-14

1617

1819202122-29

LabelAll ZerosLAR Status (Lane Ambiguity Resolution)15 140 0 no LAR in progresso 1 LAR in progress1 0 insufficient signals for LAR1 1 unsuitable geometry for LAR

System is in dead reckoning (DR) modePosition uncertainty estimate exceeds4.0 NMSystem in nonsynchronizedSystem navigating in relative modeMinor RPU failureMajor RPU failurePosition uncertainty estimate in binaryformat, with Bit 22 corresponding to0.1 NM and Bit 29 corresponding to12.8 NM

1

1

1

11

1

Label 270 Discrete Word 1 FormatTable 11-2

22-14-00Page 198.162


BIT(S) FUNCTION SET

-8: -13141516171819

:;2223242526272829

LabelAll ZeroesStation 8 (Puerto Rico) usedStation A (Norway) usedStation B (Liberia) usedStation C (Hawaii) usedStation D (Dakota) usedStation E (Reunion) usedStation F (Argentina) usedStation G (Australia) usedStation H (Japan) usedStation NWC (Australia) usedStation NDT (Japan) usedStation GBR (Great Britain) usedStation NAA (Maine) usedStation NPM (Hawaii) usedStation NSS (Maryland) usedStation NLK (Washington)used

VLF iA 1

i

OMEGA 111

v 1A 1

111

vLF 111

v 1

~: Discrete label 272 is the same format as label 271 with the exception thatthe setting of a bit means the station has been either deselected by thepilot or cutoff by the RPU.

Label 271 Discrete Word 2 or Label 272Discrete Word 3 Format

Table 11-3

22-14-00Page 198.163


YOMEGANLF

ANTENNA

-i

ANALOG

k

fUVLF

INPUTRCVR

L

t

I DIGITAL DATA

DIGITAL

H

BUFFERED CLKPROCESSOR

RCVRCONTROL

I I

GND/OPENDISCRETE ~

DISCRETE

INPUTSINTERFACE

HDGITAS > SYNCHROINPUTS INTERFACE

TO FMS

{r

DIGITALlNPUT/OUTPUT

4

RPU DC

28 VDC * POWER ‘VOLTAGES REQUIREDSUPPLY 212VDC

* TOACUAD-1 5927

02-800 Receiver Processor UnitBlock DiaqramFigure 1~-2 22-14-00

Page 198.164Apr 15/93

Useor disclosure of information on this page issubject totherestfictions onthetitle page of this document.

This page is intentionally left blank.

22-14-00Page 198.165

Apr 15/93Uaeor disclosure of information on this page issubject totherestrictiona onthetitle page of this document.

11. B. AT-800 Antenna Coupler Unit (ACU) - Teardrop H-Field (See Figure 11-3and Table 11-4.)

The AT-800 ACU contains two signal-sensing elements and the first twostages of signal amplification and frequency selection. An orthogonalpair of ferrite rod loop antenna elements are used to sense theH-(magnetic) field component of the Omega/VLF signals. Since loopantennas display an inherent directional sensitivity pattern, orthogonalpositioning of the two loops provides omnidirectional coverage.Moreover, through electronic steering of the loop reception pattern, a6-dB signal-to-noise gain can be achieved. Also, loop antennas exhibitrelative immunity to noise generated by airframe electrostaticdischarges (e.g., precipitation-staticdischarges).

Integral preamplifiers comprise the amplification stages. These twostages amplify the received signals sufficiently to allow the signals tobe cable-transmittedto the OZ-800 RPU without significantsignal-to-noisedegradation. A BITE capability whereby test signals areintroduced directly into the ferrite rods, allows monitoring of ACUsense elements and preamplifier stages.

Electromagnetic interference (EMI) protection for the ACU is provided bya combination of filtering and shielding. The frequency response of theACU provides rejection of any signals outside the Omega/VLF frequencyband. All internal circuitry in this unit is shielded by a solid

metallic shield whose electrical reference point is floating with

respect to the airframe. A circumferentially grounded Faraday shield is

provided in the cover for the ACU.

./\

!2AD-1 5620

AT-800 Antenna Coupler UnitFigure 11-3 22-14-00

Page 198.166Apr 15/93

Useor disclosure of information onthispage issubject totheresttictions on the title page of this document.

D mensions

Length.Width..Height.

Weight .....

maximum

. . . . . . .

. . . . . . .

.

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7.35 in.

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5.58 in.

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1.75 in

. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2

Power Requirements ................................. ~lz V dc

(186.69 mm)(141.73 mm)

(44.45 mm)

b (0.91 kg)

0.18 W max

Mating Connector. . . . . . . . . . . . . . . . . . . . . . . . . . . . . Tracer Part No. 24952-0002

Mounting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Hard mount using four 10-32nonmagnetic hex head screws

AT-800 Antenna Coupler UnitLeading Particulars

Table 11-4

22-14-00Page 198.167


11. c. AT-801 Antenna Coupler Unit (ACU) - Brick H-Field (See Figure 11-4 andTable 11-5.)

The AT-801 ACU receives Omega/VLF signals and converts them forprocessing by the OZ-800 RPU. The AT-801 ACU is electrically equivalentto the AT-800 ACU described in paragraph 6. The AT-BO1 ACU is used onaircraft which require internal mounting within tail cones and fin caps.

AT-801 Antenna Coupler UnitFigure 11-4

22-14-00Page 198.168

Apr 15/93

Llseor disclosure of information on this page is subject to the restrictions on the title page of this document.


Length ......................................... 6.50 in. (165.10 mm)Width. ......................................... 5.65 in. (143.51 mm)Height.......................................... 1.75 in. (44.45 mm)

Weight ................................................. 2.2 lb (1.00 kg)

Power Requirements. ................................ ~lz V dc, 0.18 W max

Mating Connector ............................. Tracer Part No. 24952-0002

Mounting .................................... Hard mount using six 10-32nonmagnetic hex head screws


Table 11-5

22-14-00Page 198.169

Apr 15/93Useor disclosure of information on this page issubject totherestrictions on the title page of this document.

11. D. AT-803 Antenna Coupler Unit (ACU) - Blade E-Field (See Figure 11-5 andTable 11-6.)

The AT-803 ACU receives Omega/VLF signals and converts them forprocessing by the OZ-800 RPU. The AT-803 contains amplifier stages thatamplify the received signals sufficiently to allow the signals to becable-transmitted to the OZ-800 RPU without significant signal-to-noisedegradation. The AT-803 ACU is used when a suitable location for theAT-800 or AT-801 H-Field ACU cannot be located by an aircraft skin mapbecause of excessive RFI aircraft noise.

AD-15619

AT-803 Antenna Coupler UnitFigure 11-5 22-14-00

Page 198.170Apr 15/93



Length ........................................ lQ.35in.Width. .......................................... 3.39 in.

364.49 mm)(86.11 mm)

Height . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6.38 in. (162.05 mm)

Weight................................................. 2.2 lb (1.00 kg)

Power Requirements................................. t12 V dc, 0.18 W max

Mating Connector............................. Tracer Part No. 24627-0307

Mounting.................................... Hard mount using four 10-32nonmagnetic hex head screws


Table 11-6

22-14-00Page 198.171

Apr 15/93Useor disclosure of information onthispage issubject totherestrictions on the title page of this document.

I12. Optional LSZ-850 Liqhtninq Sensor System

il. LP-850 Lightning Sensor Processor (See Figures 12-1 and 12-2, and Table12-1.)

~O.O>booo

O.OO

>.OO>>

AD-1 5286

LP-850 Lightning Sensor ProcessorFigure 12-1

22-14-00Page 198.172



Length ................................... 14.75 in. (375 mm)Width .1”.................................... 2.42 in. (61.5 mm)Height ...................................... 7.62 in. (194mm)

Weight .......................................... 6.751b (2.7G kg)

Power ................................................ 28Vdc, lA

Mating Connector .......................... DPX2MA-67S-67S-33B-O011

Mounting ........................................... Mounting Tray,Honeywell Part No. 7011839-901

LP-850 Lightning Sensor ProcessorLeading Particulars

Table 12-1

The processor receives signals which are naturally generated bylightning activity, and determines their range from this energydistribution. At the same time, bearing is computed by means of antennacrossed loops in a manner similar to an ADF.

The output data is converted to an ARINC 429 low-speed architecture fordisplay on the ND. However, label assignments do not conform to ARINC429. The data stream contains range, bearing, and severity data for upto 50 cells. The 429 data contains all data available for a 360-degreearea with a radius of 125 NM around the aircraft, and it is the task ofthe display device to determine which cells fall within its displayarea.

Mode selection is provided by the LU-860 Lightning Sensor Controller.

When the antenna assembly is mounted on the aircraft, four manualadjustments may be required. The adjustments are intended to correctthe magnetic and electric field distortion caused by the aircraft. Thecorrection factors for each aircraft type and antenna mounting locationwill be predetermined and plainly marked on a label (Honeywell Part No.7013068-000) to be located adjacent to the LP-850. The factors areimplemented using switches S1 - S4 located on the LP-850 front panel.

22-14-00Page 198~173


28 VDC mPOWER *12VDC7-OSUPPLY * ANTENNA

I

‘N’H’R’’’+’++

YDATA

PROCESSOR

DG VALID

wow-

TAS, TAS REF

‘AS’ASREF+‘“’T’NHB------I‘R’NC’N(’)-----I‘R’NC’N(2)-----I

-,.=’===2

t-

LIGHTNINGFLAGVALID

I= 1.OUTPUTS

INPUTIOUTPUT

PROCESSOR

DATA OUT (1)

DATA OUT (2) ARINC 429OUTPUTS

DATA OUT (3) jl~jpD

r DATA OUT (4) J

1 AD-18249-R1

LP-850 Lightning Sensor ProcessorBlock DiagramFigure 1?-2 22-14-00

Page 198.174Apr 15/93



I22-14-00

Page 198.175Apr 15/93

Use or disclosureof informationon thispage issubject to the restrictionsonthe titlepage of thisdocument.

I12. B. LU-860 Lightning Sensor Controller (See Figures 12-3 and 12-4 and Tables12-2 and 12-3.)

LU-860 Lightning SensorFigure 12-3

Controller


Length ..................................... 3.900 in. (99.06 mm)Width ..................................... 5.750 in. (146.05 mm)Height ..................................... 1.125 in. (28.58 mm)

Weight ........................................... 0.61 lb (0.28 kg)

Lightning Power ......................... 5 V ac/dc, 3.5 VA maximum

Mating Connector ................................. JTO6RE1O-35S(SR)

Mounting ....................................... Unit Dzus Fasteners

LU-860 Lightning Sensor ControllerLeading Particulars

Table 12-2

22-14-00Page 198.176


The LU-860 Lightning Sensor Controller is used to select the modes ofoperation of the lightning sensor system. The switch functions arelisted in Table 12-3.

OFF All power removed.

STBY Display is inhibited but the processor isaccumulating data.

LX System is fully operational. All accumulatedlightning data is displayed.

CLR/TEST Accumulated data is cleared from memory. After3 seconds the test mode is initiated.

LU-860 Switch FunctionsTable 12-3

22-14-00Page 198:177

Apr 15/93Use or disclosureof information on thispage issubject to the restrictionson the titlepage of thisdocument.

hED 1BtK

>6

WIT )4T >2

15~x

,? Vlo t2

7

I1

,1

I-J

ii

I

II 1

i ‘ 4 ?@6°0’

III

II

13

+5v LIGUTING

LIGHTING ~

+ 29V LIGUIING

SW COMMON

OFF ~

sTaY

cl-n / TEST

ao12738-1-L

LU-860 Lightning SensorController Schematic

Figure 12-4 22-14-00Page 198.178

Apr 15/93Use or disclosure of informationon this page is subject to the restrictions onthe title page ofthis document.


I22-14-00

Page 198.179Apr 15/93


I 12. C. AT-850 Antenna (See Figure 12-5 and Table 12-4.)

AD.139S6

AT-850 AntennaFigure 12-5


Length ................................... 11.58 in. (294.13 mm)Width ..................................... 6.06 in. (153.92 mm)Height ..................................... 1.25 in. (31.75 mm)

Weight ........................................... 2.5 lb (1.13 kg)

Power ........................................ ~12 V dc from Lp-850

Mating Connector ..................................... KJ6F12A35-SN

Mounting ................ Hard Mount Using Three No. 10 Screws WithMaximum Washer Diameter of 0.455 in.

AT-850 AntennaLeading Particulars

Table 12-4

22-14-00Page 198.180

Apr 15/93Useor disclosure of information onthispage issubject totherestrictions on the title page of this document.

The AT-850 Antenna contains crossed loop H-field antennae and an E-fieldantenna similar to an ADF antenna.

Preamplifier stages are also built into the antenna in order to enhancethe system’s immunity to noise originating in aircraft wiring.

The H-field loop antennae are designated Hn and Hw and are orientated insuch a manner that Hn will be most sensitive to signals originatingahead or behind the aircraft, and the Hw antenna will be most sensitiveto signals originating abeam the aircraft.

The E-field antenna is constructed such as to be most sensitive tovertical E-fields.

~12 V dc power for the preamplifiers is provided to the antenna from the

processor.

The antenna also contains a test winding. During test mode this windingis driven with a simulated lightning signal and couples with the E- andH-elements of the antenna. Thus, the test mode is able to provide anevaluation of the performance of the antenna, its preamplifiers, andcabling to the LP-850 Lightning Sensor Processor.

The AT-850 Antenna is designated for mounting external to the aircraftand is not to be painted after installation. Also, the antenna isencapsulated and is not repairable.

22-14-00I

Page 198.181Apr 15/93


I 12. D. AT-855 Antenna (See Figure 12-6 and Table 12-5.)

AT-855 AntennaFigure 12-6


We

Length .................................... 6.75Width .................................... G.OGOHeight ..................................... 1.67

ght ........................................... 2

Power ........................................ flz v

n. (171.45 mm)n. (152.40 mm)in. (42.42 mm)

5 lb (1.13 kg)

dc from LP-850

Mating Connector ..................................... KJ6F12A35-SN

Mounting ................... Hard Mount Using Four No. 8 Screws WithMaximum Washer Diameter of 0.350 in.

AT-855 AntennaLeading Particulars

Table 12-5

22-14-00Page 198.182


The AT-855 Antenna contains crossed loop H-field antennae and an E-fieldantenna similar to an ADF antenna.

Preamplifier stages are also built into the antenna in order to enhancethe system’s immunity to noise originating in aircraft wiring.

The H-field loop antennae are designated Hn and Hw and are orientated insuch a manner that Hn will be most sensitive to signals originatingahead or behind the aircraft, and the Hw antenna will be most sensitiveto signals originating abeam the aircraft.

The E-field antenna is constructed such as to be most sensitive tovertical E-fields.

flz V dc power for the preamplifiers is provided to the antenna from theprocessor.

The antenna also contains a test winding. During test mode this windingis driven with a simulated lightning signal and couples with the E- andH-elements of the antenna. Thus, the test mode is able to provide anevaluation of the performance of the antenna, its preamplifiers, andcabling to the LP-850 Lightning Sensor Processor.

The AT-855 Antenna is designated for radome mounting and is not to bepainted after installation. Also, the antenna is encapsulated and isnot repairable.

22-14-00I

Page 198.183Apr 15/93


Dimensions (maximum)

6 MCU Short per ARINC 600Height . . . . . . . . . . . . . . . . . . . 7.64 in. (194.06 mm)Width. . . . . . . . . . . . . . . . . . . . 7.75 in. (196.85 mm)Length (from rear of connector) . . . . . . . 12.75 in. (323.85 mm)

Weight . . . . . . . . . . . . . . . . . . . . . . . . 24.01b(10.9 kg)Power Requirements . . . . . . . . . 115 V ac 400 Hz, 80 Watts (maximum)Mating Connector . . . . . . . . . . . . . TRI-STAR PN C-O6B5-99O1-O1OOMounting: . . . . . . . . . . . . . . . . . . . . . . Non Honeywell Tray

RT-91O TCAS Computer Leading ParticularsTable 13-1

The TCAS computer contains the RF transmitter and receivers necessary tointerrogate and receive replies from transponder equipped aircraft. Dualmicroprocessors are utilized to implement the surveillance and collisionavoidance algorithms to decide whether an intruder aircraft should beconsidered a threat, and then to determine the appropriate vertical responseto avoid a midair collision or near midair incident. In addition, outputdata is provided to drive displays, and the aircraft audio system, to informthe flight crew as to what action to take or avoid. An interface isprovided with an on-board Mode S transponder in order to coordinate avoid-ance maneuvers with other TCAS equipped aircraft. The TCAS computertransmits ARINC 429 high speed output data to the SG-884 Symbol Generatorsas specified in Table 13-2.

22-14-00Page 198.185


Parameter/Signal liame Label Rate

Control Panel SetAltitude SelectTCAS Mode/SensIntruder Range *Intruder Altitude *Intruder Bearing *Own Aircraft AltitudeVertical RAHorizontal RA

Select TCAS Sensitiv-ityMaintenanceSTXData CharactersEOTRTS/ETXEquipment ID

013015016130131132203270271274

350356356356357377

2-3 Hz2-3 Hz2-3 I-Iz

2 Hz2-3 Iiz

2-3 Hz2-3 Hz

NOTE: * Labels 130, 131, and 132 are repeated foreach successive intruder for display.

RT-91O TCAS Computer ARINC 429 Output DataTable 13-2

22-14-00Page 198.186


The TCAS Computer also transmits ARINC 429 high speed output data to theMode S Transponder as listed in the following table, This data is forcoordination with other TCAS-11 equipped aircraft.

Parameter/Signal Name Label Rate

Control Panel SetAltitude SelectTCAS Mode/SensOwn Aircraft AltitudeVertical RAHorizontal RA

Select TCAS Sensitivity

Maintenance

013015016203 2 Hz270 2-3 Hz271 2-3 Hz271 2-3 Hz273274 2-3 HZ275276277350

RT-91O TCAS Computer to Mode S Transponder DataTable 13-3

The Mode S Transponder transmits ARINC 429 high speed output data to theTCAStion

Computer as-listed in the following tabl~. This data is for coordina-with other TCAS-11 equipped aircraft.

Parameter/Signal Name Label Rate

274Select TCAS Sensitivity 274 2-3 HZ

275

Mode S Transponder To TCAS Computer DataTable 13-4

22-14-00Page 198.187


SUPPRESSIONBUS

TOP ANT’

COAX

BOITOM ANTCOA%

PART OFARINC 600

CONNECTOR

I

PART OFARINC 600CONNECTOR

i

4I i*I

-; 4I * ~

- ,’4 I* RHO A6 ~ SUR;;;~FNCE

I REC%ER ~ 1!0PROCESSOR

III i%

4:: 4

ITRANsh4~ER _

IIIIIIII

i-1

i115 V AC 1

Al 14400 Hz -

IPOWER

POWER SUPPLY I- OUTPUT DISCR=ES

II 2,

I- AURAL OUTPUTS

;4 !

SITE 1 2,ANALOG INPUTS

A2INDICATORS ~

DIGITAL VIDEOCAS CPU} ~ INTERFACE

(SYNCI-IRO. RADIO _ALTIT~OE) z AIRCRAFT 1

1/0 I- aECORDEE INTERFACE

OISCR~ INPUTS-;14 I

I -pROCESSOR I 4

~ ARINC 429 0LJT~L17s

ARINC 429 ,5 1INPUTS I * I

I I:}

ARINC 615 LOAOEFINTERFACE

I II I

c86-1632#

RT-91O TCAS Computer Block DiagramFigure 13-2

22-14-00Page 198.188

~pr 15/93Use or disclosure of information on this page is subject to the restrictions on the title page of this document,

13. B. AT-91O Directional Antenna (See figure 13-3 and table 13-5.)

AD-32828@

AT-91O Directional AntennaFigure 13-3


Height Outside Aircraft . . . . . . . . . . . . . .806 in. (20.47 mm)

Height Inside Aircraft . . . . . . . . . . . . . 1.56 in. (39.62 mm)Diameter . . . . . . . . . . . . . . . . . . . 9.31 in. (236.47 mm)

Weight . . . . . . . . . . . . . . . . . . . . . . . . .2.801b(l.30 kg)Mating Connectors:(4) . . . . . . . . . . . . . . . . . . . .. TypeTNCMounting Adapter . . . . . . . . . . . . . . . . . . . . . HPN 7514081-910

AT-91O Directional Antenna Leading ParticularsTable 13-5

The AT-91O directional antenna mounted on top of the aircraft fuselagein conjunction with four receivers in the TCAS computer unit provide thecapability to determine the bearing of the intruder. Since TCAS II is avertical-only system, intruder bearing is not used in the computation ofthe escape or limit maneuver. Intruder bearing is used only to enablethe flight crew to more easily locate the intruder visually. TheHoneywell TCAS II will also accommodate a bottom mounted directionalantenna if so desired by the user (refer to paragraph 13.C.). Such aninstallation will slightly increase the areas in which bearing informa-tion is available. The AT-91O directional antenna is connected to theTCAS computer unit by four coaxial cables.

22-14-00Page 198.189


The antenna is capable of receiving replies from all directions simulta-neously with bearing information using amplitude-ratio monopulsetechniques. Insertion loss differences in coaxial cable lengths fromthe antenna to the TCAS computer need only be matched to within 0.5 dBwhich corresponds to a 5 to 10 foot difference in length depending onthe specific cable type. Losses between the antenna and the computerunit must be 2.5 t 0.5 dB, including line connections.

All antenna lobes work across the face of the antenna. J1 is located atthe rear of the antenna and produces the forward looking lobe. Inaddition the antenna ports are both color-coded and electrically codedwith a resistance to ground that is checked by the TCAS computer and canbe checked by the technician.

J1

J2

J3

J4

O degree port. Color coded yellow. Located on the rear of theantenna and installed to the rear of the aircraft. Resistance toground is 1000 ohms.

270 degree port. Color black. Installs toward right wing.Resistance to ground is 8000 ohms.

180 degree port. Color blue. Installs toward front of aircraft.Resistance to ground is 4000 ohms.

90 degree port. Color red. Installs toward left wing. Resistanceto ground is 2000 ohms.

22-14-00Page 198.190


13. c. Typical Bottom Omnidirectional Antenna (See figure 13-4.)

——— ——— ——

AD32827

Typical Bottom Omnidirectional AntennaFigure 13-4

The leadinq particulars for the bottom omnidirectional antenna must beobtained f~orn the selected antenna manufacturer.

The omnidirectional antenna mounted on the bottom of the afuselage does not provide any directional information, butthe TCAS computer to interrogate and receive replies from ~aircraft located beneath own aircraft. Since TCAS-11 is asystem, intruder bearing is not used in the computation of

limit maneuver.

The omnidirectional antenna mustexhibit 50 ohms fixed resistancetest. Identical omnidirectionalbottom antennas connected to the

rcraftdoes enablentrudervertical-onlythe escape or

exhibit 50 ohm imDedance., and must alsoto ground. This is necessary for selfantennas may be used for the top andMode S transponder.

22-14-00Page 198.191


14. Optional MLZ-850 Microwave Landing System (MLS)

A. ML-850 Microwave Landing System Receiver (See Figures 14-1 and 14-2,Tables 14-1 and 14-2.)

AD-1374s.m

ML-850 MLS ReceiverFigure 14-1

and


Length ......................................... 14.01 in. (355.9 mm)Width ........................................... 3.08 in. (78.2 mm)Height .......................................... 3.36 in. (85.3 mm)

Weight ................................................. 4.9 lb (2.22 kg)

Power Requirements ....................... 18t032Vdc, 15.5 VA nominal

Mounting Kit ........................................ Part No. 7510664-901

NOTE : The mounting kit contains the following components that can beordered separately:

Mount, Frame (Tray) ............................. Part No. 7510654-901Spring, Finger ...................................... Part No. 7510835Connector, 67 Pin ................................ Part No. 7500297-67Screw, CSKH CRES 4-40 x 3-8 ......................... Part No. 0920-15

ML-850 MLS ReceiverLeading Particulars

Table 14-1 22-14-00Page 198.192

Apr 15/93Use or disclosure of information on this page is subject to the restrictions on the title page of this document

NOTE : The following RF connector kits contain two RF connectors and tworetainer rings. The retainer rings can be ordered separately underPart No. 0493-18.

IK-858RF Part No. 7500436-90X

RF Connector Kit for use with Mounting Tray. Two kits required foreach MLS Receiver installation. Select from the following:

Part No. 7500436-907 - Connector Kit with RF Straight Connector(RG-214)

Part No. 7500436-908 - Connector Kit with RF Right Angle Connector(RG-214)

Part No. 7500436-909 - Connector Kit with RF TNC Adapter Connector

ML-850 MLS ReceiverLeading ParticularsTable 14-1 (cent)

The ML-850 MLS Receiver decodes and processes data from an MLS groundstation and provides an accurate indication in both azimuth (equivalent tolocalizer) and elevation (equivalent to glide slope) of the deviation fromthe desired flight path. The deviation data is displayed on the PFD and isoutput to the FZ-820 Flight Guidance Computer for use in the approach modeof operation.

The ML-850 operates in the frequency range of 5031.0 to 5090.7 MHz on 200channels spaced 3000 kHz apart. Selection of the desired azimuth andelevation angle and tuning is accomplished with the CM-850 MLS Con-trol/Display Unit.

The ML-850 provides ARINC 429 digital outputs (Table 14-2) that conform tothe data standards of ARINC characteristic 727 for MLS receivers.

22-14-00Page 198.193


Labelw

017033035036105130131132133134135136137140141142143151152153154155156157160161162163164165173174175176177256257270376377

Data Descrir)tion

Runway HeadingILS/MLS FrequencyMLS/DME FrequencyMLS Channel NumberRunwav HeadinqMLSMLSMLSMLSMLSMLSMLSMLSMLSMLSMLSMLSMLSMLSMLSMLSMLSMLSMLSMLSMLSMLSMLSMLSMLSMLSMLSMLSMLSMLSMLSMLSMLSGenMLS

Afixiliary-DataAuxiliary DataAuxiliary DataAuxiliary DataAuxiliary DataAuxiliary DataAuxiliary DataAuxiliary DataAuxiliary DataAuxiliary DataAuxiliary DataAuxiliary Data

Word 1AWord 2AWord 3AWord 4AWord lBWord 2BWord 3BWord 4BWord lCWord 2CWord 3CWord 4C

Azimuth DeviationGlidepath DeviationSelected Azimuth AngleMax Selectable Glidepath AngleSelected Glidepath AngleBasic Data Word 1Basic Data Word 2Basic Data Word 3Basic Data Word 4Basic Data Word 5Basic Data Word 6Absolute Glidepath AngleAbsolute Azimuth AngleLocalizer DeviationGlide Slope DeviationSelected Back Azimuth AngleBack Azimuth Absolute AngleBack Azimuth DeviationGround Station Identification 1Ground Station Identification 2DiscretesAV Equipment IdentificationSpecific Equipment Identification

ML-850 MLS ReceiverARINC 429 Outputs

Table 14-2

22-14-00Page 198.194

Apr 15/93Use or disclosure of information on this page issubject totheresttictions onthetitle page of this document.

MLS ANTENNAs

TTF———— ——— aI CPU CARD

It + DPSK

II ANTENNA

SWITCHZso

MAIN

A4

PROCESSOR

I

II TCXO !

IIII

I RC6 TO

II BITECONTROL

UNIT

L——————— ————— —.——

+5 +28+15 -15

y

r——. ——— ——— ——

II POWER

SUPPLY

I

L ——— ——. ——— ——— ——— .—— ——— d28 VDCINPuT

POWER AO.19591

ML-850 MLS Receiver Block DiagramFigure 14-2

22-14-00Page 198.195

Apr 15/93Useor disclosure of information onthispage issubject totherestrictions onthetitle page of this document,

14. B. CM-850 Control/Display Unit (see Figures 14-3 and 14-4 and Table 14-3.)

GLIDEPATHSELECTPUSHBUTTON=

CHANNELSELECTPUSHBUTTON—_

MODESELECTSWITCH.

ON/OFFSWITCHVOLUMECONTROL~KNOB(OPTIONAL)

BfiCKAZIMUTH f$jBunoNSELECTPUSHBUTTON

CM-850 MLS Control/Dis~laY Unit

PUSHBUTTON

YTUNINGKNOBS

AD-20722

Figure 14-3 ‘ -

22-14-00Page 198.196



Length (from back of bezel) ...................... 5.90 in. (149.9 mm)Width ........................................... $.:; ~;. (60.3 mm)Height .......................................... . . (66.7 mm)

Weight (maximum) ........................................ 1.251b. (567g)

Power Requirements ......................... 18t032Vdc, 4VA maximum

Lighting ............................... 5 V acldc or 28 V dc; Brightnessis 0.25 to 0.75 FT-L at 5077 of

maximum voltageCurrent = 10 mA maximum

Display Type ............................. Dichroic LCD, white characterson black background

Mating Connector ......................................... MS3126F20-41SW

Mounting ..................................... Panel mount, self clamping

CM-850 MLS Control/Display UnitLeading Particulars

Table 14-3

The CM-850 MLS Control Display Unit (CDU) is the main tuning source forthe MLS system. The CDU provides access to four separate functions;channel number/ident (CH), glidepath (GP, azimuth (AZ), and blackazimuth (BK). Each function has an annunciator, an arrow cursor and apushbutton associated with it. Depressing any one of the pushbuttonsselects that function for programming, and turns on the associated arrowcursor. The pushbuttons are all labeled as to their function. Thefollowing paragraphs describe the operation of each control.

14. B. (1) Rotary MODE Switch

Rotating the MODE switch (clockwise or counterclockwise)to thenext position will cause the mode to change. If in AUTO mode, themode will change to MAN mode. If in MAN mode, the mode will changeto AUTO mode.

22-14-00Page 198.197


14. B. (2) CH Pushbutton

Pressing the CH pushbutton will allow the channel to be changed viathe rotary tuning knobs. The channel selector is turned on and thechannel number will be displayed. If the CH pushbutton is held for2 seconds then the CH annunciator should turn off and the identshould be displayed. The ident will be displayed until the channelpushbutton is released or after 30 seconds (whichever occursfirst) . This will allow the channel number to be displayed in theevent that the CH pushbutton were to stick. After the 30 secondtimeout has occurred then the CH pushbutton must be released andpressed again before the ident will be displayed. The CH annuncia-tor will turn on again when the pushbutton has been released orwhen the timeout occurs.

If BACK is being displayed and the CH pushbutton is pressed thenthe BACK annunciator will be turned off and the AZ display willthen be displaying the forward azimuth. Also, the glidepath digitswill be unblanked and will display the glidepath angle.

(3) AZ Pushbutton

Pressing the AZ pushbutton will allow.the azimuth angle to bechanged via the rotary tuning knobs. If BACK is being displayedand the AZ pushbutton is pressed then the BACK annunciator will beturned off and the AZ display will then be displaying the forwardazimuth. Also, the glidepath digits will be unblanked and willdisplay the glidepath angle.

(4) GP Pushbutton

Pressing the GP pushbutton will allow the glidepath angle to bechanged via the rotary tuning knobs. If BACK is being displayedand the GP pushbutton is pressed then the BACK annunciator will beturned off and the AZ display will then be displaying the forwardazimuth. Also, the glidepath digits will be unblanked and willdisplay the glidepath angle.

(5) BK Pushbutton

This pushbutton performs a toggle function between forward azimuthand back azimuth modes. Pressing this pushbutton selects the backazimuth mode if previously displaying forward azimuth and allowsthe back azimuth angle to be changed. Pressing this pushbuttonselects the forward azimuth mode if previously displaying backazimuth.

The glidepath display will be blanked when the MLS CDU is in BACKAZ mode.

22-14-00Page 198.198

Apr 15/93Use or disclosure of informationon this page is subjectto the restrictionson thetitle pageof this document,

14. B. (6) Rotary Tuning Knobs

(a) Channel Tuning

The rotary tuning knobs can change the channel number wheneverthe channel selector is on. Whenever the channel number ischanged, the mode is automatically changed to AUTO.

~ Inner Knob

The inner knob controls only the least significant digit ofthe channel number. Turning the knob clockwise will causethe number to increase in the range from O to 9 by 1 unit.When the number reaches 9, the next rotation will cause thenumber to change to O.

Turning the knob counterclockwisewill cause the number todecrease in the range from 9 to O by 1 unit. When thenumber reaches O, the next rotation will cause the numberto change to 9.

~ Outer Knob

The outer knob changes only the most significant 2 digits.Turning the knob clockwise will cause the number toincrease in the range from 50 to 69 by 1 unit. When thenumber reaches 69, the next rotation will cause the numberto change to 50. Turning the knob counterclockwise willcause the number to decrease in the range from 69 to 50 by1 unit. When the number reaches 50, the next rotation willcause the number to change to 69.

(b) Azimuth Selector

The rotary tuning knobs can change the azimuth angle wheneverthe azimuth selector is on. If the mode is AUTO when thisswitch is turned, the first position change of the rotaryswitch will change the mode to MAN. The next position changeof the switch will change the azimuth value.

~ Inner Knob

The inner knob will increment or decrement the azimuthangle by 1 degree steps for each position of rotation. Theazimuth angle will rollover from 359 to 000 for anincrement and from 000 to 359 for a decrement. Clockwiserotation will increment the display and counterclockwiserotation will decrement the display.

22-14-00Page 198.199


14. B. (6) (b) ~ Outer Knob

The outer knob controls only the most significant 2 digitsof the azimuth angle. Turning the knob clockwise willcause the number to increase in the range from 00 to 35 by1 unit. When the number reaches 35, the next rotation willcause the number to change to 00. Turning the knobcounterclockwise will cause the number to decrease in therange from 35 to 00 by 1 unit. When the number reaches 00,the next rotation will cause the number to change to 35.

(c) Glidepath Selector

The rotary tuning knobs can change the glidepath anglewhenever the glidepath selector is on. If the mode is AUTOwhen this switch is turned, the first position change of therotary switch will change the mode to MAN. The next positionchange of the switch will change the glidepath value.

~ Inner Knob

The inner knob increments or decrements the glidepath angleby 0.1 degree steps for each rotation of the switch. Norollover of the glidepath takes place for an incrementabove the maximum glidepath or a decrement below 2.0degrees. A 9 to O or O to 9 transition of the LSD doescarry over into the MSD’S. Values are incremented withclockwise rotation and decremented with counterclockwiserotation.

~ Outer Knob

The outer knob is required to increment or decrement theglidepath by 1.0 degree steps for each rotation of theswitch. The glidepath angle is to be set to the maximumglidepath (this includes the LSD) if the increment causesthe maximum glidepath to be exceeded. The glidepath angleis to be set to 0.0 (this includes the LSD) if the two mostsignificant digits are O after a decrement and the switchis rotated (decremented) again. The two most significantdigits never rollover. Values are incremented withclockwise rotation and decremented with counterclockwiserotation.

22-14-00Page 198.200


14 B. (6) (d)

(7)

Back Azimuth Selector

Back Azimuth selected angles can be changed if the backazimuth selector is on. If the switch is rotated when inAUTO, the first position change of the rotary switch willcause the mode to change to MAN. The next position changewill change the back azimuth value.

~ Inner Knob

The inner knob will increment or decrement the back azimuthangle by 1 degree steps for each position of rotation. Theback azimuth angle will rollover from 359 to 000 for anincrement androtation willrotation will

~ Outer Knob

f~om 000 to 359 for a decrement. Clockwiseincrement the display and counterclockwisedecrement the display.

The outer knob controls only the most significant 2 digitsof the back azimuth angle. Turning the knob clockwise willcause the number to increase in the range from 00 to 35 by1 unit. When the number reaches 35, the next rotation willcause the number to change to 00. Turning the knobcounterclockwise will cause the number to decrease in therange from 35 to 00 by 1 unit. When the number reaches 00,the next rotation will cause the number to change to 35.

TEST Pushbutton

Pressing and holding the TEST pushbutton for 2 seconds will causethe MLS radio to be tested. Pressing this pushbutton for less than2 seconds is the same as not pressing it at all. The CDU willdisplay “TST” in the channel display during the time period thatthe radio is being tested and until the radio test results havebeen reported. The test is performed only when the pushbutton isbeing held. The test shall terminate and the MLS CDU shouldfunction normally if the pushbutton is released or after 30seconds, whichever occurs first.

The alphanumeric channel/ident display is used for displaying theresults of the radio test. If the MLS radio passes the test theMLS CDU will display “OK”. If the MLS radio fails its test thenthe MLS CDU will display “ERR”.

The azimuth and glidepath displays will be blanked when in the TESTmode.

22-14-00Page 198.201


II

RCB TX4 RS422 DATA4

RCB RX~TRANS-

9DISPLAY

CEIVER ENABLE- DRIVERSDISPLAY

.—

1RS422 NAV CTL + )

IPROCESSOR

RS422 FMS RECEIVER 4 )

1 I DATA

RS422 DME ~ RS4224

CLOCK SWITCH

a

9

TUNING DRIVER ISHIH SWITCHES

ENABLE- REGISTERS

2x50ME I 2x5

TUNINGLATCH d

DRIVERS 1

=“’:’m

L .—— —. —4I I

TEST INHIBIT *DISCRETE

1/04

,

lLU%%4T~LU%f&&T

28 VDC

ON/OFF

INPUT POWER_-30 VDC

/ POWER

OUTPUT ~ SUPPLY~-15 Voc

++5 VDC

\

(DIMMING 5vAC/DCINPUTS 28 VDC AO-17049-R2

CM-850 MLS Control/Display Unit Block DiagramFigure 14-4

22-14-00Page 198.202


This page is intentionally left blank.

22-14-00Page 198.203


15. opt onal Global Positioning System (GPS)

Global Positioning System Sensor Unit (GPS) (See F gures 15-1 and 15-2, andTables 15-1 through 15-5.)

Global Positioning System Sensor Unit (GPSSU)Figure 15-1


Length . . . . . . . . . . . . . . . . . . . . .8.50 in. (215.9mm)

Width. . . . . . . . . . . . . . . . . . . . . . 8.50 in. (215.9 mm)

Height . . . . . . . . . . . . . . . . . . . ..2.20 in. (55.8 mm)

Weight . . . . . . . . . . . . . . . . . . . . . . . . . 5.01b(2.27 kg)

Power Requirements . . . . . . . . . . . . . . 28 V dc, 36 Watts (maximum)

Mating ConnectorJ1 . . . . . . . . . . . . . . . . . . . . . . . . . . . M83723/77R2041NJ2 . . . . . . . . . . . . . . . . . . . . . . . . . . . .. TNC Female

Mounting . . . . . . . . . . . . . Hard Mount Using Four 10-32 Cap Screws

Global Positioning System Sensor Unit Leading ParticularsTable 15-1 22-14-00

Page 198.204Apr 15/93


The GPSSU contains the necessary power supplies, receiver circuitry andelectronics to receive signals transmitted by the NAVSTAR satellites, and tocompute present position, altitude, true track and groundspeed, and tooutput these data on an ARINC 429 databus. If the GPSSU is not able tomaintain track of at least four satellites, it uses pressure altitude fromthe ADC, and received data from the remaining satellite(s) to computepresent position. If the GPSSU is not able to track any satellites for 30seconds, it reverts to the Acquisition Mode. During this mode, the GPSSUaccepts position data from the FMS, and transmits that data (which isidentified as FMS data), until it has acquired at least four satellites,when it re-enters the Navigation Mode.

The GPSSU operates in five normal modes and indicates the current mode ofoperation on the ARINC 429 output bus maintenance label. In addition to theGPSSU modes, the signal processor circuits operate in two modes. Anadditional mode, the acceptance test procedure (ATP) mode, is used fordevice testing. Each mode operates as described below:

●

●

●

●

9

9

Self-Test Mode - During the self-test mode, the GPSSU tests its circuitsto verify proper operation. If the self-test passes, the GPSSUimmediately enters the initialization mode. If the self-test fails, theGPSSU enters the fault mode.

InitializationMode - When the circuits have been tested, the GPSSUinitializes those circuits for operation.

Acquisition Mode - During the acquisition mode, the GPSSU beginsacquiring satellite data which includes ephemeris and almanac data. Whenthe GPSSU has acquired enough satellite data to compute position, itenters the NAV mode. The GPSSU reverts to the acquisition mode from thealtitude-aiding/clock-coastingsubmode when it is unable to track anysatellites for 30 seconds.

Navigation (NAV) Mode - In the NAV mode, the GPSSU updates and transmitsdata on the ARINC 429 data bus to its interfaces. The data, whichincludes latitude, longitude, altitude, time, and velocity, are derivedfrom pseudo range and pseudo range rate measurements. These measurementsare performed seven times a second. The GPSSU remains in the NAV mode aslong as it is able to track four satellites. If it is unable to track~~~;o;~tellites, the GPSSU enters the altitude-aiding/clock-coasting

.

Altitude-Aiding/Clock-Coasting Submode - The GPSSU enters the altitude-aiding/clock-coasting submode from the NAV mode when it is unable totrack four satellites. In this submode, the GPSSU uses inertial orpressure altitude inputs to determine position and other data. The GPSSUremains in this submode as long as one to three satellites are beingtracked. When the GPSSU has acquired four satellites, the GPSSUre-enters the NAV mode. If the GPSSU cannot track any satellites for 30seconds, the GPSSU reverts to the acquisition mode.

Fault Mode - The fault mode occurs when built-in test equipment (BITE)detects a critical failure. In this mode, all outputs are”inval~d. “

The signal processor circuits have two modes of operation; continuoustracking and automatic frequency control (AFC) sequencing.

22-14-00Page 198.205


. Continuous Tracking Mode - In this mode, the signal processor circuitsperforms continuous phase lock loop (PLL) tracking and navigational datacollection.

. Automatic Frequency Control (AFC) Sequencing Mode - In this mode, channelone of the signal processor circuits performs AFC tracking withoutcollecting navigational data. This mode is used to perform measurements.

The ATP mode is used only during shop testing to verify that the GPSSUperforms within specifications. The GPSSU enters the ATP mode when itreceives a test discrete input and specific ARINC test inputs. The ATP modeconsists of an alternating sequence of power-up BITE tests and satellitesignal tracking.

The accuracy and resolution of the output data provided by the GPSSU islisted in Table 15-2.

PARAMETER

1. Present Position

2. Ground Speed

3. Track AngleTrue****

4. Vertical Velocity

5. Altitude

6. N-S Velocity

7. E-W Velocity

8. Time***

SELECTIVEAVAILABILITY

OFFON

OFFON

OFFON

OFFON

OFFON

OFFON

OFFON

OFFON

LIMITATION

HDOP = 1.72

HDOP = 1.72

HDOP = 1.72

HDOP = 1.72VDOP= 2.2

HDOP = 1.72VDOP = 2.2

HDOP = 1.72

HDOP = 1.72

HDOP = 1.72

AUTONOMOUS

25 meter100 meter

1.8 Knots**1.8 Knots**

0.30**0.30**

100 ft/min100 ft/min

138 ft551 ft



350 ns450 ns

HYBRID

25 meter100 meter

0.3 Knots*0.7 Knots*

0.07°0.18”

20 ft/min25 ft/min

138 ft551 ft



350 ns450 ns

* The resolution of the ARINC output is 0.125 knots.** The track anqle, velocity and speed errors will qrow to Iarqer values

during aircr;ft maneuver;. In a 2G turn, this e;ror nominaily grows to6.0 knots for the velocities and 2.2” for track angle.

*** The time accuracy is internal. The time is associated with a hardware“time mark.” The system capability depends on the receiving equipment.

**** Based on 160 knot ground speed.

GPSSU Digital Accuracy and ResolutionTable 15-2

22-14-00Page 198~206

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The GPSSU provides ARINC 429 high speed output data in three different dataformats as listed in the following tables. Table 15-3 lists those wordsthat are transmitted in Binary (BNR) data format. Table 15-4 lists thosewords that are transmitted in Binary Coded Decimal (BCD) data format. Table15-5 lists those words that are transmitted in the DIS data format, andwhose bits denote specific discretes. The output data is updatedapproximately once per second. Specific digital data word format can befound in the Installation Manual for the GPSSU, Pub. No. 95-8698.

Binary (BNR) Data Format Label Units DigitalParameter/Signal Name Range Resolution

Pseudo RangePseudo Range FinePseudo Range RateDelta RangeSatellite Position XSatellite Position X FineSatellite Position YSatellite Position Y FineSatellite Position ZSatellite Position Z FineUTC Measure TimeGPS Altitude (MSL)HDOPVDOPTrack Angle - TrueGPS LatitudeGPS LongitudeGPS Ground SpeedLatitude FractionsLongitude FractionsVertical Figure of MeritUTC FineUTC Fine FractionsUTCVertical VelocityN/S VelocityE/W VelocityHorizontal Figure of Merit

061062063064065066070071072073074076101102103110111112120121136140141150165166174247

MetersMetersMeters/SecondMetersMetersMetersMetersMetersMetersMetersSecondsFeet

DegreesDegreesDegreesKnotsDegreesDegreesMetersSecondsSecondsHR:MIN:SECFeet/MinuteKnotsKnotsMeters

&26&13!j456

256~4096*4096~67108864

64~6710886464~671088646410~131072

10241024~180~18t)~18(140961.716E-41.716E-4102419.537E-79.537E-723:59:59*32768*4096~40961024

2560.1250.00390.0039640.0039640.0039640.00399.5367E-60.1250.0310.0310.00551.716E-41.716E-40.1258.38E-88.38E-80.031

9.313E-111.00.1250.1250.031

GPSSU ARINC 429 Output Data (BNR Format)Table 15-3

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Binary Coded Decimal (BCD)Data Format Label Units Digital

Range ResolutionParameter/Signal Name

UTC 125 HR:MIN 23:59.9 0.1 Min.

Date 260 D:M:Y 1

Equipment ID 377

GPSSU ARINC 429 Output Data (BCD Format)Table 15-4

Discrete (01S) Data Format Label UnitsDigital

Parameter/Signal Name Range Resolution

GPSSU Status 273

Maintenance Discrete No. 2 352

System Time Counter 354 Seconds 262144 1

Maintenance Discrete No. 1 355

GPSSUARINC 429 Output Data (DIS Format)Table 15-5

22-14-00Page 198.208


Y

cDADC 1 2

cDADC2 2

cFMC/lRS1 2

cFMc/lRs2 2

DADC INPUT4191429SELECT -

429OUTPUTHSILSSELECT -

BEGINATP +

+28V -+28 VRETURN -

+G202lAB GPSSU

)REALTIME

ARINC419(575) CLOCKOR 429 1Hz12.5KHZ

[

)ARINC 42912.5 KHz

J’OR100 KHz ARINC 429

12.5 KHzOR

1

100KHz

OPEN/

GROUND

DISCRETE

J INPUTSOPEN/GROUND

DISCRETE OUTPUT

}POWER

b2 TIMEMARK #1

b2 TIME MARK #2

P2 TIME MARK #3

B2 429OUT #1

2 429OUT #2

P2 429OUT #3

l--GPSSUFAULT

8698/301OS 60092-01 PT1 ~

Global Positioning System Sensor UnitBlock DiagramFigure 15-2

22-14-00Page 198.209/198.210


SECTION 3SYSTEM OPERATION

1. General

This section describes the operation of the SPZ-8000 DAFCS by separating itinto three major subsystems that are contained in the following paragraphs.

EDZ-884

DFZ-820

FMZ-800

Subsvstem paraclraDh

Electronic Display System 2

Dual Flight Guidance Control System 3

Flight Management System 4

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2. EDZ-884 Electronic Dis~lav Svstem

The cockpit’s display system configuration is shown in Figure 201. Eachpilot has a display controller and corresponding EFIS displays (PFD and ND).The center panel houses the EICAS displays (ENGINE and CAS). The remotedisplay unit dimming panel is located in the pedestal area.

A. Display System Formats

Each display format will be discussed in detail in the followingsections. Generally, the following four formats are available fordisplay in the GIV system:

. Primary Flight Display (PFD)● Navigation Display (ND) - MAP, COMPASS, or PLAN● Engine Instrument Display (EI)● Crew Alerting System (CAS) With System Page Display

B. EFIS/EICAS System Components

(1)

(2)

(3)

Display Units (DU)

The GIV system contains six color display units as shown in Figure201. The DU is the media by which flight data is conveyed to boththe pilot and copilot.

Display Controllers (DC)

The GIV display system contains two display controllers. The DCSare used to control the various display unit formats. Basically,each pilot can control the formats for the on-side PFD and NDdisplays in addition to the ENGINE and CAS displays.

Symbol Generators (SG)

The GIV display system has three symbol generators, each onecontaining the four formats described in paragraph 2.A. The purposeof the SG is to supply all the displayed data to the individualdisplay units in the system. Each symbol generator has thecapability to drive all six display units in the event the other twofail. With two or three SGS operational, all six display units haveunique formats displayed. With one SG operational, the Copilot’s

displays (PFD and ND) are repeaters of the pilot’s displays.Normally (3 SGS operational), SG 1 drives the pilot’s displays, SG 2drives the copilot’s displays, and SG 3 drives the EICAS displays.

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PILOT’S DC

Z-I

COPILOT’S DC

-j

1.PILOTND

ENGINE

CASSYSTEM

Elm

‘1’’”%=\Cockpit Layout of the EFIS/EICAS Display System

Figure 201

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2. B. (4) Fault Warning Computers (FWC)

The GIV system contains two fault warning computers. The FWCprovides all the messages to the SGS for display on the CAS format.Each FWC is doing the same computations on aircraft system inputs.Only one FWC is displayed at a time, making the other a hot spare.

(5) Data Acquisition Unit (DAU)

The GIV system contains two data acquisition units. The DAUSprovide all the aircraft engine information in addition to theanalog data of some subsystem parameters to the SG for display onthe ENGINE display. DAU 1 provides left engine data and DAU 2provides right engine data.

(6) Display Dimming Panel

One display dimming panel has been incorporated into the GIV systemfor the purpose of adjusting the brightness of the six displayunits.

(7) Display Power Panel

In addition to the circuit breaker panel, power may be removed fromthe display system by the display power panel located in theoverhead on the pilot’s side of the cockpit. This panel is shown inFigure 202.

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r-——— ———— ———— -——— ———— ———— ——-. ———- ____

1IIIIIIIII1II1IIII1

PILOT

L -—-— ———— —,——.

DI SPLAYSEICAS

OFF

COPILOT

OFF

IIIIIIIII

!IIIIII1

.—— —— ———— ———— ———— ———— ____ j

Display Power PanelFigure 202

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2. c. Primary Flight Display (PFD)

The display components which make up the PFD are discussed in detail in

the following paragraphs, but in general the PFD is comprised of thefollowing functions:

●

●

●

●

●

●

●

●

●

●

●

●

●

●

●

●

●

●

●

●

●

●

●

●

●

●

●

Attitude Sphere with Linear Pitch and Roll ScalesAttitude Source AnnunciationFlight Director Command Bars - SC/CPFlight Director Mode AnnunciationsAutopilot/Autothrottle Engage AnnunciationTCS Mode AnnunciationAutothrottle Mode AnnunciationsRadio Altitude Rising RunwayRadio Altitude Rising Brown Raster On Altitude TapeRadio Altitude Digital ReadoutRadio Altitude Source AnnunciationRadio Altitude Set Digital ReadoutVertical Nav Scale (GS/VNAV)Lateral Nav Scale

Airspeed Display Presentation

- Moving Analog Airspeed Tape with Vspeeds- Airspeed Target Bug with Digital Readout- Fixed Airspeed Current Value Pointer with Rolling Digits- V~O Thermometer on Airspeed Tape- Digital Mach Readout- Amber VSPD Annunciation

Mach Display Presentation

- Moving Analog Mach Tape- Mach Target Bug with Digital Readout- Fixed Mach Current Value Pointer with Rolling Digits- M~O Thermometer on Mach Tape- Digital Airspeed Readout

Moving Analog Altitude Tape with Trend VectorBaro Set Function - inHg/millibarsFixed Altitude Current Value Pointer with Rolling DigitsPositive Altitude Indication Below 10,000 FeetAltitude Select Bug with Digital ReadoutPositive Alt Sel Indication Below 10,000 FeetAltitude Alerting IndicationDigital Readout of Altitude in MetersVNAV Target Alt Bug with Digital ReadoutVNAV Altitude Alerting IndicatorDADC Source Annunciation

22-14-00Page 206


●

●

●

●

●

●

●

●

180-Degree Compass RoseHeading Source AnnunciationMagnetic/True AnnunciatorHeading Bug with Digital ReadoutNavigation Source AnnunciationNavigation Source Digital Distance ReadoutCourse Pointer/CDI with Digital ReadoutTo/From IndicatorBearing Pointers Circle/DiamondBearing Pointer Source AnnunciationMarker Beacon AnnunciationFixed Vertical Speed Tape with Moving PointerVertical Speed Current Value Digital Readout within Pointer (CurrentValue Range = t9900)Vertical Speed Target BugFixed Normalized AOA Scale with Moving PointerAOA Current Value Digital Readout within PointerAOA Target BugAbnormal AOA Range Indication

Comparison Monitor Annunciators

- Attitude- Heading- Air Data- ILS/MLS Data- EICAS Wraparound- Fault Warning Computer

Category II ILS Mode AnnunciationsCategory II ILS Excessive Deviation Monitor

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Honeywell !i!)!!r.c’2. c. (1) Attitude Displays

(a) Attitude Sphere - Moves with respect to the aircraft symboldisplay actual pitch and roll attitude.

(b) Pitch Attitude - The pitch attitude scale has white scalereference marks above and below the horizon line and markedfollows:

~ Down

;: ;:30 3040 4560 6090 90

to

as

There are reference markings every 2.5 degrees up to ten (10)degrees and every five (5) degrees between 10 and 30 degrees.Additionally, there are red warning chevrons which appear at 45and 65 degrees pitch up and 35, 50, and 65 degrees pitch down.

(c) Roll Attitude - The roll attitude scale displays actual rollattitude through a movable white index pointer and fixed whitescale reference markings at O, 10, 20, 30, 45, and 60 degrees.The 30 degrees reference mark is highlighted and shows up as abrighter shade of white.

(d) Aircraft Symbol - The stationary yellow aircraft symbol is usedto develop the relationship between aircraft pitch and rollattitudes and the movable sphere. Additionally, the symbol isused to align to the flight director command cue in order tosatisfy the commands of the selected flight director mode.

(e) Attitude Source Annunciation - The selected attitude source isnot annunciated if it is the normal on-side source for thatindicator. As other attitude sources are selected they areannunciated in white left of the attitude sphere. When thepilot and copilot source are the same, they are annunciated onboth displays in amber. Source selection for attitude isaccomplished through the display controller as shown in Figure203.

22-14-00Page 208


2. c. (1) (f)

(9)

(h)

(i)

(j)

Flight Director Command Cue(s) - The magenta flight directorcommand cues are superimposed over the attitude display and canbe used as reference commands to capture and maintain a desiredflightpath. The command cue format single cue (SC) or crosspointers (CP) is selected by the crew through the displaycontroller as shown in Figure 204. The SC combines the pitchand roll commands into one cue presentation and is flown byaligning the fixed aircraft symbol with the cue. The CP hasthe pitch and roll commands separated into two cuepresentations and is flown by aligning the aircraft symbol withthe center of each command cue.

Flight Director Autothrottle Modes - Flight director (FD)vertical and lateral modes are annunciated along the top of theattitude sphere as shown in Figure 203. The lateral moderegion appears to the upper left of the attitude sphere. Thevertical mode region appears to the upper right of the attitudesphere. The center region is reserved for autothrottle (AT)mode annunciations as shown in Figure 204. Armed modes areannunciated in white and located in the left half of theappropriate region. Captured modes are annunciated in greenand located in the right half of the appropriate region. Asthe modes transition from armed to captured, a green box isdrawn around the capture mode for 5 seconds. Table 201 lists

all the annunciated lateral and vertical FD modes, as well asautothrottle modes.

Autopilot (AP)/Autothrottle (AT) Engage Annunciation - The APand AT engage status is displayed adjacent to the right rollindices below the FD lateral mode box as shown in Figure 205.

Touch Control Steering (TCS) Mode Annunciation - TCS isannunciated in place of the AP engage annunciation whenever TCSis selected on either the pilot’s or copilot’s control wheel.The annunciation is normally white except when CAT II ILSmonitors are active in which case the annunciation is amber.

Radio Altitude Rising Runway - Theprovides an indication of absoluteThe symbol appears at 200 feet andat touchdown. The runway symbol g“corresponding digital readout.

yellow rising runway symbolaltitude above the terrain.contacts the aircraft symbolves analog cues to its

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Aug-15/91Use or disclosure of information on this page is subject to the restrictions on the title page of this document.

v

CRossPOINTERS

ALTITUDE-TRENDVECTOR

Vs&kRf000 ml

RAD

‘SOtiR~E

Primary Flight Display Format(SENSOR)

Figure 203 22-14-00Page 210


sw7-

tlACH_NUMBER

T D

-.55M 293;~SI;

CRSm

NAV1 9l.o– 000 \\u 1!/,

* ~l$f’

23. 4DME.< ?//

-~

;.8– \\ /

.6— 4 .~ Q%%

0

0 ‘;A:

o-

.4—==00

t

1AOA

.20>HDG

SCALE/’:

0 m ‘OADFf000 0 VOR2

;

Primary Flight Display Format(DISP)

Figure 204

ALTITUDE

\

\

B;;:

METR1 CALTITUDE

,_VEFF&\AL

BEAR INGPOINTERS

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FD Lateral Mode

Heading SelectLNAV FMSLNAV VORApproach ILSBack Course

I MLS

I

FD Vertical Mode

Altitude HoldVertical SpeedAltitude PreselectFLCH IASFLCH MACHApproach ILSTakeoffGo-AroundVNAV ALT PreselectVNAV ALT HoldVNAV FLCH IASVNAV FLCH MACHVNAV PathVNAV ARMMLS

AT Mode

TakeoffGo-AroundFlight Level ChangeSpeed - IASSpeed - MachHold ModeVHO LimitedMHO LimitedPower LimitedFlap LimitedGear Limited

Mode Annunciation

HDGFMSVORLOCBCAZ

Mode Annunciation

ALTVsASELIASMACH

!;GAVASELVALTVIASVMACHVPATHVNAVGP

Mode Annunciation ~

TO xGA xFLCHIAS :MACH xHOLDVMOMMOPOWERFLAPSGEAR

Qy

x

;xxx

~

xxxxx

;xxxxxx

x

~

xxxxxxxxxxx

Flight Director/AutothrottleMode AnnunciationsTable 201

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2. C. (1) (k)

(1)

(m)

(n)

Radio Altitude Rising Brown Raster Display - Another indicationof radio altitude is given on the barometric altitude tape (seeFigure 205). At 600 feet AGL, a rising brown raster fills thebackground of the altitude tape displacing the normal grayraster field. The bottom of the altitude scale corresponds to600 feet and the baro altitude present value window correspondsto O feet (Figure 206). The brown raster fills in the scaleproportionally between 600 and O feet AGL.

Radio Altitude Digital Readout - The white radio altdigital readout and RA label is located to the lowerthe attitude sphere. The digital readout comes into2500 feet and is displayed in 10-foot increments abo~and 5-foot increments below 200 feet.

tuderight ofview ate 200 feet

Radio Altitude Source Annunciation - The selected radioaltitude source is not annunciated if it is the normal on-sidesource for that indicator. As other sources are selected, awhite side designator (1 or 2) is added to the RA legend. Whenthe pilot’s and copilot’s source are the same, it isannunciated on both displays in amber. Source selection forradio altitude is accomplished through the display controlleras shown in Figure 203.

Radio Altitude Reference Set Display - The crew can set agreen radio altitude reference display directly above thedigital radio altitude display. When the actual value ofradio altitude is less than the reference value, the actualradio altitude value is boxed in white. The box flashes for5 seconds after the transition has occurred. The referencedisplay is set using the display controller asshown inFigure 205. The reference value is settable to 2500 feet in10-foot increments above 200 feet and 5-foot increments below200 feet.

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Honeywell !!$!!g~.c’

CATEGORY I I

AP/AT ENGAGEILS MODE

ANNUNCIATORSANNUNCIATOR

\VSPEEDS \RAD ALTSET DATAs I ~%%

CASTARGET

OFF-SCALE

\

RISINGRUNWAY

AOATARGET

\

Jet-l ~ I %+2////7

+

*RS1.0 000

.8

0.59

1

.4

.2HDG000

\\\ \’\\ +

\\.~ Q\f?l

/11

k=75

Primary Flight Display Format(FLT REF)Figure 205

BELOW-1:E:go

R;::::

RASTER

BELOWRA:E$LT

- RADIOALTITUDE

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“ 220

. 23M

1.0

.8

.6i

I HDG I

L E

10=10

\ /w

TO I

.

CRS000

HDG000

Primary Flight Display Format(AUTO VSPD - CONFIG Mismatch)

Figure 20622-14-00

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2. c. (1) (o) Vertical Navigation Scale - The vertical navigation scaleappears to the right of the attitude sphere. Glideslo~;e(G)and VNAV (V) deviations are displayed on this scale.source selection for this scale is done through the displaycontroller. The selected source type is labeled with theletters shown in parenthesis of the above listed source types.The VNAV scale is enabled for display only when a VNAV flightdirector mode is active (see Figure 208). The verticalnavigation scale has a fixed white (G)/cyan (V) deviation scaleshowing t2 dots of deviation. The green (G)/cyan (V) deviationpointer moves along this scale to show deviation from pathcenter. Aircraft displacement from path is indicated by therelationship of the aircraft to the deviation pointer. Thescaling definition for glideslope, VNAV, and VNAV approach areas follows:

GS/GP 75 pA/dotVNAV 250 ft/dotVNAV APP 75 ft/dot

(2) Calibrated Airspeed/Mach $cale

The CAS/Mach scale presentation is located to the left of theattitude sphere as in a conventional “T” arrangement. The CAS scaleis presented for altitude less than 25,000 feet (Figure 204). Thedisplay can be selected to a Mach scale above 25,000 feet (Figure203) via the display controller or through the flight guidancesystem. Each display is configured to be a moving tape with fixedcurrent value pointer. The tape markings, current value data, andcurrent value pointer are white with gray background shading fordisplay enhancement.

(a) Calibrated Airspeed Scale (CAS)

~ Calibrated Airspeed Analog Scale - The white scale markingson the tape are shown at 10-knot intervals. The markingsare labeled every 10 knots below 200 knots (Figure 204) andare labeled every 20 knots above 200 knots.

~ Calibrated Airspeed Digital Readout - A digital readout ofthe actual airspeed is magnified and displayed within thecurrent value pointer. The digital data is displayed inwhite using a digital rolling drum presentation readable toa 1 knot resolution.

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2. c. (2) (a) 2 V Vector - The V~O“$in oard side of the


vector is a fixed red bar located on theCAS scale, originating from the CAS

scale endpoint extending out to VMO. The V~o vectorsymbology is identical to the MM vector shown in Figure203. When the value in the poln%er exceeds VN~MHO, thedigital CAS data, in the current value pointer, IS turnedred.

Speed Target Digital Readout - The speed target digital datais displayed in green at the top of the CAS scale. Thespeed target data is set by the pilot using the guidancepanel or is available automatically with the FMS.

Speed Target Bug - The green triangular speed target bugtravels along the inboard side of the airspeed tape. Thebug position on the scale corresponds to the digital speedtarget value displayed at the top of the CAS scale.

Vspeed Bugs - In addition to the speed target bug describedabove, the pilot has the ability to set Vspeed bugscorresponding to reference speeds for various phases offlight. This data is available from the FMS PZ-800Performance Computer automatically or is manually setthrough the display controller as shown in Figure 205. TheVspeeds travel along the airspeed tape in the same area asthe speed target bug previously described. The Vspeed bugsare displayed in green.

The default in the display controller is AUTO VSPD selected.In the event that a performance computed configurationmismatch is detected, prior to takeoff (aircraft on ground),the performance computed Vspeeds are displayed on the DCunboxed with an asterisk. Additionally, an amber VSPDannunciation is displayed in the top portion of the PFDairspeed tape as shown in Figure 206. The VSPD annunciationis also active with a DC failure on the ground.

The source for Vspeed information is the on-side displaycontroller.

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2. c. (2) (a) z Digital Mach Readout - A Mach number digital readout isdisplayed below the speed scale whenever the CAS tapeconfiguration is active. The data is labeled with a white“M” to designate Mach. The digital readout is normallywhite. When V~~M~O is exceeded, the Mach digits turn red.

(b) Mach Scale (MACH)

Mach Number Analog Scale - The white scale markings on thetape are shown at 0.02 Mach intervals. The markings arelabeled in white every 0.02 Mach interval.

Mach Number Digital Readout - A digital readout of theactual Mach number is magnified and displayed within thecurrent value pointer. The digital data is displayed inwhite using a digital rolling drum presentation readable toa 0.01 Mach resolution.

M Vector - The M~O vector is a fixed red bar located on the‘%in oard side of the Mach scale, originating from the Machscale endpoint extending out to MO (Figure 203). When thevalue in the pointer exceeds VH~~O, the digital Mach data,in the current value pointer, 1s ~urned red.

Mach Target Digital Readout - The speed target digital datais displayed in green at the top of the Mach scale. Thespeed target data is set by the pilot using the flightguidance controller, or it is available automatically withthe FMS.

Mach Target Bug - The green triangular Mach target bugtravels along the inboard side of the Mach tape. The bugposition on the scale corresponds to the digital Mach targetvalue displayed at the top of the Mach scale.

Digital CAS Readout - A CAS digital readout is displayedbelow the speed scale whenever the Mach tape configurationis active. The data is labeled with a white KTS todesignate knots. The digital readout is normally white.When v~O/M~Ois exceeded, the CAS digits turn red.

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2. C. (3) Barometric Altitude Scale

(a)

(b)

(c)

(d)

(e)

(f)

Altitude Analog Scale - The barometric altitude scale is amoving tape display with fixed current value pointer. Thewhite scale markings on the tape are in 100-foot intervals.The scale is labeled every 200 feet below 10,000 feet (Figure205) and every 500 feet above 10,000 feet (Figure 203). Thetape markings, current value data, and current value pointerare white with gray background shading for display enhancement.

Altitude Digital Readout - A digital readout of the currentaltitude is magnified and displayed within the current valuepointer. The current data is displayed in white using adigital rolling drum presentation readable to a 20-footresolution. For climb/descent rates greater than 1000 feet perminute, the rolling drum digits are replaced by two dashes toenhance altitude scale readability (Figure 203). Below 10,000feet, an indication for absence of the most significant digitis given (Figure 205).

Altitude Trend Vector - The magenta altitude trend vectororiginates at the tip of the current value pointer and travelsalong the inboard side of the altitude scale (Figure 203). Thetrend vector is an altitude predictor term with a lengthequivalent to 6 seconds.

Altitude Select (ALT SEL) Digital Readout - ALT SEL digitaldata is displayed in green at the top of the altitude scale.The ALTSEL target data is set by the pilot using the flightguidance controller.

Altitude Select Bug - The green triangular ALT SEL bug travelsalong the inboard side of the altitude scale. The bug positionon the scale corresponds to the ALT SEL digital target valuepreviously discussed. The bug is present on the scale at alltimes with the exception of glideslope (GS) capture modeactive. Glideslope capture has no effect on the digital ALTSEL display.

Altitude Alert Indications - Approaching the ALT SEL reference,the ALT SEL bug and readout turn from green to amber wheneveractual altitude is within 1000 feet of the ALT SEL referenceindicators. These displays remain amber until the actual valueis within 250 feet at which point they turn back to green.Departing a selected altitude, the ALT SEL bug and readout turnfrom green to amber whenever the current altitude value isgreater than 250 feet from the ALT SEL reference indicators.If the excursion from altitude exceeds 1000 feet, the bug andreadout return to their normal green color.

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2. c. (3) (g)

(h)

VNAV Altitude Reference Bug/Readout - The VNAV altitude targetdigital readout is displayed in cyan below the FD vertical modewindow (Figure 208). This target is active whenever an FD VNAVmode is active and the target is different than the altitudepreselect value. A cyan pointer shaped bug travels along theinboard side of the altitude scale. The bug position on thescale corresponds to the VNAV altitude target digital readout.This data is transmitted automatically by the FMS when VNAV isactive.

VNAV Altitude Alert Indications - Approaching the VNAV ALTtarget, the VNAV ALT bug and readout turn from cyan to amberwhenever actual altitude is within 1000 feet of the ALT SELreference indicators. These displays remain amber until theactual value is within 250 feet at which point they turn backto cyan. Departing a selected altitude, the VNAV ALT SEL bugand readout turn from cyan to amber whenever the currentaltitude value is greater than 250 feet from the VNAV ALT SELreference indicators. If the excursion from altitude exceeds1000 feet, the bug and readout return to their normal cyancolor.

(i) Barometric Altimeter Setting - The magenta baro set digitalreadout is located directly below the altitude tape. Eachpilot can set the barometric reference for the on-side DADCthrough the baro set knob located on the display controller.The pilot has the ability to select the altimeter setting fordisplay in either inHg (IN) or millibars (MB) as selected onthe display controller (Figure 204).

(j) Metric Altitude Display - Directly below the baro set dataappears a white digital readout of altitude in meters. Thisdisplay appears only if it has been selected on the displaycontroller (Figure 204).

(k) Digital Air Data (DADC) Source Annunciation - The selected DADCsource is not annunciated if it is the normal on-side sourcefor that indicator. As DADC sources are selected they areannunciated in white to the right of the attitude sphere. Whenthe pilot and copilot source are the same, they are annunciatedon both displays in amber. Source selection for DADC isaccomplished through the display controller as shown inFigure 203.

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2. c. (4) Horizontal Situation Indicator (HSI)

(a) Aircraft Symbol - The white aircraft symbol provides a quickvisual cue as to the actual aircraft position in relation toheading, selected heading, or selected course.

(b) Heading Display - The displayed compass rose consists of a180-degree arc that is divided into 5-degree increments andlabeled every 30 degrees. The compass rose rotates about thestationary aircraft symbol to provide actual headinginformation. The compass indices and labels are displayed inwhite.

(c) Heading Source Annunciation - Displayed heading comes from theselected IRS.

(d) Heading MAG/TRU Annunciations - The MAG/TRU reference label isnot displayed if the selected source is the normal on-side MAGcondition. For on-side true data a white TRU is displayedabove the lubberline. Cross-side data is displayed in white asMAG or TRU with a side designator 1, 2, or 3 attached (Figure203). When the pilot’s and copilot’s source are the same, theyare annunciated on both displays in amber.

(e) Heading Bug with Readout - The notched magenta heading bug ispositioned around the rotating heading dial by the headingselect knob located on the flight guidance controller. Adigital heading select readout of bug position is provided andlocated at the lower left-hand side of the compass. This datais labeled with a white HDG. The digital readout is displayedin magenta corresponding to the heading bug color. The centerof the heading select knob contains a button labeled SYNC.Selection of the SYNC button causes the heading bug to alignwith the current aircraft heading.

(f) Course Deviation, Course Pointer with Readout - The coursedeviation bar represents the centerline of the selectednavigation or localizer course. The aircraft symbolpictorially shows aircraft position in relation to thedisplayed deviation.

The course pointer is positioned around the rotating headingdial in one of two manners:

● Course select knob located on the flight guidance controllerfor short range NAV (SRN) sources (i.e., NAV, ILS, MLS).

. Long range NAV (LRN) sources automatically select thedesired track data for the active flight plan (i.e., FMS,LASERTRAK’”).

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Honeywell %!!k!f.c’The digital readout corresponding to the pointer position isprovided and located to the upper left of the compass arc. Thedata is labeled CRS forSRNs and DTRK for FMS. The pointer,deviation indicator, and digital readouts are green for SRNsources and cyan for FMS. The center of the course select knob

contains a button labeled SYNC. Selection of the SYNC buttoncauses the course pointer to align with the bearing of theselected NAV source causing the deviation bar to center. TheSYNC mode is only operational when the displayed navigationsource is VOR.

The scaling for course deviation for each displayed parameteris as follows:

SRNS 75 flldotFMS 2.5 miles/dotFMS AW 0.75 miles/dot

2. C. (4) (g) Distance Display - The distance display indicates the nauticalmiles to the selected DME station or FMS waypoint. Thedistance display is located to the upper right of the compassarc. The data is labeled DME (Figure 207) for SRN sources andNM (Figure 203) for FMS. The digital display color isconsistent with the course select pointer (i.e., green for SRNSand cyan for FMS). DME hold is indicated by displaying anamber “H” adjacent to the distance readout (Figure 207).

(h) NAV Source Annunciations - Annunciation of the navigationsource is displayed in the upper right-hand corner of the

(i)

compass arcthe abilitygiven insta-Figure 207.the pilot’sannunciated

TO/FROM Indof the HSI.

directly above the distance readout. The crew hasto select all navigation sources available in anylation through the display controller as shown inThe labels are normally displayed in white. Whenand copilot’s source are the same, they areon both displays in amber.

cater - A magenta arrow head appears in the centeroriainatinq at the fixed aircraft symbol nose or

tail as appropriate. The arrow indicates whether the selectedcourse will take the aircraft TO or FROM the station orwaypoint. The TO/FROM annunciator is not in view duringlocalizer operation.

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2. c. (4) (j) Bearing Pointers with Annunciation - Each pilot can display up

to two bearing pointers on the compass arc simultaneously. Thebearing pointers are distinguishable by shape (circle ordiamond) and color (white or yellow). Each bearing pointersource is labeled and displayed in the lower right-hand cornerof the HSI. The circle pointer is always displayedpositionally above the diamond pointer. Bearing pointersources are selected in the MAP (Figure 212) and COMP(Figure 217) modes of the display controller when the BRGselection in the DISP mode has been made (Figure 204).

(k) Marker Beacon Annunciation - Marker beacon information isdisplayed inside the lower right-hand side of the attitudesphere (Figure 207). The marker colors are blue for outer (0),amber for middle (M), and white for inner (I). A boxidentifies the location of the marker beacon annunciationswhenever a localizer has been selected for display.

(1) LASERTRAKW Option - LASERTRAKW course (LRN) and distanceinformationmay be displayed on the PFD HSI when selected onthe display controller. HSI displays function as previouslydescribed exce~tthat the fliqht quidance computer cannot be--coupled to this data.

(5) Vertical Speed Display (VS)

(a) VS Analog Scale - The VS tape is a fixedmoving current value pointer. The whitedisplay appear at O, t500, t1000, t2000,minute (FPM). The scale labels O, 1, 2,

white scale withscale markings on theand t3000 feet perand 3 appear at the O,

t1000, jZOOO, and t3000 FPM marks respectively. The rangebetween *1OOO FPM is expanded and marked with a short tickmarkat the t500 FPM point. A white zero rate of climb index markis also provided on this display.

(b) VS Digital Readout - A digital readout of the actual verticalspeed is magnified and displayed in the current value pointer.The digital data is displayed in white and is readable to a 50FPM resolution (Figure 207) below t1000 FPM and 100 FPMresolution (Figure 208) above t1000 FPM.

(c) VS Target Bug - The green triangular VS target bug travelsalong the inboard side of the VS scale. The bug position onthe scale corresponds to the pilot selected reference bug onthe flight guidance controller.

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Honeywell !!!#!#b.cE

Fi:M

SELECTEIHEADING

Primary Flight Display Format(NAV)

Figure 207

22-14-00Page 224


VNAVALTITUDE/TARGET

I

29.921N

:=

DTRKT

ms 1 31.0 000

%V

\\u’ ‘/’//

~~

23.4 NM 2.8 %/ t

.<.6 .~ Q ;

e: 01%

A.4

liDG=> o 0

t

II ;50.20 0 ‘–OADF1 s

325 / \\ ~ VOR2

VNAV

Primary Flight Display Format(VNAV)

Figure 208

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Honeywell ~f~~~~fE

2. c. (6) Normal ized Angle Of Attack (AOA) Display

(a)

(b)

(c)

AOA Analog Scale - The AOA tape is a fixed scale with movingcurrent value pointer. The white scale markings on the displayrange from 0.2 to 1.1 in 0.1 AOA increments. The scale islabeled in white at 0.2 AOA intervals. The scale is color-

banded to show AOA exceedance ranges. The scale is whitebetween O and 0.64 AOA, amber between 0.64 and 0.80 AOA, andred between 0.80 and 1.1 AOA.

AOA Digital Readout - A digital readout of the actualnormalized AOA is magnified and presented in the current valuepointer. The digital data is readable to a 0.01 AOAresolution. The digit colors change to match the indicated

range on the scale as defined above.

AOA Target Bug - The green triangular shaped AOA target bug

moves along the inboard side of the VS scale. The target valueis settable by the pilot through the use of the displa~controller (Figure 205).

(7) Comparison Monitor Annunciation

Selected pilot and copilot input data is compared by the symbolgenerator. If the difference between the data exceeds predeterminedlevels, the out-of-tolerance symbol is displayed. Theseannunciations are displayed in amber and are located to the left,adjacent to the attitude sphere. A list of the compared signals anddisplayed cautionary symbols are given below.

Com~ared Sictnals Displayed S~mbol Tri~ Threshold

Pitch AttitudeRoll AttitudeHeadingAirspeedAltitudeLocalizer

Glideslope

Fault Warning ComputerEngine DisplayCAS Display

PIT (ATT*)ROL (ATT*)HDGIASALTLOC (ILS**)

(AZ***)GS (ILS**)

(GP***)FWCENGDU4

6 degrees6 degrees6 degrees20 knots200 feet38 PA

49 pll

Red Messages

* PIT and ROL active produce a ATT symbol.

** LOC and GS active produce a ILS symbol.*** When MLS is selected EL and Az symbols are displayed. EL and

AZ active produce an MLS symbol.

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2. c. (8) Category II (CAT 2) ILS Operation

CAT1/CAT2 annunciations and ILS deviation monitors have been addedto the PFD display format for aircraft operators receiving CAT2certification (CAT2 wiring installed).

(a) CAT1/CAT2 Annunciations - The CAT2 annunciation is displayedautomatically when the APR mode is selected on the flightguidance controller and lateral and vertical deviation validwith the displays configured as follows:

● FWC JIA-82 (CATII Installed) will be read by the FWC at

. ~fle~L~pset < 200 feet on at least one side● One radio altimeter (minimum) valid● Two symbol generators (minimum) valid● On-side ILS (NAV/ILS) source selected for display on each

side and valid. Either dual couple mode or radio altitude greater than

800 feet (Dual couple mode occurs at approximately1200 feet)

CAT2 is annunciated in green directly below the FD verticalmode window (Figure 205) when all the conditions listed aboveare met. In the event that one of the enabling conditionschanges, the green CAT2 is replaced with an amber CAT1. CAT1appears flashing for 5 seconds and then remains displayedsteady. Changes back to CAT2 are handled the same way in theevent that the enabling condition is restored. The CAT2annunciation is presented only if the displays are configuredas above. If these conditions are not met when APRmode isselected, no CAT1/2 mode annunciations are displayed.

(b) Excessive Deviation Monitor - When the CAT2 mode annunciationis active, and displayed radio altitude is less than 500 feet,the excessive deviation monitor is activated. When thethreshold of the localizer (LOC) or glideslope (GS) deviationsare exceeded, the deviation pointer and deviation scale areturned amber and are flashed. This condition remains as longas the deviation is above the appropriate threshold. Thethreshold for LOC and GS are as follows:

Localizer 320 /LA Glideslope ~65 PA

Excessive deviation indications are also shown on the NDcompass mode when the CAT2 mode is active. CAT2 will becancelled:

● at touchdown● when there is a change in the selected lateral FD mode● when there is a change in the selected vertical FD mode● when approach mode is deselected.

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2. C. (9) PFD Caution and Warning Displays

(a) Inertial Reference System (IRS) Failures and Flags - Thedisplay symbology associated with IRS failures is shown inFigure 209.

~ Failure of either the pitch or roll data is indicated byremoving the pitch scale markings, turning the entireattitude sphere cyan, and displaying, in red, the label ATTFAIL in the top center of the attitude sphere.

~ Failure of the displayed heading is shown by removing thedisplayed bearing pointers, course pointer and deviation,and displaying, in red, the label HDG FAIL at the top centerof the compass rose.

(b) IRS Test mode is shown by displaying the labels ATT TEST andHDG TEST in the same location as the fail annunciationsdiscussed above (see Figure 211).

(c) Flight Director Failure - In the event of a flight directorfailure, an amber FD annunciation is displayed in the left-most(lateral modes) FGC mode annunciation box. Additionally, theflight director mode annunciations and cue are removed (Figure210).

(d) Autothrottle Failure - In the event of failure in the selectedautothrottle, an amber AT annunciation is displayed in the(center) autothrottle mode annunciation box (Figure 210).

(e) Radio Altitude Failure - In the event of a failure of the radioaltimeter, amber dashes will replace the digital radio altitudevalue and the rising runway/brown raster will be removed fromthe display if present (Figure 210).

(f) Vertical Navigation Scale - A failure of the source driving thevertical navigation scale (SRN only) is shown by removing thedeviation pointer and displaying a red “X” through thedeviation dots as shown in Figure 210. When VNAV is invalid,the VNAV vertical deviation scale is inhibited for display.

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I

—-- M

!1.0

.

.

.4

.2

I

FAIL

I

0

0

al-0

---

-

=xHDG---

/

m=

3

K

21

0

123

PFD Failure Indications (IRS/DADC)Figure 209

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I FD I AT I1 v

20 — 20

lo— 10

to= 10~ 1

200----

. 55f’1

PFD Failure Indications (Mist)Figure 210

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2. c. (9) (g) Digital Air Data (DADC) Failures) - In the event of a failureof the displayed DADC, CAS or Mach, altitude, vertical speed,and AOA, the failure indication is as shown in Figure 209.

~ In the case of the CAS (Mach) and altitude scales, data isremoved from the current value pointer, the scale markingsare removed, and a red “X” is drawn through the scale. Thedigital Mach (CAS) display failure is shown by replacing thenumerical value with amber dashes.

~ In the case of the vertical speed and AOA scales, thecurrent value pointer is removed and a red “X” is drawnthrough the scale.

(h) Heading Select Failure - Failure of the heading select signalswill cause the digital display to be replaced by amber dasheswith the heading bug removed from the display. This indicationwill also be given in the event of an invalid heading display(Figure 209).

(i) Course Select Failure - Failure of the course select signalswill cause the display to be replaced by amber dashes with the

course pointer removed from the display. This indication willalso be given in the event of an invalid heading display or FMSsource (Figure 209).

(j) Distance Display Failures - Failure of either the DME or FMSdistance signals is indicated by replacing the digital distancevalue with amber dashes (Figure 211).

(k) Course Deviation Failure - A failure of the course deviationindicator is shown by removing the deviation bar and displayinga red “X” through the deviation dots (Figure 211).

(1) EFIS Self-Test - This test can be initiated through the on-sidedisplay controller (TEST mode) when airspeed is valid and lessthan 60 knots and the WOW switch is active. Selecting the EFISself-test shows the invalid flags for the following PFD displayinformation:

. IRS

. DADC● Flight Director. All Comparison Monitor Annunciations Active. Radio Altitude. Distance Display● Vertical Deviation Display

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.55M

1.4

i

DTRK---

1 I

O RA

FMS 1 r3

F21

123

IRS Test Mode IndicationsFigure 211

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2. D. Navigation Display (ND) Formats

The navigation display is comprised of the following three display modes:

c Map Mode● Compass (COMP) Mode. Plan Mode

The above modes are selected by the crew via the on-side displaycontroller. The following paragraphs contain specific descriptions ofeach NAV display mode.

(1) Navigation Display - General

(a) Static AirTemp (SAT), True Airspeed (TAS), and Groundspeed(GSPD) Displays - The SAT, TAS, and GSPD digital readoutsappear in all three of the above listed NAV display formats.The data is positionally stable in the MAP, COMP, and PLANmodes. This data is displayed in the extreme lower right-handcorner of the ND. In each case, the data is labeled in dimwhite and the actual value is displayed in bright white. Datafor SAT and TAS comes from the selected DADC source and GSPDcomes from the selected ND FMS source.

(b) Weather Radar (WX) Mode Annunciations - Weather radar modeannunciations appear in the MAP, COMP, and PLAN modes. Thedata is positionally stable in all three ND modes. This datais displayed in the extreme lower left corner of the ND.Available modes for the P-800 Weather Radar Svstem are asfollows:

WX (Amber) -STBY (Green).-WX (Green) -

GMAP (Green) -TEST (Green) -FAIL (Amber) -RCT (Green) -GCR (Amber) -TGT (Green) -TGT (Amber) -VAR (Amber) -WX (Flashing

WXR Off/WXR FailStandby ModeWeather ModeGround Map ModeTest ModeTest Mode Failed AnnunciationREACT ModeGround Clutter Reduction ModeTarget Alert OnTarget Alert ActiveVariable Gain ModeAmber) - ND is not in PLAN mode to display

Test, Weather, or Ground Map

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2. D. (1) (c)

(d)

Honeywell ##&f~.cE

Available modes for the P-870 Weather Radar System are asfollows:

WX-OFF (Green) -WX (Green) -WX (Amber) -

WX/T (Green) -GMAP (Green) -TEST (Green) -FAIL (Amber) -RCT (Green) -GCR (Green) -TGT (Green) -TGT (Amber) -STBY (Green) -FSBY (Amber) -WAIT (Amber) -STAB (Amber) -VAR (Amber) -

W (Flashing

Weather Radar Tiltdisplayed in greenannunciation box.

P-870 RT Turned OffWeather ModeWXR Interface FailWeather Mode with TurbulenceGround Map ModeTest ModeWXR Failed AnnunciationREACT ModeGround Clutter Reduction ModeTarget Alert OnTarget Alert ActiveStandby ModeForced Standby ModeRT Warmup ModeStabilization Off IndicationVariable Gain ModeAmber) - ND is not in PLAN mode to display

Test, Weather, or Ground Map

Display - Weather radar tilt information ison the bottom line of the WX modeThe information is displayed in half-decwee

increments for values between tlOO and one-d~gree incrementsfor values greater than tlOO. The P-870 autotilt mode isindicated by an “A” following the digital value.

Weather Radar Fault Codes - When the amber FAIL annunciation isdisplayed (P-870 installationsonly) selection of the TEST modecauses P-870 fault codes to be displayed (amber) in place ofthe digital tilt data.

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2. D. (2) Map (MAP) Mode

The map mode is comprised of the follow.

. 120-Degree Compass Rose● Digital Heading Readout. Magnetic/True Annunciation. Heading Bug with Digital Readout. Actual Track Line. Wind Display in Vector or X-Y Formats● FMS 1/2 Source Annunciation● Bearing Pointers Circle/Diamond. Bearina Pointer Source Annunciation

ng funct”ons:

. Flight-Management System (FMS) Mode Annunciations

. FMS Waypoint, Airport, and Navaid Display Provisions

. FMS Desired Track Lines

. Map/Weather Radar Range Rings

. Joystick Slewable Waypoints● Vertical Profile Presentation. Weather Radar Display

(a)

(b)

(c)

Aircraft Symbol - The white aircraft symbol provides a visualcue to aircraft position relative to actual heading andselected heading.

Actual Track Vector - Actual track from the IRS is displayedfrom the nose of the aircraft symbol to the map lubber line asa magenta, dashed line. The selected IRS source provides thisdata (Figure 212).

Heading Display - The displayed compass rose consists of anexpanded 120-degree arc that is divided into 5-degreeincrements and labeled every 30 degrees. The compass roserotates about the stationary aircraft symbol to provide actualheading information. The compass indices and labels aredisplayed in white. The display is augmented with a whitedigital readout of actual heading above the compass lubber-line(Figure 212).

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HEADING 1WAY~~INT MAG/TRU FMsM::

WAYPOINT= 7\ \

Y00 TRU FMS1

SELECTEIHEAD ING

\

WIND-

M%%EBOX

\

iHDG315

1

-WxVARGCR

o ADF 1@OR2

r

SAT ‘- 56

TAS234

GSPD345

FMS~ ALERTMESSAGES

BEAR ING‘POINTERS

HALF RANGELABEL/DATA

-BEARINGSOURCES

Map Mode FormatFigure 212

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2. D. (2) (d)

(e)

(f)

(9)

Heading MAG/TRU Annunciations - The heading source for thisdisplay is the same IRS source which drives the on-side PFD.The MAG/TRU reference label is not displayed if the selectedsource is the normal on-side condition. For on-side true data,a white TRU is displayed to the right of the heading digitalreadout (Figure 212). Cross-side data is displayed in white asMAG or TRU with the side designator 1, 2, or 3 as appropriate.When the pilot and copilot source are the same, they areannunciated on both displays in amber.

Heading Bug with Readout - The notched magenta heading bug ispositioned around the rotating heading dial by the headingselect knob located on the flight guidance controller. Adigital heading select readout of bug position is provided andlocated at the top of the WX mode annunciator box (lower leftcorner). This data is labeled with a white HDG. The digitalreadout is displayed in magenta corresponding to the headingbug color. The center of the heading select knob contains abutton labeled SYNC. Selection of the SYNC button causes theheading bug to align with the current aircraft heading.

NAV Source Annunciations - Annunciation of which FMS is drivingthe displayed flight plan data is located in the upper right-hand corner of the ND. The pilot’s source is FMS1 when hisactive navigation source is any SRN or FMS1. When FMS2 isselected as the active NAV source, the map source also becomesFMS2. The copilot’s map source is FMS2 when his activenavigation source is any SRN or FMS2. When FMS1 is selected asthe active NAV source, the map source also becomes FMS1. Theseselections are made through the display controller as shown inFigure 207. When the pilot and copilot map sources are thesame, the source label is displayed on both NDs in amber.

Bearing Pointers with Annunciation - Each pilot can display upto two bearing pointers on the map arc simultaneously. Thebearing pointers are distinguishable by shape (circle ordiamond) and color (white or yellow). Each bearing pointersource is labeled and displayed in the lower right-hand cornerof the HSI. The circle pointer is always displayedpositionally above the diamond pointer. Bearing sourceselections are made using the display controller as shown inFigure 212.

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2. D. (2) (h) Wind Display - The wind format is displayed in white in eithera vector (Figure 212) or X-Y (Figure 213) component format asselected through the display controller. The wind displayabeams the aircraft symbol to the left, above the headingselect digital readout. This data is displayed from theselected ND FMS source.

(i) Range Rings - Range rings are displayed to aid in determiningthe position of radar returns and active flight planparameters. The range ring boundary is the compass card arcand represents the select range on the weather radarcontroller. A half-range ring (arc) is provided andlabeled with the half-range distance at each end of the arc(Figure 212).

(j) Weather - Weather information from the radar is displayed onthe map format, within the WX boundary (Figure 213.1), wheneverat least one weather radar (side independent) controller hasbeen turned on.

With both WX controllers ON (STBY, TEST, WX, or GMAP modesselected), each pilot has independent control over alternatesweeps of the radar yielding independent weather displaycapability relative to range, tilt, and WX mode. With one WXcontroller ON and the other in the OFF position, the activecontroller selects weather radar modes and range data on bothside map displays (slave-mode).

Activation of the TEST, WX, or GMAP modes, with the aircraft onthe ground, and the COMP or PLAN mode displayed on the ND,causes the mode annunciation to turn amber and flash.

With the MAP mode display on the ND and the TEST, WX, orGMAPmode active, selecting the FPLN range on the WX controllerremoves the displayed radar returns and flashes the activeradar mode annunciation in amber.

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NAG/TRUUIND

‘COMPONENTS0

looaMAG2i

YFMS 1

4+

HDG315 4

Ux TIUS / tlILEN SATVAR SAV12 FL41O - S6GCR TAS

234GSPD

34s

*NAVAID

J!A~#D

AIRPORTH ID

-AIRPORT

JEJ?EJ?IICXJL

Map Mode With Vertical ProfileFigure 213

22-14-00Page 239


6!%r5tE151 f‘mo FII!S1 VOR1 ADFt AUTO

?~ w

c BS* Ff’ls2 VOR2ADF2 3 0sm9cuFLw TmsYsTm0000 ID UAYPT

c ID NAVAID VERT PROF SET

Bfi E ID AIRPT WIND )(Y~ “o&

Mif%EBOX

FMS 1 ‘

UX DISPLAY-BOUNDARY

HALF RANGELABEL/DATA

Map Mode With Weather Radar DisplayFigure 213.1

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—— .- MAINTENANCE

2. D. (2) (k)

(1)

(m)

(n)

(o)

Active F1ight Plan Waypoints (WPTS) - The active flight plancan be displayed on the MAP whenever the WPT selection has beenmade on the display controller (Figure 212). The pilot isgiven the choice of displaying the waypoints with or without ID1abels. The flight plan WPTS are displayed in white with whiteconnecting path line (DTRK). The TO WPT is always displayed inmagenta (consistentwith the FMS CDU) for a visual indicationof the active flight plan leg. A maximum of 16 WPTS can bedisplayed on the MAP at any time.

Navaids - The FMS outputs data for the 10 (max) closest navaidsto present position. These are displayed, when selected, inyellow using the standard vortac symbol (Figure 213). Navaidscan be displayed with or without IDs as selected through thedisplay controller.

Airports - The FMS outputs data for the nine (max) closestairports to present position. These are displayed, whenselected, in cyan using a circular symbol (Figure 213).Airports can be displayed with or without IDs as selectedthrough the display controller.

Waypoint Annunciations - Track changes, both lateral andvertical, are given at the upper right-hand side of the MAP arc(Figure 212). The annunciations are displayed in white andlabeled as follows:

VERTALRT - Vertical Track ChangeTRK CHG - Lateral Track Change

These annunciations appear flashing for 5 seconds uponactivation and remain steady for the duration of the mode.When coming out of the alert mode, the annunciation is boxedfor 5 seconds.

The OFFSET annunciation (Figure 212) is displayed below the TRKCHG annunciation whenever the lateral track offset mode hasbeen selected on the FMS CDU.

Joystick - The aircraft installed joystick is used to move acursor around on the MAP format. The cursor is a green diamondshape attached to a green dashed line which originates at theaircraft symbol. When the cursor is at the desired location,the enter button is depressed. This transmits the positiondata (LAT/LON) into the CDU scratchpad of the displayed FMS.This information can then be line selected into the active FMSflight plan. Selection of the clear pushbutton returns thejoystick to the origin (i.e., hidden under aircraft symbol).

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2. D. (2) (p) Vertical Profile - The vertical profile mode allows the crew toview a portion of the flight plan in the vertical dimension.This mode is selected through the display controller (Figure213). The previous discussion on waypoints apply. Navaids andairports are not displayed in the vertical profile area.

The vertical profile display uses the symbols previouslydiscussed in the lateral map section of this document. Thisdisplay is limited to five symbols (maximum) to make the bestuse of the limited display area. These symbols includewaypoints, top of climb (TOC), top of descent (TOD), andtransition points. The display vertical dimension is altitudein feet. The vertical dimension is scaled to be t20,000 feetabout the aircraft altitude. The lateral dimension is distancein miles and is scaled to be equivalent to the lateral MAPrange.

The altitude is identifiedfollows:

*

FL300

FL300

FL300

FL300

The waypoint identifier is

further by type of

Tv~e of Constraint

At or Below

At or Above

At.

Predicted

normally displayed

constraint as

in white tomaintain consistency with the lateral flight plan. The TOwaypoint is the exception in that it will-be displayed inmagenta for consistency with lateral map and CDU. The waypointaltitude is displayed in cyan and is located below theappropriate waypoint.

~ TOC/TOD Waypoints - The TOC and TOD (top of climb/descent)waypoints are represented on the vertical profile displayusing a diamond symbol (Figure 214).

~ Transition Point - A transition point is defined as a pointthat has no ident or altitude labels but is on the verticalpath (Figure 214).

~ Altitude Preselector - The altitude preselector is shown onthe map as a horizontal green dashed line (Figure 214).This value corresponds to the value displayed above thealtitude tape on the PFD.

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2. D. (2) (p) ~ Holding Pattern - When a flight plan holding pattern isactive, a hold annunciation will be displayed at the holdingpoint on the profile. In this situation, the vertical mapwill be held static in the distance axis and updated in thealtitude axis only. The racetrack pattern appears on thelateral map only.

(3) MAP Caution and Warning Displays

(a) IRS Failures and F1ags - Failure of the displayed heading fromthe IRS is shown by removing the flight plan, bearing pointers,track line, and wind vectors from the display. The digitalheading readout is replaced by amber dashes. Additionally, ared HDG FAIL label is displayed below the lubber line as shownin Figure 215.

IRS test mode is shown with the label HDG TEST in the samelocation as the fail annunciation discussed above (Figure 216).

(b) Headinu Select Failure - Failure of the heading select signals. .will c~use the display to be replaced by amber-dashes with theheading bug removed from the display. This indication willalso be given in the event of an invalid heading display(Figure 215).

/

ALTITUDE PRESELECTOR

TOC TILLSWx --–------mvT- +

——-. ————- —. SATVAR L 10 - 56GCR TAS

234GSPD

345)I

TRANSITION POINT

Vertical Profile SymbolsFigure 214

22-14-00Page 243


w——— FMS 1

HDG---

Wx

50

\

+50

1 0 ADF1\

O VOR2

SAT---

TAS---

GSPD---

Map Caution/Warning DisplaysFigure 215

22-14-00Page-244


Honeywell:;W::ANCE

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Ft4S1 “as345 v NAV 1‘

23. 4DrI~

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IRS Test Mode DisplayFigure 216

22-14-00Page 245


2. D. (3) (c)

(d)

(e)

(f)

DADC Failures - DADC failures are indicated by replacing theSAT and TAS numerical values by amber dashes (Figure 215).

FMS Failure - A failure of the FMS removes the active flightplan waypoints, navaids, and airports from the display. Thisindication will also be given in the event of an invalidheading display (Figure 215). The digital GSPD window isreplaced by amber dashes.

Weather Radar Failure - A failure is shown by displaying anamber WX in the weather radar mode annunciation box (Figure215).

EFIS Self-Test - This test is initiated through the on-sidedisplay controller TEST mode when airspeed is valid and lessthan 60 knots and the WOW switch is active. Selecting the EFISself-test shows the invalid flags for the following MAP display

-parameters:

● IRS● DADC. VERT TRK, TRK CHG Annunciations

(4) Compass (COMP) Mode

The COMP mode is comprised of the following functions (Figure 217):

● 360-Degree Compass Rose● Digital Heading Readout● Magnetic/True Annunciation. Heading Bug with Digital Readout. Wind Display in Vector or X-Y Formats● Course Pointer/CDI with Digital Readout● Navigation Source Digital Distance Readout● Navigation Source Annunciation● To/From Indicator● Bearing Pointers Circle/Diamondc Bearing Pointer Source Annunciation● Vertical NAV Scale (GS/VNAV)● NAV Preview Mode

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2. D. (4) (a) Aircraft Symbol - The white aircraftcue to aircraft position relative toand selected course.

symbol provides a visualheading, selected heading,

(b) Heading Display - The displayed compass rose rotates through360 degrees of aircraft heading, is marked in 5-degreeincrements, and is labeled every 30 degrees. The compass roserotates about the stationary aircraft symbol to provide actualheading information. The compass indices and labels aredisplayed in white. The display is augmented with a whitedigital readout of actual heading above the compass lubber-line(Figure 217).

(c) Heading MAG/TRU Annunciations - TheMAG/TRU annunciationoperation and location is identical to the MAP mode (paragraph2.D.(2)(d)) as previously discussed.

(d) Heading Bug with Readout - The heading bug operation andidentification is as previously discussed in the MAP mode(paragraph 2.D.(2)(e)).

(e) Wind Display - The wind format is displayed in white in eitheran X-Y (Figure 217) or vector (Figure 218) format as selectedthrough the display controller. The wind display abeams theaircraft symbol to the left, above the heading select digitalreadout. This data is displayed from the selected ND FMSsource.

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Honeywell !!!$C.CE

BEAR INGPOINTERS

TO/FROM DISTANCE NAVSOURCE

J >

CRS000 H23.%;;

DME- HOLD

COURSEDEVIATION‘INDICATOR

VERTICAL~ NAVSOURCE

VERTICAL% NAVSCALE

\BEARINGSOURCES

Compass Mode Display Format(COMP)

Figure 217 22-14-00Page 248


2. D. (4) (f) Course Deviation, Course Pointer with Readout - The coursedeviation bar represents the centerline of the selectednavigation or localizer course. The aircraft symbolpictorially shows aircraft position in relation to thedisplayed deviation.

The course pointer is positioned around the rotating headingdial in one of two manners:

● Course select knob located on the flight guidance controllerfor short range NAV (SRN) sources (i.e., NAV, ILS, MLS).

● Long range NAV (LRN) sources automatically select thedesired track data for the active flight plan (i.e., FMS).

A digital readout corresponding to the pointer position isprovided and located to the upper left of the compass rose.The data is labeled CRS for SRNS and DTRK for FMS. Thepointer, deviation indicator, and digital readouts are greenfor SRN sources and cyan for FMS. The center of the courseselect knob contains a button labeled SYNC. Selection of theSYNC button causes the course pointer to align with the bearingof the selected NAV source causing the deviation bar to center.The SYNC mode is only operational when the displayed navigationsource is VOR.

(9)

(h)

Distance Display - The distance display indicates the nauticalmiles to the selected DME station or FMS waypoint. Thedistance display is located to the upper right of the compassarc. The data is labeled DME (Figure 217) for SRN sources andNM (Figure 218) for FMS. The digital display color isconsistent with the course select pointer (i.e., green for SRNSand cyan for FMS). DME hold is indicated with an amber Hdisplayed adjacent to the distance readout.

NAV Source Annunciations - Annunciation of the navigationsource is displayed in the upper right-hand corner of thecompass rose directly above the distance readout. The crew hasthe ability to select all navigation sources available in anygiven installationthrough the display controller as shown inFigure 207 (note the PFD displayed NAV source is also the COMPmode displayed NAV source). The labels are normally displayedin white. When the pilot and copilot NAV sources are the same,they are annunciated on both displays in amber.

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2. D. (4) (i) TO/FROM Indicator - A magenta arrow head appears in the centerof the HSI, originating at the fixed aircraft symbol nose ortail as appropriate. The arrow indicates whether the selectedcourse will take the aircraft TO or FROM the station orwaypoint. The TO/FROM annunciator is not in view duringlocalizer operation.

(j)

(k)

Bearing Pointers with Annunciation - Each pilot can display upto two bearing pointers on the compass arc simultaneously. Thebearing pointers are distinguishable by shape (circle ordiamond) and color (white or yellow). Each bearing pointersource is labeled and displayed in the lower right-hand cornerof the HSI. The circle pointer is always displayedpositionally above the diamond pointer. The sources availableon each bearing pointer are selected through the displaycontroller (Figure 217).

Vertical Navigational Scale - The vertical navigation scaleappears to the right of the compass rose (Figure 217). Thepointer source is selected through the display controller. Thevertical navigation source displayed is associated with theactive lateral navigation source selected for display (i.e.,glideslope with ILS, VNAV with FMS, etc.). The display formatand operation is identical to the description already providedin paragraph 2.C.(1)(0).

(1) Waypoint Annunciations - Track changes, both lateral andvertical, are given at the upper right-hand side of the compassrose (Figure 218). Directly below these annunciations is thetrack offset annunciation. These annunciations are displayedin white and labeled as follows:

VERT ALRT - Vertical Track ChangeTRK CHG - Lateral Track ChangeOFFSET - Track Offset Mode

These annunciations are active as previously discussed inparagraph 2.D.(2)(n).

22-14-00Page-250


\

/WIND

DTRK000

FMS f23.4 NM

HDG000

1!:R --.-—— --. —-— --r-

GCR I ‘Lcw ii----i CRS NAV2;[ 330 23. 4DMEI

L ----- ----- —--- A\

TAS234

GSPD345

Navigation Preview Mode(NAV)

Figure 218

FMS ALERT-MESSAGES

PREVIEWED/ NAVSOURCE

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2. D. (4) (m) Navigation Preview Mode - The crew has the capability topreview other navigation sources simultaneously with the activeNAV source. This mode is selected through the displaycontroller as shown in Figure 218. When active, a ghosted (dimwhite-dashed lines) course/CDI is presented on the compassrose. Additionally, a preview mode annunciation box is broughtup on the display centered between the weather mode box and theSAT/TAS and GSPD box. This box contains the previewed NAVsource label, course select readout, and distance display. Allprevious discussion of these modes for the compass format applywith the exception of color. All preview mode associateddisplays follow the convention of using dim-white, dashed lineswith dim-white digital readouts.

(5) COMP Caution and Warning Displays

(a) IRS Failures and Flags - Failure of the displayed heading fromthe IRS is shown by removing the bearing pointers, coursepointer/deviation bar, wind display, and TO/FROM indicator.The digital heading readout is replaced by amber dashes.Additionally, a red HDG FAIL label is displayed below thelubber line as shown in Figure 219.

lRS test mode is shown with the label HDG TEST in the samelocation as the fail annunciation discussed above (Figure 216).

(b) Heading Select Failure - Failure of the heading select signalswill cause the digital display to be replaced by amber dasheswith the heading bug removed from the display. This indicationwill also be gi~en in the event of an invalid(Figure 219).

(c) Course Select Fai1ure - Fai1ure of the coursewill cause the digital display to be replacedwith the course pointer removed from the disp’indication will also be given in the event ofheading display or FMS source (Figure 219).

heading display

select signalsbv ambera~. Thian inval

(d) Course Deviation Failure - A failure of the course dev.indicator is shown by removing the deviation bar and d“a red “X” through the deviation dots (Figure 220).

dashes

d

ationsplaying

22-14-00Page 252


Honeywell #&$~.cE

CRS———

I-IDG---

WxVARGCR

NAV 1 ‘23. 4DME

s+ ,Q’\\\\\

I 0

0

0

o

0

() ADF 1@/0R2

SAT- 56

TAS234

GSPD—--

Compass Caution/Warning Displays (IRS)Figure 219

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M#JJJHJANCE

Honeywell ...~.~.

CRS000

HDG000Wx

IL!

[\,\\’\s 33

\\

P1

NAV 1----DME

SAT——-TAS--—

GSPD345

Compass Caution/Warning Displays (Mist)Figure 220

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2. D. (5) (e) Distance Display Failures - Failure of either the DME or FMSdistance signals is indicated by replacing the digital distancevalue with amber dashes (Figure 220).

(f) Vertical Navigation Scale - A failure of the source driving thevertical navigation scale is shown by removing the deviationpointer and displaying a red “X” through the deviation dots asshown in Figure 220.

(g) DADC Failures, Weather Radar Failures - The same indicationsare given as previously described in the MAP mode (paragraph2.D.(3)(c)).

(h) Previewed NAV Source Failure - A failure of the previewed NAVsource is shown in the same manner as the active NAV sourcefailure previously discussed. A red, dashed “X” is displayedthrough the deviation scale to preserve the ghosted image ofthe preview symbology.

(i) EFIS Self-Test - This test is initiated through the on-sidedisplay controller TEST mode when airspeed is valid and lessthan 60 knots and the WOW switch is active. Selecting the EFISself-test shows the invalid flags for the following compassdisplay parameters:

● IRS● DADCc VERT TRK, TRK CHG Annunciations. Distance Display. Vertical Deviation Display

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Honeywell !!!#&#r.cE2. D. (6) Plan (PLAN) Mode

The PLAN mode is comprised of the following functions:

. True North-Up Map Presentation● Heading Source Annunciation. Wind Display in Head/Tail and Left/Right Components● FMS 1/2 Source Annunciation. Flight Plan Range Ring. FMS Waypoint Annunciations● FMS Waypoint, Airport, and Navaid Display Provisions● FMS Desired Track Lines. Flight Plan Scroll

(a) Aircraft Symbol - The green aircraft symbol moves about thedisplay area as a function of present position. The aircraftsymbol provides a visual cue as to the actual aircraft positionin relation to true north and the active flight plan.

(b) Heading Displaypresentation isnorth-up arrow ~adjacent to theheading indicat

- In this display format, the headingfixed at true north-up as shown by the whiten the upper right corner of the displayFMS source (Figure 221). This is the onlyon given in the plan mode.

(c) Wind Display - The wind format is displayed in white using heador tail wind and left or right cross-wind labels with theassociated magnitude (Figure 221). The left/right componentsare displayed positionally above the head/tail components atall times. The symbols used are defined as follows:

LABEL DEFINITION

H-10 Head wind of 10 knotsT-10 Tail wind of 10 knotsR-10 Right cross-wind of 10 knotsL-10 Left cross-wind of 10 knots

(d) NAV Source Annunciations - NAV Source annunciation functionsare identical to the MAP mode source annunciations previouslydiscussed in paragraph 2.D.(2)(f).

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rml

WIND‘COMPONENTS

NORTH-UP~INDICATOR

FtlSsol

L-4T-5

-wXVARGCR

KBNLo

+MILEN

“wy

25

VERT ALRTTRK CHG

OFFSET

\25 0

KNBC

/‘2

0KLHU

- 56TAS234

GSPD345

Use or disclosure of reformation

!CE

‘MS ALERTMESSAGES

PLANMODE

POSITION

Plan Mode Display(PLAN)

Figure 221

Format

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2. D. (6) (e) Range Ring - A range circle is displayed to aid in determiningthe position of the active flight plan parameters. The circlecorresponds to half the selected range on the weather radarcontroller. The circle is labeled on both sides with the half-range data.

(f) Active F1ight Plan Waypoints (WAYPT) - The active flight plancan be displayed in the plan mode whenever the WAYPT selectionhas been made on the display controller (Figure 221). Thepilot is given the choice of displaying the waypoints with orwithout ID labels. The flight plan waypoints are displayed inwhite with a white connecting path line (DTRK). The TOwaypoint is always displayed in magenta (consistentwith theFMS CDU) for a visual indication of the active flight plan leg.A maximum of 16 waypoints can be displayed on the MAP at anytime.

(g) Navaids - The FMS outputs data for the 10 (max) closest navaidsto present position. These are displayed, when selected, inyellow using the standard vortac symbol (Figure 221). Navaidscan be displayed with or without IDs as selected through thedisplay controller.

(h) Airports - The FMS outputs data for the nine (max) closestairports to present position. These are displayed, whenselected, in cyan using a circular symbol (Figure 221).Airports can be displayed with or without IDs as selectedthrough the display controller.

(i) Waypoint Annunciations - Track changes, both lateral andvertical, are given at the upper right-hand side of the planrange ring (Figure 221). Directly below these annunciations isthe track offset annunciation. The annunciations are displayedin white and labeled as follows:

VERT ALRT - Vertical Track ChangeTRK CHG - Lateral Track ChangeOFFSET - Track Offset Mode

These annunciations are active as previously discussed inparagraph 2.D.(2)(n).

(j) Joystick - Joystick operation is as previously discussed inparagraph 2.D.(2)(0). The one difference is that the joystickcursor remains fixed in the center of the display and theactive flight plan is moved about this point with the joystick.

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2. D. (6) (k) F1ight Plan Scroll Mode - Once a flight plan has been selectedfor display, the crew can then scroll through the entire flightplan using the FPLN SCROLL buttons on the display controller(Figure 221). Selection of the scroll FORE allows the pilot tomove the waypoints of the active flight plan one at a timethrough the center of the plan range ring in the forward(origin to destination) direction. Selection of the scrollBACK allows the pilot to move the waypoints of the activeflight plan one at a time through the center of the plan rangering in the backward (destinationto origin) direction.

The maximum number of waypoints that can be displayed at onetime is limited to 16. The 16-waypoint window for scrollingpurposes is made up of the past three FROM waypoints and the TOthrough the TO+12 waypoints. The scroll window and selectedrange will determine the number of waypoints that will bedisplayed while scrolling the flight plan.

(7) PLAN Caution and Warning Displays

(a) IRS Failures and Flags - Failure of the displayed heading isshown by removing the active flight plan and wind display anddisplaying, in red, the label HDG FAIL at the top center of theplan range ring (Figure 222).

IRS test mode is shown with the label HDG TEST in the samelocation as the fail annunciation discussed above (seeFigure 216).

(b) FMS Failure - A failure of the FMS removes the active flightplan, navaids, and airports from the display. This indicationwill also be given in the event of an invalid heading display(Figure 222). The digital GPSD window is replaced by amberdashes.

(c) DADC Failures, Weather Radar Failures - The same indicationsare given as previously described in the MAP mode (paragraph2.D.(3)(c)).

(d) EFIS Self-Test - This test is initiated through the on-sidedisplay controller TEST mode when airspeed is valid and lessthan 60 knots and the WOW switch is active. Selecting the EFISself test shows the invalid flags for the following plandisplay parameters:

. IRS● DADC. VERT TRK, TRK CHG Annunciations

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4N

FMS1I

HDGFAIL

25 25

rSAT---

TASI ---I GSPD

I---

Plan Mode Caution/Warning DisplaysFigure 222

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2. E. Engine (ENG) Display

The ENG display is comprised of the following functions (Figure 223):

●

●

●

●

●

●

●

●

s

●

●

●

●

●

●

●

●

●

●

●

●

●

●

●

●

●

●

●

●

●

●

●

●

●

Analog EPR Scales with Digital ReadoutFMS (PZ) Computed EPR Limit/Target BugsFMS (PZ) Target Bug Digital ReadoutFMS Source/EPR Mode AnnunciationsManual EPR Limit Set Capability

Analog TGT Scales with Digital ReadoutAbnormal TGT Range Indication

Analog LP Scales with Digital ReadoutLP SYNC AnnunciationLP Anti-Ice AnnunciationsAbnormal LP Range Indication

Analog HP Scales with Digital ReadoutHP SYNC AnnunciationAnnunciation of Start Valve Open [SVO)Annunciations of Ignition (IGN)Abnormal HP Range Indication

Analog FF Scales with Digital ReadoutFuel Flow Valve Status Indications

Digital Readouts of Oil PressureOil Pressure Warning IndicationsDigital Readouts of Oil TemperatureOil Temperature Warning Indications

Digital Readouts of Engine Vibration Monitor (EVM) LPDigital Readouts of Engine Vibration Monitor (EVM) HP

Digital Readout of Combined Hydraulic PressureDigital Readout of Flight Hydraulic PressureCOMB/FLT Hydraulic Pressure Valve Status IndicationsDigital Readout of Auxiliary Hydraulic PressureDigital Readout of Utility Hydraulic Pressure

Digital Readout of Engine TemperatureEngine Fuel Temperature Warning IndicationsDigital Readouts of Fuel Quantity, Left, Right, and TotalDigital Readout of Combined Hydraulic PressureFuel Quantity Low Indication

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2. E. (1) Engine Instruments - General

The configuration of the GIV system includes two DA-884 DataAcquisition Units (DAU) which receive signals from the enginesensors for display on the engine and CAS (paragraph 2.F.) displays.DAU 1 is tied to the left engine sensors and DAU 2 is tied to theright engine sensors. Additionally, each DAU has two channels (Aand B) that are completely independent from one another. Eachchannel within a particular DAU is performing the same computationson the same engine parameters. The pilot has the ability to displaydata from one channel at a time from each DAU. The selectionbetween the two channels is accomplished through the displaycontroller (Figure 223). This is the crew’s means of reversion inthe event that the displayed DAU channel fails.

The primary engine instruments reside on the left two thirds of theengine display format. This area is divided into five rows ofengine gauges as follows:

. Engine Pressure Ratio (EPR)● Turbine Gas Temperature (TGT). Low Pressure Tach (LP). High Pressure Tach (HP)● Fuel Flow (FF)

Each engine gauge is a 225-degree arc with moving pointer and fixedcurrent value window located above the pointer rotation point. Eachengine gauge type is labeled in white between the left and rightgauges as shown in Figure 223.

The secondary engine parameters are displayed digitally in theremaining one-third of the display format. The parameters whichcomprise the secondary engine parameter display are:

●

●

●

●

●

●

●

●

●

●

Oil PressureOil TemperatureEngine Vibration Monitor (EVM) - LPEngine Vibration Monitor (EVM) - HPCombined Hydraulic PressureFlight Hydraulic PressureUtility Hydraulic PressureAuxiliary Hydraulic PressureEngine Fuel TemperatureFuel Quantity

22-14-00Page 262


2. E. (2) Primary Engine Instruments

(a)

(b)

Engine Pressure Ratio (EPR) Display - The EPR instruments(left/right) are the top row of scales on the primary enginedisplay format (Figure 223). The EPR arc boundaries range from

0.85 EPRto 2.00 EPR. The scale is marked every 0.1 EPRstarting from 0.9 and going to 2.00. The 0.85, 1.0, 1.5, and2.0 tick marks are highlighted with a longer tick. The whiteEPR current value corresponds to the pointer position on thescale and is readable to a 0.01 EPR resolution. If the actualdata goes below 0.85, the pointer remains stationary at thelower limit and the digits show actual EPR value. If theactual data goes above 2.00, the pointer remains stationary atthe upper limit and the digits show the actual EPR value. TheEPR arc, scale markings, current value window, and pointer arewhite.

Engine Pressure Ratio Transmitter (EPRT) Interface - Enginepressure ratio (EPR) is displayed on the engine display(DU No. 3) and on the standby engine instrument panel(Figure 224). The EPRT provides an indication of engine powerin the form of the ratio of engine fan duct total pressure tointake total pressure. Engine fan duct pressure is receivedpneumatically from the on-side engine. Total pressure is

received from the on-side AZ-81O Digital Air Data Computer(DADC) via ARINC 429. Cross-side DADC is used if the on-sideis invalid. To compensate for variations in the actual thrustversus EPR of different Rolls Royce Tay engines, a seven-wiretrim plug discrete input is provided. This allows the enginetechnician to select a trim appropriate for each engine. Thetrim plug input is read by the EPRT during power-upinitialization and is used by its software to trim the computedEPR. The computed EPR is transmitted out, over two redundantlow-speed ARINC 429 buses, to the DA-884 Data Acquisition Unit(DAU) and to the standby engine instrument signal conditioner.The DAU reads the information off of ARINC 429 and retransmitsit on ASCB for the display system to read.

(c) EPR Limit/Target Source Annunciation - The FMS source for thelimit (except manual limit) and target is annunciated above theEPR label on the engine display format (Figure 223). Thesource is always the PFD commanded (priority) side FMS. Thisinformation is also displayed on the TRS menu of the displaycontroller.

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2. E. (2) (d) FMSareThe

(PZ) EPR Limit Annunciation and Tickmark - EPR 1imit modesselected on the main TRS menu of the display controller.annunciation is displayed in white directly below the EPR

scale label. A white, enh~nced, tickmark is d~splayed alongthe EPR scale in correspondence with the EPR limit modeselected (Figure 225). The available FMS limit modes are asfollows:

Limit Mode Annunciation

Takeoff TOGo-Around GAClimb CLBCruise CRZMax Continuous MCT

(e) Manual EPR Limit Annunciation - The FMS calculated limits maybe overridden by selecting the second page of the TRS menu andmanually dialing an EPR limit value. The symbology is asdiscussed above for the limit marking on the scale. A MANannunciation is displayed in white in the same location as theFMS limit annunciation (Figure 226).

(f) EPR Target Bug and Digital Readout - The FMS-PZ calCU1ates atarget bug based on the selected EPR limit mode. This bug isdisplayed on the EPR scale in green (Figure 225) correspondingpositionally to the green digital readout above each EPRcurrent value window. Generally, there is not a modeannunciation given for the target bug since the targetcorresponds to the selected limit mode. FLEX mode is the oneexception. When FLEX mode is selected, takeoff (TO) is thedisplayed limit mode, and a green FLEX mode annunciation isdisplayed below the TO annunciation. The green bug ispositioned at the FLEX mode value.

(g) Turbine Gas Temperature (TGT) Display - The TGT instrumentsform the second row of scales on the primary engine instrumentdisplay (Figure 223). The TGT arc is banded with white, amber,and red colors to indicate normal, caution, and warning rangesrespectively on the arc. The arc boundaries range nonlinearlyfromO “C to 1000 “C. A set point tick mark is displayed atthe 785 “C point on the arc. The color band definition for TGTis as follows:

WHITE O “C< TGT < 715 ‘CAMBER 715 OCS TGT < 800 “CRED 800 “C< TGT < 1000 ‘C

22-14-00Page 264


The TGT current value corresponds to the pointer position onthe scale and is readable to a 1.0 “C resolution. If theactual data goes above 1000 “C, the pointer remains stationaryat the upper arc limit and the digits show the actual TGTvalue.

TGT exceedance (amber/red regions) is shown by growing thecolored band to twice the normal width and changing the pointerand current TGT digital value (normally white) color to matchthe arc region color.

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REDBAND%

AIM&;R_

wtsr; .

A:::~

LP *SYNCHED

START/VALVE

/IGNIT ON

EPR

TGT

LPSYNC

/’

HP

FF pj mo,

Engine instrument Display Format(SENSOR)

Figure 223 22-14-00Page 266


.—— — —.A2-S1ODADC NO. 1 9J1B

%

3031

3233

ARINC429[

HOUTPUTNO 3 L

1———————

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——— —rDA41B’IDATA

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L ——— —-

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37

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——— — J

STANDBYENGINEINSTRUMENTS

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——. —— ——SG-SS4SYMBOL GENERATORNO. 2

r 1——— —

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I ( hOP-SIDEDADC H =

IARINC 429 L 49

I

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I EPR OUTPUT

(

H 31

I

NO 1 ARINC 429 L 39

I

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(

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NO 2 ARINC 429 L 52

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——— —XbSSODISPLAYUNIT NO. 4

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)

PRIMARY SG/DU

L BUS II

H

)

ALTERNATE NO i

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)

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——— —— J

——— —hA4B4 DATA

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EnciinePressure RatioTransmitter Interface Diagram

Figure 224 22-14-00Page 267/268

Aug 15/91Use or disclosure of Informallon on Ihm page is subject 10 the restrictions on the Mle page of this document.

EPR EPRTARGET EPR LIMITr BUG > TARGET ~BUG

OIL PRESS

p4 m{OIL TEMP

mmEVM

~] LP~l

-I HPl~]

HYD PRESSCOMB FLT

ml plmq

ENG FUEL TEMP

mmFUEL QTY

& 50:

Engine instrument Display Format(TRS)

Figure 22522-14-00

Page 269Aug-15/91


2. E. (2) (h)

(i)

(j)

(k)

Low Pressure (LP) Tach Display - The LP instruments comprisethe center row of gauges on the primary engine display (Figure223). The LP arc is banded with white and red colors toindicate the normal and warning ranges respectively on the arc.The arc boundaries range nonlinearly from O%RPM to 110% RPM.The color range definition for LP is as follows:

WHITE O%RPM s LP < 95.5% RPMRED 95.5% RPM ~ LP < 110% RPM

The LP actual value corresponds to the pointer position on thescale and is readable to a 0.1% RPM resolution. If the actualdata goes above 110% RPM, the pointer remains stationary at theupper arc limit and the digits shown the actual LP value.

LP exceedance (red region) is shown by growing the red band totwice the normal width and changing the pointer and digital LPvalue color (normally white) to match the arc region color.

Anti-Ice (A/1) Annunciator - This annunciation is displayed ingreen above the LP digital readout windows. The annunciationlabel isA/I (Figure 223).

LP Synchronization (SYNC) Annunciator - If the crew chooses tosync the engines using LP, a SYNC annunciation is displayed ingreen directly below the LP scale label (Figure 223).

High Pressure (HP) Tach Display - The HP instruments form thefourth row of gauges on the primary engine display (Figure223). The HP arc is banded with red, white, amber, and redcolors to show the low end warning, normal, high end caution,and high end warning ranges on the arc. The arc boundariesrange nonlinearly between O% RPM to 110% RPM. The color rangedefinition for HP is as follows:

RED O%RPM s HP < 46.7% RPMWHITE 46.7% RPM ~ HP < 97.5% RPMAMBER 97.5% RPM s HP < 99.7% RPMRED 99.7% RPM g HP < 110% RPM

The HP current value corresponds to the pointer position on thescale and is readable to a 0.1% RPM resolution. If the actualdata goes above 110% RPM, the pointer remains stationary at theupper arc limit and the digits shown the actual HP value.

HP exceedance (amber and red region) is shown by growing thecolored band to twice the normal width and changing the pointerand digital HP value (normally white) color to match the arcregion color.

22-14-00Page 270


2. E. (2) (1) Start Valve Open (SVO) Annunciation - The SVO annunciation islocated adjacent (outboard side) to the HP scales as shown inFigure 223. When the start valve is open, the SVO label coloris determined by the following logic:

Condition Color

HP < 42% 61ueHP ~ 42% Flashing Amber

(m) Ignition (IGN) Annunciation - The green IGN annunciation islocated positionally below the SVO annunciation (Figure 223).This label is displayed whenever the igniters are active.

(n) HP Synchronization (SYNC) Annunciator - If the crew chooses tosync the engines using HP, a SYNC annunciation is displayed ingreen directly below the HP scale label (Figure 225).

(o) Fuel Flow (FF) Display - The FF instruments comprise the bottomrow of gauges on the primary engine display (Figure 223). TheFF arc boundaries range between 250 PPH (pounds per hour) to8000 PPH. The digital readout window is located directly abovethe pointer rotation point. The white FF actual valuecorresponds to the pointer position on the scale and isreadable to a 1O-PPH resolution. If the actual data goes below250 PPH, the pointer remains stationary at the lower limit andthe digits show actual FF value. If the actual data goes above8000 PPH, the pointer remains stationary at the upper limit andthe digits shown the actual FF value. The FF arc, scalemarkings, current value window, and pointer are white.

(p) Fuel Flow Valve Status Indications - When the fuel flow valvesare in the closed position, amber cross-hatching is displayedwithin the FF digital readout window (Figure 226). When thefuel flow valves are in transit, amber dashes are displayedwithin the FF digital readout window. The pointer remainsdisplayed to distinguish this condition from the invalid datacondition (paragraph 2.E. (4)).

22-14-00Page 271


MAINTENANCE

Honeywell %!!!~b.2. E. (3) Secondary Engine Instruments

(a) Engine Oil Pressure (OIL PRESS) Display - The engine oilpressure is displayed digitally in white in the first row ofthe secondary engine parameters (Figure 223). The resolutionof the readout is 1 PSI. The display digits change color as afunction of oil pressure as defined below:

RED O PSI < OIL PRESS < 16 PSIAMBER 16 PSI ~ OIL PRESS < 30 PSIWHITE 30 PSI ~ OIL PRESS

(b) Engine Oil Temperature (OIL TEMP) Display - The engine oiltemperature is displayed digitally in white in the second rowof the secondary engine parameter display (Figure 223). Theresolution of the digital readout is 1 “C. The display digitschange color as a function of oil temperature as defined below:

RED OILTEMP < -40 “CAMBER -40 “C < OIL TEMP < -30 “CWHITE -30 “C ~ OIL TEMP < 105 ‘CAMBER 105 “C < OIL TEMP < 120 ‘CRED 120 “C < OIL TEMP

(c) Engine Vibration Monitor (EVM) Display - The EVM displays(Figure 223) consists of two sets of windows (row 3 and 4)labeled LP and HP respectively. EVM is displayed digitally inwhite to a 0.01 IPS resolution.

(d) Combined Hydraulic Pressure (FLT HYD PRESS) - The COMB HYDPRESS digital readout is the left-most display in the fifth rowof the secondary engine parameter display (Figure 223). COMBHYD PRESS is displayed digitally in white to a 1OO-PSIresolution.

(e) F1ight Hydraulic Pressure (FLT HYD PRESS) - The FLT HYD PRESSdigital readout is located right of the COMB HYD PRESS displayin the fifth row of the secondary engine parameter display(Figure 223). FLT HYD PRESS is displayed digitally in white toa 1OO-PSI resolution.

22-14-00Page-272


VA~~EIN<

TRANSIT

FFVALVE “CLOSED

G ‘G

SvoIGN

OIL PRESS

D mlOIL TEMP

I EVM

W LT

AUXl-l -1

ENG FUEL TEMP

mmFUEL CITY

& 50:

COMB

FLT‘PRESSVALVECLOSED

Fuel Flow/Hydraulic Pressure Valve Symbology(TRS-MAN)Figure 226

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2. E. (3) (f) COMB/FLT Hydraulic Pressure Valve Status Indications - When theCOMB/FLT HYD PRESS valves are in the closed position, ambercross-hatching is displayed within the appropriate hydraulicpressure digital readout window (Figure 226). When theCOMB/FLT HYD PRESS valves are in transit, amber dashes aredisplayed within the appropriate hydraulic pressure digitalreadout window.

(9) Utility Hydraulic pressure (UTIL HYD pRESS) - The UTIL HYDPRESS digital readout is the left-most display in the sixth rowof the secondary engine parameter display (Figure 223). UTILHYD PRESS is displayed digitally in white to a 100 PSIresolution.

(h) Auxiliary Hydraulic Pressure (AUX HYD PRESS) - The AUX HYDPRESS digital readout is located right of the UTIL HYD PRESSdisplay in the fifth row of the secondary engine parameterdisplay (Figure 223). AUX HYD PRESS is displayed digitally inwhite to a 1OO-PSI resolution.

(i) Engine Fuel Temperature (ENG FUEL TEMP) Display - The ENG FUELTEMP is displayed digitally in white in the seventh row of thesecondary engine parameter display (Figure 223). Theresolution of the digital readout is 1 ‘C. The display digitschange color as a function of oil temperature as defined below:

RED FUEL TEMP < -40 “CWHITE -40 “C s FUEL TEMP < 90 “cAMBER 90 ‘C s FUEL TEMP < 120 ‘CRED 120 ‘C s FUEL TEMP

(j) Fuel Quantity (FUEL QTY) - This FUEL QTY display (Figure 223)consists of three digital windows comprising rows 8 and 9. Twowi;ows (row 9) corresponds to the fuel quantity in each (L/R)

The third window (row 8) is a running total (T) of thefuel”quantity from both tanks. All three digital displays arewhite (normal) and readable to a 50-pound resolution. Thedisplay digits go to amber when a fuel level low indication isreceived. The total window turns amber only if both sidesindicate fuel level low.

22-14-00Page 274


Honeywell f~~~~vc’

2. E. (4) Engine Instruments Failure Indications

(a)

(b)

(c)

(d)

(e)

(f)

Data Acquisition Unit (DAU) Failure - A DAU failure isindicated on the primary engine display by removing the datapointers from all the arc displays. Additionally, the currentvalue digits are replaced with amber dashes. The failureindication on the secondary engine display consists ofreplacing the digital readout with amber dashes. Figure 227

shows the indication for a single DAU (No. 1) channel failure.

Engine Sensor/Interface Failure - The failure indication of aparticular sensor or interface is a subset of the DAU failurepreviously discussed. The difference is that the failureindication is shown only on the affected parameter asappropriate (Figure 227).

Display Controller Failure - In the event of a dual displaycontroller failure, EPR mode annunciations and command/limitbugs are removed from the EPR scale and the EPR target digitalreadout is replaced with amber dashes (Figure 227).

Bus Controller Failures - In the event the three buscontrollers fail the indication given would be that of a dualdisplay controller failure and a dual DAU failure as previouslydiscussed.

FMS Failure - Failure of the PFD commanded FMS-PZ causes thecommand bug to be removed from the display with thecorresponding digital readout replaced with amber dashes. FMScalculated limit tickmark and mode annunciations are alsoremoved from the display.

EICAS Self-Test - This test is initiated through either displaycontroller TEST mode when airspeed is valid and less than 60knots and the WOW switch is active. Selecting the EICAS self-test gives the following indications on the engine display:

● DAU lA-lB Failure

Indication. DAU 2A-2B Over-Range- TGT- LP- HP- Oil Temperature- Engine Fuel Temperature

. Fuel Quantity Low Indication

. A/I, IGN, and SVO Annunciations

The above indications are also given on the system pages (referto paragraph 2.G.) as appropriate.

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Honeywell !!!!!!5.”

-----

FF

G

OIL PRESS

OIL TEMP

emEVM

HYD PRESSCOMB FLT

UTIL AU)(

c1 l-mm-l

:NG FUEL TEMP

FUEL QTYT

Engine Instrument Display Failure IndicationsFigure 227

22-14-00Page 276


2. F. Crew Alerting System (CAS) Display

The crew alerting system (Figure 228) is comprised of the followingfunctions:

● 23 Message Lines Max. 18 Characters Per Line Max● 1 Message Status Line. CAS Display Declutter Via Scroll Buttons● Master Caution Button Acknowledgement of Messages● CAS Computed Messages

(1) CAS Display - General

The G-IV system is configured with two FC-880 Fault WarningComputers (FWC) which receive warning, caution, and advisory signalsfrom various subsystems in the aircraft. Each FWC is performing thesame computations on identical signal inputs. The selection betweenwhich FWC is displayed is accomplished through the DC-884 DisplayController (Figure 228). This is the crew’s means of reversion inthe event of a FWC failure.

The CAS display unit is located positionally under the enginedisplay. The CAS format comprises the left-hand one-third of theusable display space. The color convention and displayed priorityfor messages on the CAS display is defined as follows:

RED - Warnings (Top message Stack)AMBER - Cautions (Middle Message Stack)BLUE - Advisories (Bottom Message Stack)

. Master Warning/Caution Panel - The main crew interface with theCAS display is through the warning/caution panel mounted in frontof each pilot on the instrument panel. This panel is shown inFigure 229. Each annunciator shown is also a push to selectswitch. The upper left annunciator/switch is the redwarning/message acknowledgement pushbutton. Below this switch isthe amber warning/message acknowledgement pushbutton. The upperright button, labeled with amber UP, is the message scroll uppushbutton. The lower right button, labeled with an amber DN, isthe message scroll down pushbutton. Each annunciator andpushbutton function will be discussed in later paragraphs.

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Honeywell !!!!L!!!%H.CE

CABIN PRESSURE LOWBAGGAGE DOORL-R FUEL PRES LOWAOA HEAT 2 FAILL-R AIL HYD OFFL-R PITOT HT FAILBATT 1-2 CHGR FAILL-R AC POWER FAILICE DETECTEDTAT PROBE HT FAILTONE GEN FAILVOICE REC FAILFLIGHT REC FAILAC EXT POWERDC EXT POWERVHF COMM 1-2 FAILISOLATION VLV OPENSERVICE DOORSL OIL FILT BPASSR COWL A/IE BATT 1 FAILAPU ALT OFFCPL DATA INVALID

I’1’r J3$

LAMBER ‘ ~BLUE MESSAGE STATUS

MESSAGE STATUS

Crew Alerting System (CAS) Display FormatFigure 228

22-14-00Page 278


l—————— -————— —————— —-———— -

iMASTERII1II~

II

WARN

IIIIII

n]

cI1

WARNVHIBIT

CASSCROLL

ONJ

AP OFFI

❑UPIIIIII1III1

❑✽1I

DNIII[I

.—— — ———— ——-— ———— ———— ———— d

Master Warning/Caution PanelFigure 229

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Honeywell !#!#r.cE2. F. (2) CAS Messages and Controls

Tables 202, 203, and 204 show the Phase II messacielist. with thesource and” the type of inputexample of how to use Tables

Messaqe

L-R Fuel Filter DAU 1,2

*Message enable state/normal

to the source. The-following is an202, 203, or 204.

Source Tvs)e

JIB-24 WSP 15, BIT 15 *28/o

operation

For the CAS message L FUEL FILTER to be displayed, DAU 1 must have28 V (message enable state) on JIB-24 and then DAU 1 will set bit 15of WSP15 on ASCB. The FWC upon recognizing this bit set willtransmit the ASCII code for this message over ASCB to the SGdisplaying CAS on DU 3 or 4. This message is not included in theINHIBIT function.

(a) Warning (Red) Messages - The available red messages are listedin Table 202. When a red message becomes active the followingindications are given:

. red message displayed flashing on the CAS display

. red message aural alert is sounded

. red annunciator on the master warn panel is lit

When the red annunciator/pushbutton is selected on the masterwarning panel, the red annunciator is extinguished and the redmessage is displayed steady on the CAS display.

If there are red messages already on the CAS display when a newred message becomes active, the new message is put at the topof the red message stack. Acknowledged messages are removedfrom the CAS display when the condition causing the message isremedied.

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Message Source Type ~

ACFT CONFIGURATIONAFT EQUIP HOTAPU FIREBAGGAGE DOORCABIN PRESSURE LOWCABIN DFRN - 9.8CHECK PFDDAU 1-2 MISCMP-MSGENG FIRE LOOP ALRTFLAME DETECTGND SPOILERL-R ENGINE HOTL-R FUEL FILTERL-R FUEL PRESS LOWL-R PYLON HOTL-R REV UNLOCKL-ROIL PRES LOWMAIN DOORSMOKE DETECT

FWCDAU 1DAU 1DAU 2DAU 1DAU 1FWCFWCDAU 1DAU 2DAU 2DAU 1,2DAU 1,2DAU 1,2DAU 1,2DAU 1,2DAU 1,2DAU tDAU 1

J1B-10 28/0JIB-48 WSP 17, BIT7 28/0JIB-47 WSP 17, BIT 6 28/0 ~JIB-58 WSP 18, BIT 1 28/0JIB-20 WSP 15, BIT 11 2;;: ~JIB-70 WSP 18, BIT 12CALCULATEDCALCULATEDJIB-69 WSP 18, BIT 11 G;OJIB-31 WSP 16, BIT6 28/0JIB-55 WSP 17, BIT 14 28/0JIB-13 WSP 15, BIT4 28/0 ~JIB-24 WSP 15, BIT 15 28/0J1B-10 WSP 15, BIT 1 28/0JIB-12 WSP 15, BIT3 28/0JIB-9 WSP 15, BITO 28/0 ~JIB-14 WSP 15, BIT5 28/0JIB-20 WSP 15, BIT 11 28/0JIB-31 WSP 16, BIT6 28/0

A91 Messa e enable state/normal operation.

A2 Message enable activates checklist auto callup. Engine firediscrete (DAU 1,2 JIB-66, WSP 18, BIT 8) also activates.

CAS Red Warning MessagesTable 202

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2. F. (2) (b) Caution (Amber) Messages - The available amber messages arelisted in Table 203. When an amber message becomes active thefollowing indications are given:

. amber message displayed flashing on the CAS display

. amber message aural alert is soundedc amber annunciator on the master warn panel is lit

When the amber annunciator/pushbutton is selected on the masterwarning panel, the amber annunciator is extinguished and theamber message is displayed steady on the CAS display. Althoughthe red annunciator is not lit, selection of the red pushbuttonacknowledges amber messages also.

If there are amber messages already on the CAS display when anew amber message becomes active, the new message is put at thetop of the amber message stack. If the existing amber or bluemessages are scrolled off the display, the amber and bluemessage stacks are brought back on the CAS display with the newmessage displayed at the top of the amber stack. Acknowledgedmessages are removed from the CAS display when the conditioncausing the message is remedied.

(c) Advisory (Blue) Messages - The available blue messages arelisted in Table 204. When a blue message becomes active thefollowing indications are given:

. blue message displayed flashing on the CAS display

. blue message aural alert is sounded

The blue message appears flashing on the CAS display for5 seconds after which time the message is displayed steady.No pilot interaction is required to acknowledge this type ofmessage.

If there are blue messages already on the CAS display when anew blue message becomes active, the new message is displayedat the top of the blue message stack. If the existing amber orblue messages are scrolled off the display, the blue messagestacks are brought back on the CAS display with the new messagedisplayed at the top of the stack. Blue messages are removedfrom the CAS display when the condition causing the message isremedied.

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AHRS COOL FAIL FWC JIA-27 G/OALT MODE OFF FGC WSP 39, BIT 8ANTI-SKID FAIL DAU 2 JIB-74 WSP 19, BIT O O~GAOA HEAT 1-2 FAIL FWC JIA-73,74 O/GAP OFF FGC WSP 38, BIT 11AP TRIM FAIL FGC WSP38, BITOAPU ALT BRG FAIL DAU 2 JIB-75 WSP 19, BIT 1 G~OAPU ALT HOT DAU 1 JIB-34 WSP 16, BIT9 28/0APU MASTER WARN DAU 1 JIB-75 WSP 19, BIT 1 G/OAUX AC POWER FAIL DAU 1 JIB-46 WSP 17, BIT 5 28/0AUXILIARY HYD HOT DAU 1 JIB-50 WSP 17, BIT9 28/0BATT 1-2 CHGR FAIL DAU 1,2 JIB-80 WSP 19, BIT6 O/GBRAKE FAIL DAU 1 JIB-74 WSP 19, BIT O O/GBRAKE PEDAL DAU 2 JIB-73 WSP 18, BIT 15 O/GBRAKE OUHT FWC 1 JIA-84 G/O ~CABIN DFRN - 9.6 DAU 2 JIB-70 WSP 18, BIT 12 G/OCABIN OXYGEN ON DAU 1 JIB-45 WSP 17, BIT 4 28/0CABIN PRES MANUAL DAU 2 JIB-45 WSP 17, BIT 4 28/0CAT 2 INVALID FWC ASCB LOGIC IN FWCCHECK DU1-2-3-4-5-6 FWC CALCULATEDCHECK VSPEEDS PERF CALCULATEDCMB-FLT HYD FAIL DAU 1,2 JIB-23 WSP 15, BIT 14 28j0CPL DATA INVALID FGC ASCB + WOW LOGIC IN FWCDAU 1-2 MISCMP-ENG FWC CALCULATEDDAU 1-2 MISCMP-MSG FWC CALCULATEDDU FAN 1-2 FAIL DAU 1,2 JIB-78 WSP 19, BIT4 OjGEL CMB-FLT HYDOFF DAU 1,2 JIB-33 WSP 16, BIT8 28/0 ~EL MISTRIM NOSE DN FGC WSP 38, BIT 3EL MISTRIM NOSE UP FGC WSP38, BIT 4ENG FLT LOOP ALRT DAU 2 JIB-69 WSP 18, BIT 11 G;OEPMP BATT SW OFF DAU 1 JIB-56 WSP 17, BIT 15 28/0EPMP POWER FAIL DAU 1 JIB-59 WSP 18, BIT 2 28/0FGC 1-2 FAIL FGC ASCB LOGIC IN FWCFWD RADIO RACK HOT DAU 2 JIB-48 WSP 17, BIT 7 28/0ICE DETECTED DAU 2 JIB-52 WSP 17, BIT 11 28/0IRS 1-2 ALN FAULT IRS WSP33, BIT 10 OR 12 -IRS 1-2 COOL FAIL IRS WSP 34, BIT 4IRS 1-2 ON BATTERY IRS WSP33, BIT 6AND OL-R AC POWER FAIL DAU 1,2 JIB-36 WSP 16, BIT 11 28j0L-RAIL HYD OFF DAU 1,2 JIB-25 WSP 16, BITO 28/0 ~L-RALT BRG FAIL DAU 1,2 JIB-72 WSP 18, BIT 14 G/OL-R ALT FUEL FAIL DAU 1,2 JIB-42 WSP 17, BIT 1 28/0L-RALT HOT DAU 1,2 JIB-29 WSP 16, BIT 4 28/0L-R BLEED AIR HOT DAU 1,2 JIB-26 WSP 16, BIT 1 28/0L-R BLEED PRES HI DAU 1,2 JIB-27 WSP 16, BIT 2 28/0L-R CONV FAN FAIL DAU 1,2 JIB-32 WSP 16, BIT 7 28/0

CAS Amber Caution MessaaesTable 203 “ 22-14-00

Page 283Apr 15/93



L-R CONV HOT DAU 1,2 JIB-30 WSP 16, BIT 5 28/0L-R COOL TURB HOT DAU 1,2 JIB-35 WSP 16, BIT 10 28/0L-R COWL A/I OVHT DAU 1,2 JIB-64 WSP 18, BIT 7 28/0L-R COWL PRESS LOW FWC JIB-17,18 28/0 ~L-R DC POWER FAIL DAU 1,2 JIB-37 WSP 16, BIT 12 28/0L-R FUEL LEVEL LOW DAU 1,2 JIB-11 WSP 15, BIT 2 28/0L-RMAIN FUEL FAIL DAU 1,2 JIB-41 WSP 17, BIT O 28/0L-R PITOT HT FAIL DAU 1,2 JIB-71 WSP 18, BIT 13 G/OL-R WING HOT DAU 1,2 JIB-40 WSP 16, BIT 15 28/0L-R WING TEMP LOW FWC JIB-22,23 0/28 ~MACH TRIM LIMIT FGC WSP38, BIT 13MACH TRIM OFF FGC ASCB + DISCRETES LOGIC IN FWCRD CMB-FLT HYD OFF DAU 1,2 JIB-53 WSP 17, BIT 12 28/0 ~RETRIM L WING DOWN FGC WSP38, BIT 8RETRIM RWING DOWN FGC WSP38, BIT 7SNGL RUDDER LIMIT DAU 1 JIB-54 WSP 17, BIT 13 28j0SSEC 1-2 DISABLED DADC WSP 17, BIT 15STAB-FLAP FAIL DAU 1 JIB-55 WSP 17, BIT 14 28~0STALL BARR 1-2 FL DAU 1,2 JIB-38 WSP 16, BIT 13 0/28STALL BARRIER OFF DAU 1 JIB-57 WSP 18, BIT O 28/0STALL BARRIER 1-2 DAU 1,2 JIB-44 WSP 17, BIT 3 28/0STBY PITOT HT FAIL DAU 1 JIB-76 WSP 19, BIT 2STEER BY WIRE FAIL FWC JIA-26 :{: ATAT PROBE HT FAIL DAU 2 JIB-76 WSP 19, BIT 2 G/OTRIM LIMIT FGC WSP38, BIT 1TRU FAIL DAU 1 JIB-58 WSP 18, BIT 1 28j0TRU HOT DAU 2 JIB-50 WSP 17, BIT 9 28/0UTILITY HYD OFF DAU 1 JIB-60 WSP 18, BIT 3 28/0YAW DAMPER OFF FGC ASCB + DISCRETES LOGIC IN FWC

~ Message enable state/nortnaloperation.

~ CCA certification only.

~ GEAR DOWN discrete (FWC JIA-67) must be active ground.

~ Corresponding COWL A/I discrete (DAU 1,2 JIB-18, WSP 15,BIT 9) must be active (28 V) for at least 15 seconds.

~ Corresponding WING A/I discrete (DAU 1,2 JIB-17, WSP 15,BIT 8) must be active (28 V) for at least 2 minutes.

A -905 FWC Only.

CAS Amber Caution MessagesTable 203 (cent)

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AC EXT POWER DAU 2 JIB-46 WSP 17, BIT 5 28/0AP CTLR SW STUCK FGC WSP 39, BIT OAP ENGAGE INHIBIT FGC WSP 39, BIT 4APU ALT OFF FWC JIB-28 28j0APU EXCEEDANCE FWC CALCULATED 5 SECONDAT ENGAGE INHIBIT AT WSP 4, BIT 7AT NOT IN HOLD AT ASCB LOGIC iN FWCAT 1-2 FAIL AT ASCBBC 1-2-3 TEST FAIL FWC CALCULATEDBRAKE MAINT REQD DAU 1 JIB-73 WSP 18, BIT 15 OjGBUS CTLR 1-2-3 FL FWC JIA-97,98,99 O/GCALL DAU 1,2 JIB-15 WSP 15, BIT 6 28/O(OR’D)CAT 2 RAD ALT FWC ASCB LOGIC IN FWCCAT 2 RADIOS FWC ASCB LOGIC IN FWCCAT 2 SG REV FWC ASCB LOGIC IN FWCCDU 1-2-3 FAIL FWC JIA-75,76,77 O/G &CHECKLIST MISMATCH FWC CALCULATEDCMB-FLT HYD HOT DAU 1,2 JIB-22 WSP 15, BIT 13 28;0CPL DATA INVALID FGC WSP 39, BIT 9DADC MISCOMPARE FGC WSP 38, BIT 5DADC 1-2 FAIL DADC ASCBDAU lA-lB-2A-2B FL DAU 1,2 ASCBDAU 1-2 MISCMP-MSG FWC CALCULATEDDC EXT POWER FWC JIB-9 28j0DISP CTLR 1-2 FAIL DC ASCBDU 1-2-3-4-5-6 HOT FWC J1A-34,35,36,37,38,39 G;OE BATT 1-2 DISCH DAU 1,2 JIB-39 WSP 16, BIT 14 28/0E BATT 1-2 FAIL FWC JIA-71,72 O/GENGINE EXCEEDANCE FWC CALCULATED 5 SECONDEPR 1 - DADC 2 DAU 1 WSP 11, BITS 0,1EPR 2 - DADC 1 DAU 2 WSP 11, BITS 0,1EXT BATT SWITCH ON DAU 1 JIB-52 WSP 17, BIT 11 28;0FGC NOT USING IRS 1(2) FGC WSP 3, BITS 14,15FGC SYSTEM TEST FGC WSP 39, BIT 1FGC 1(2) MASTER FGC WSP 1, BIT 11 5 S~CONDFLIGHT REC FAIL FWC JIB-14 0/28FMS 3 ACTIVE FWC JIA-780RJ1A-79 LOGIC IN FWC ~FUEL INTTANK OPEN DAU 2 JIB-68 WSP 18, BIT 10 G/OFUEL XFLOW OPEN DAU 1 JIB-68 WSP 18, BIT 10 G/OFWC 1-2 FAIL FWC ASCBGND PROX FAIL DAU 2 JIB-61 WSP 18, BIT4 28~0GND SPOILER UNARM FWC JIA-70 G/OIRS MISCOMPARE FGC WSP 38, BIT 6IRS MONITOR FAIL FGC WSP 38, BIT 15IRS 1-2 FAIL IRS WSP 33, BIT 4IRS 1-2 HI LATALN IRS WSP 34, BIT 6

CAS Blue Advisory MessagesTable 204

22-14-00Page-285


I Message Source Type ~ I

IRS 1-2 NAV READYIRS 1-2 ON DCISOLATION VLV OPENL-R COWL A/IL-R OIL FILT BPASSL-R WING A/IMAINT SWITCH ONMLS FAILMLS 1 FAILMLS 2 FAILNAV MISCOMP L SELNAV MISCOMP R SELNZ 1-2-3 FAILPERF 1-2 FAILRADALT 1-2 FAILRECENTER TURN KNOBRUDDER LIMITSELECT INHIBITSERVICE DOORSSG 1-2-3 FAILSG 1-2-3 HOTSPD BRAKE EXTNDEDSPD BRAKE SWITCHTCAS FAILTONE GEN FAILTREND RECORDVHF COMM 1-2-3 FLVNAV TRACK CHANGEVOICE REC FAILVOR COURSEWS FAILWS UNAVAILABLE

IRSIRSDAU 1DAU 1,2FWCDAU 1,2FWCSGSGSGFGCFGCNZPERFDAU 1,2FGCDAU 2FGCDAU 2SGFWCDAU 2FWCSGDAU 2FWCFWCNZFWCFGCFWCFWC

WSP 34, BIT 2WSP 34, BIT 3JIB-51 WSP 17, BIT 10 28;0JIB-18 WSP 15, BIT 9 28/0JIB-12,13 0/28JIB-17 WSP 15, BIT8 28/0JIA-85 G/OWSP 6, BIT 5 :&WSP 6, BIT 6WSP 6, BIT 7 -bWSP 39, BIT 6WSP 39, BIT 7ASCB (NZ 3=FWC JIA-87) ~ ~ASCBJIB-19 WSP 15, BIT 10 Oj28WSP 38, BIT 14JIB-54 WSP 17, BIT 13 28j0WSP 39, BIT 5JIB-59 WSP 18, BIT 2 28j0ASCBJIA-51,52,53 G~OJIB-49 WSP 17, BIT 8 28/0ASCB (DAU VS. FGC) LOGIC IN FWCWSP 6, BIT 4JIB-57 WSP 18, BIT O 0j28 ~CALCULATEDJIB-19,20 0j28ASCB LOGIC IN FWCJIA-25 O/GWSP 23. BIT 2JIA-33’JIA-69

~ Message enable state/normal operation.

~ WS INSTALLED discrete (FWC JIA-32) must be active ground.

~ SPARE FMS INSTALLED discrete (FWC JIA-90) required to activateFMS 3 messages. FMS 3 ACTIVE message requires valid NZ 3 andvalid CDU 3.

~ WS INSTALLED (FWC JIA-32) and WS VALID (FWC JIA-33) discretesmust be active ground.

I ~ -905 FWC only. I

CAS Blue Advisory MessagesTable 204 (cent)

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2. F. (2) (d) CASout

Message Scroll - Amber and blue messages can be scrol 1edof the CAS display area providing the pilot with a

decl uttered CAS display. The scroll UP/DOWN buttons, locatedon the master warning/caution panel (Figure 229), provide themeans to do the scroll function.

To scroll the messages up, the following criteria must be met:

. all amber messages acknowledged● all blue messages displayed steady

To scroll the messages down, the following criteria must bemet:

● 23 messages minimum must be displayed● all amber messages acknowledged● all blue messages displayed steady

New amber messages bring the scrolled amber and blue message(up to 23 messages) stacks back onto the CAS display. New bluemessages bring the scrolled blue message stack back onto theCAS display.

(e) CAS Message Status Line - The bottom line of the CAS display isprovided to give a status of the scrolled messages. Thedisplay consists of a total message number with arrow(s) toindicate the direction the message was scrolled. The amberreadout and arrow correspond to the amber messages scrolled offthe display and is located on the left side of the status line.The blue readout and arrow correspond to the blue messagesscrolled off the display and is located on the right side ofthe status line. It is possible to have a total number witharrows pointing in both directions. The number indicates thetotal number of undisplayed messages. The arrows indicate thatsome are scrolled off in the up direction while some arelocated in the down direction.

(f) Message Inhibit Function - The inhibit select switch is locatedon the master warning/caution panel (Figure 229). This mode isactivated by selecting the inhibit switch with either gear downor valid radio altitude less than 400 feet AGL.

The master caution light, amber aural alert, and blue auralalert are disabled whenever the inhibit function is active(except as noted below). Display and acknowledgement of themessages is as previously described.

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The master warn and red aural alert are disabled for any of thefollowing red messages when the inhibit function is active:

● APU FIRE. CABIN DFRN - 9.8. MAIN DOOR. CABIN PRESSURE LOWc DAU 1-2 MISCMP - MSG

All other red messages activate the master warn and red auralalert.

The following amber messages are not inhibited when the inhibitfunction is active:

. CAT 2 INVALID (Option)● CPL DATA INVALID

2. F. (3) CAS Computed System Messages

The majority of the messages displayed by the crew alerting systemare generated directly from logic within a particular aircraftsystem. The messages described in this section are computed by thefault warning computer based on multiple system inputs.

(a)

(b)

(c)

DAU 1-2 MISCMP-MSG Red Message - The fault warning computercompares the red message activation logic between each channel(A and B) of a particular DAU. If a difference is detectedbetween the DAU channels, the miscompare message is activatedas appropriate (i.e., DAU 1, 2, or 1-2).

CHECK PFD Red Message - This message is active anytime theCHECK DU message is active from both the pilot’s and copilot’sprimary flight displays (PFD). The CHECK DU messages, relative

to the PFDs, are suppressed whenever the CHECK PFD message isactive. Refer to the next paragraph for the CHECK DU messagedefinition.

CHECK DU 1-2-5-6 Amber Message - Each display unit (DU)transmits data to the fault warning computer over a private-line data bus. The content of the data varies depending on DUposition and display format. All DUS transmit activity counterinformation. The DUS displaying the PFD format additionallytransmit attitude (pitch and roll), speed (CAS or Mach), andaltitude information. The CHECK DU message is activatedwhenever the conditions for any one of the following criteriaare met:

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. The fault warning computer monitors for activity on the DUwrap-around bus. The CHECK DU message is displayed asappropriate for the DU’S positional location (1 through 6)in the aircraft.

● The fault warning computer compares PFD format data againstactual sensor data. If the difference between the displayand the sensor exceed the appropriate threshold, for any oneparameter, the CHECK DU message is activated as appropriatefor the DU’S positional location in the aircraft.

2. F. (3) (d) ENG/DU4 Message (PFD) - These messages are active on the PFDwhenever the conditions for any one of the following criteriaare met:

. The fault warning computer monitors for act-ivityon the DUwrap-around bus. The ENG and DU4 messages are displayed asappropriate for the DU’S positional location in theaircraft.

. The fault warning computer compares engine format dataagainst the actual sensor data. If the difference betweenthe display and sensor exceed the appropriate threshold, forany one parameter, the ENG annunciation is activated. Ifthis condition exists with compacted EICAS format(paragraph 2.H.) on DU4, the ENG message is displayed andthe DU4 message is suppressed.

(e) DAU 1-2 MISCMP-MSG Amber Message - The fault warning computercompares the amber message activation logic between eachchannel (A and B) of a particular DAU. If a difference isdetected between the DAU channels, the miscompare message isactivated as appropriate (i.e., DAU 1, 2, or 1-2).

(f) DAU 1-2 MISCMP-ENG Amber Message - The fault warning computercompares primary engine data (EPR, TGT, LP, HP, FF) betweenchannels A and B of each DAU. If the difference exceeds anyone of the following thresholds for 2 or more seconds, thismessage will indicate the appropriate DAU: EPR - 0.5,TGT - 200, LP - 20, HP - 20, FF - 1000.

(g) cWX VSPEEDS Amber Message - When the AUTO VSPD mode isactive, the FWC compares the Vspeeds from FMS-PERF 1 andFMS-PERF 2. If the Vspeeds differ by more than 3 knots or onlyone performance computer is outputting valid Vspeeds, themessage is enabled.

(h) DAU 1-2 MISCMP-MSG Blue Message - The fault warning computercompares the blue message activation logic between each channel(A and B) of a particular DAU. If a difference is detectedbetween the DAU channels, the miscompare message is activatedas appropriate (i.e., DAU 1, 2, or 1-2).

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2. F. (4) CAS Display Failure Indications

(a) Fault Warning Computer (FWC) Failure - In the event that thedisplayed FWC fails, all messages and status line data areremoved. A red “X” is drawn through the entire CAS displayarea as shown in Figure 230.

(b) Bus Controller Failures - In the event the three buscontrollers fail, a FWC failure indication is given asdiscussed in the previous paragraph.

(c) EICAS Self-Test - This test is initiated through either displaycontroller (TEST mode) when airspeed is valid and less than60 knots and the WOW switch is active. Selecting the EICASself-test gives the following indications in the cockpit:

c All Red Messages Displayedc AP OFF Horn Activec AP OFF Light On. Red Aural Alert Active. EICAS Fail Discrete Active. Master WARN/CAUTION Lights On

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Honeywell ##&!i!wcE

CAS Display Failure IndicationsFigure 230

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Honeywell !!!!#!R~NcE2. G. System Page Displays

The system mode is comprised of the following system pages:

c Hydraulic System (Figure 231). Fuel System (Figure 233). APU/Bleed System (Figure 235)● Engine Start (Figure 237)● Exceedance System (Figure 239). Checklist (Figure 242)● Waypoint List (Figure 244)● System Page Declutter Mode (Figure 245)

(1) System Page Displays - General

The data acquisition units and fault warning computers, previouslydiscussed, provide all the required data to build the system pagedisplay formats. The system pages are displayed in the right-handtwo-thirds of the CAS display, adjacent to the CAS messages. Theindividual formats, which make up the system mode, can be selectedby either pilot using the display controller as shown in the abovereferenced figures.

(2) Hydraulic System Page

The hydraulic system page (Figure 231) is comprised of the followingfunctions:

Combined Hydraulic Pressure (COMB PRESS) Digital ReadoutFlight Hydraulic Pressure (FLT PRESS) Digital ReadoutCOMB/FLT Hydraulic Valve Status IndicationsUtility Hydraulic Pressure (UTIL PRESS) Digital ReadoutAuxiliary Hydraulic Pressure (AUX PRESS) Digital ReadoutCombined Hydraulic Quantity Analog Scale (Vertical Tape)Flight Hydraulic Quantity Analog Scale (Vertical Tape)Applied Brake Pressure Analog Scales (Vertical Tape)

(a) Combined Hydraulic Pressure (COMB PRESS) - COMB PRESS isdisplayed digitally in white on the left side, first row, ofthe hydraulic system page (Figure 231). The digital readoutresolution is 100 PSI.

(b) F1ight Hydraulic Pressure (FLT PRESS) - FLT PRESS is displayeddigitally in white adjacent to the COMB PRESS readout in thefirst row of the hydraulic system page. The digital readoutresolution is 100 PSI.

(c) COMB/FLT Hydraulic Pressure Valve Status Indications - When theCOMB/FLT JiYDPRESS valves are in the closed position, ambercross-hatch is displayed within the appropriate hydraulicpressure digital readout window (Figure 226). When theCOMB/FLT HYD PRESS valves are in transit, amber dashes aredisplayed within the appropriate hydraulic pressure digitalreadout window (Figure 226). 22-14-00

Page 292Aug 15/91


HYDRAULICS

COMB PRESS FLT PRESS

m mUTIL PRESS AUX PRESS

m m

HYDRAULIC APPLIEDQUANTITY BRAKE PRESS

r(

- FULL

- ADD

- LOW

d<

Hydraulic System PageFigure 231

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2. G. (2) (d) Utility Hydraulic Pressure (UTIL PRESS) - UTIL PRESS isdisplayed digitally in white on the left side, second row, ofthe hydraulic system page (Figure 231). The digital readoutresolution is 100 PSI.

(e) Auxiliary Hydraulic Pressure (AUX PRESS) - AUX PRESS isdisplayed digitally in white on the right side, second row, ofthe hydraulic system page. The digital readout resolution is100 PSI.

(f) Combined Hydraulic Quantity (COMB) - The hydraulic quantityanalog scales (COMB/FLT) are the left set of scales located inthe bottom half of the hydraulic system page (Figure 231). Thecombined HYDRAULIC QUANTITY (left) scale is labeled with awhite COMB. The flight HYDRAULIC QUANTITY (right) scale islabeled with a white FLT.

The white quantity scale is marked with the labels LOW, ADD,and FULL. The triangular shaped scale pointers change colordepending on the indicated quantity range. The colorrelationship is as follows:

RED HYD QTYs LOWWHITE HYD QTY > LOW

(g) Applied Brake Pressure Left/Right (LT/RT) - The applied brakepressure scales are located in the lower right-hand corner ofthe hydraulic system page. The left (LT) and right (RT)applied brake pressure scales are marked with tickmarks every100 PSI below 1000 PSI and every 200 PSI above 1000 PSI. Thescales are labeled in white as follows:

5, 10, 20, and 30

The labels are interpreted as 100 PSI units as labeled at thebottom center of the scale. The data resolution is readable to50 PSI below 1000 PSI and to 100 PSI above 1000 PSI. Thedisplay thermometer is a moving ladder type. The ladder rungsare left at the labeled tickmarks only (i.e., 5, 10, 20, 30) asthe data increases.

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2. G. (3) Hydraulic System Page Failure Indications

DAU Failures - A DAU failure is indicated by removing the datadisplay thermometers and pointers and replacing all the digitalreadouts with amber dashes (Figure 232).

. The above indications are given on the COMB PRESS, UTIL PRESS,COMB quantity, and LT applied brake pressure for a DAU No. 1failure.

● The above indications are given on the FLT PRESS, AUX PRESS, FLTquantity, and RT applied brake pressure for a DAU No. 2 failure.

(4) Fuel System Page Displays

The fuel system page (Figure 233) is comprised of the followingfunctions:

● LeftlRight Fuel Flow (FF) Digital Readout● Fuel Flow Valve Status Indicationss Left/Right Engine Fuel Temperature (ENG FUEL TEMP) Digital

Readout● Engine Fuel Temperature Warning Indications. Fuel Tank Temperature (FUEL TANK TEMP) Digital Readout● Fuel Tank Temperature Warning Indications. Fuel Quantity Digital Readouts

(a) Fuel Flow (FF) Digital Scale - The FF digital readouts arelocated at the top, first row, of the fuel system page. Thewhite FF label is located between the left and right digitaldisplay windows. The white current value readouts are readableto a 10 pounds per hour (PPH) resolution.

(b) Fuel Flow Valve Status Indications - When the fuel flow valvesare in the closed position, amber cross-hatch is displayedwithin the FF digital readout window (Figure 226). When thefuel flow valves are in transit, amber dashes are displayedwithin the FF digital readout window (Figure 226).

(c) Engine Fuel Temperature (ENG FUEL TEMP) Display - The ENG FUELTEMP digital readouts are located in the second row of the fuelsystem page. The white ENG FUEL TEMP label is located abovethe left and right digital display windows. The white currentvalue readouts are readable to a 1 “C resolution. The displaydigits change color as a function of fuel temperature asdefined below:

RED FUEL TEMP < -40 “CWHITE -40 “C s FUEL TEMP < 90 “CAMBER 90 “C s FUEL TEMP< 120 “CRED 120 “C: FUEL TEMP

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Honeywell !ff!g~.c’

1

HYDRAULICS

COMB PRESS F$T PREYS

UTIL PRESS AUX PRESS

HYDRAULICtlUANTITY

– FULL–

–ADD –

– LOW —

APPLIEDBRAKE PRESS

mmEl

Hydraulic System Page Failure IndicationsFigure 232

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ENG FUEL TEMP

mm

FUEL TANK TEMP

mFUEL QTY

+%’

EElm

Fuel System Page DisplayFigure 233

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Honeywell !k!!!r.”2. G. (4) (d)

(e)

(5) Fuel

(6)

Fuel Tank Temperature (FUEL TANK TEMP) Display - The FUEL TANKTEMP digital readouts are located in the third row of the fuelsystem page. The white FUEL TANK TEMP label is located abovethe digital display windows. The white current value readoutshas a 1 “C resolution. The display digits change color as afunction of tank temperature as defined below:

RED TANK TEMP c -40 ‘CWHITE -40 “C s TANK TEMP < 54 “CRED 54 “C s TANK TEMP

Fuel Quantity (FUEL QTY) - The FUEL QTY display consists ofthree digital windows (comprising rows 4 and 5). Two windows(row 5) corresponds to the fuel quantity in each tank (L/R).The third window (row4) is a running total (T) of the fuelquantity from both tanks. All three digital displays are white(normal) and readable to a 50-pound resolution. The displaydigits go to amber when a Fuel Level Low indication isreceived. The total window turns amber only if both sidesindicate fuel level low.

System Page Failure Indications

DAU Failures - A DAU failure is indicated by replacing the currentvalue digits with amber dashes (Figure 234).

. The above failure indications are given on the FF (left), ENGFUEL TEMP (left), FUEL TANK TEMP, and FUEL QTY (left) displaysfor a DAU 1 failure.

. The above failure indications are given on the FF (right), ENGFUEL TEMP (right), FUEL TANK TEMP, and FUEL QTY (right) displaysfor a DAU 2 failure.

APU/BLEED System Page

The APU/BLEED system page (Figure 235) is comprised of the followingfunctions:

● Analog APU EGT Scale With Digital Readout. Analog APU RPM Scale With Digital Readout. Left Bleed Air Pressure Digital Readout● Right Bleed Air Pressure Digital Readout

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FUEL

ENG FUEL TEMP

FUEL TANK TEMP

FUEL QTY

+%’

Fuel System Page Failure Indications

Figure 234

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)t?mEE&lrl -(iHYD ENS START

E%CEEDANCES

~ UAYPT LIST

APU/BLEEDAPU

EGT

m

b

BLEED AIR

RPM

PRESS

APU/BLEED System Page DisplayFigure 235

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Use or disclosure of information on this page is subject to the restrmtions on the title page of this document,

2. G. (6) (a) APU EGT Display - The EGT scale is the left scale at the top ofthe display under the white label APU (Figure 235).

The EGT arc is banded with white, amber, and red colors toindicate normal, caution, and warnin ran es respectively. The

8 “?.arc boundaries range between O and 1 00 The color rangedefinition for APU EGT is as follows:

WHITE O ‘Cg EGT < 680 “C&l~ER 680 “C< EGT < 732 “C

732 “C< EGT < 1000 ‘C

The EGT current value readout corresponds to the pointer

!osition on the scale and is readable to a 1 ‘C resolution.f the actual data goes above 1000 ‘C, the pointer remainsstationary at the upper arc limit and the digits show theactual EGT value.

APU EGT exceedances are shown by growing the colored band totwice the normal width and than ing the pointer and current EGT

!di ital value color (normally w ite) to match the arc region7co or.

(b) APU RPM Dis lay -!

The RPM scale is the right scale at the topof the disp ay under the label APU (Figure 235).

The RPM arc is banded with white, amber, and red colors toindicate normal, caution, and warnin ran es respectively. The

!!arc boundaries range between O and 1 O% R M. The color rangedefinition for APU RPM is as follows:

WHITE O% RPM s RPM < 104% RPM&4~ER 104% RPM s RPM < 110% RPM

110% RPM s RPM < 120% RPM

The RPM current value readout corresponds to the pointerposition on the scale and is readable to a 1% RPM resolution.If the actual data goes above 120% RPM, the pointer remainsstationary at the upper arc limit and the digits show theactual RPM value.

APU RPM exceedances are shown by growing the colored band totwice the normal width and than ing the pointer and current RPM

!di ital value color (normallyw ite) to match the arc region7co or.

(c) Left Bleed Air Pressure Digital Readout - Directly below theAPU EGT arc is the left bleed air ressure di ital readout.

k’ !The white digital readout is reada le to a 1 S1 resolution.

(d) Right Bleed Air Pressure Digital Readout - Directly below theAPU EGT arc is the right bleed air ressure di ital readout.

Y !The white digital readout is readab e to a 1 P I resolution.

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2. G. (7) APU/BLEED System Display Indications

. DAU Failures - A DAU failure is indicated on the arc displays byremoving all color bands and data pointers from the arc andreplacing the current value di its with amber dashes.

7The

indication shown on the digita readout is that of replacing thedigital readout with amber dashes (Figure 236).

The above indications are given on the APU EGT, APU RPM, and leftbleed air pressure displays for a DAU 1 failure. The failureindications are given on the right bleed air pressure display fora DAU 2 failure.

(8) Engine Start Page (ENGINE START) Display

The ENGINE START system page display consists of the followingfunctions and annunciations (Figure 237).

. Left Bleed Air Pressure Digital Readout● Right Bleed Air Pressure Digital Readout. Annunciations of Fuel Pressure Low. Annunciations of Ignition (IGN)● Analog TGT Scales. Analog HP Scales● Di ital Readouts of Oil Pressure

7. Oi Pressure Warning Indications. Sin le Rudder Limit Annunciation

i!● Cornined Hydraulic Pressure (COMB HYD PRESS) Digital Readout● Fli ht Hydraulic Pressure (FLY HYD PRESS) Digital Readout

8. COM /FLT Hydraulic Pressure Valve Status Indications● Utility Hydraulic Pressure (UTIL HYD PRESS Digital Readout

4s Auxiliary Hydraulic Pressure (AUX HYD PRES ) Digital Readout

The ENGINE START format is available for on-ground use only. Thedisplay controller ENG START prompt is removed when weight-off-wheels is detected.

The ENGINE START format (Figure 237) shown is divided in half torepresent the left and right side engine parameters. The format isdivided into five functional rows as follows:

Row 1 - ;~~cm~stsof the left and right.

Row 2 - &sts of the left and right.

Row 3 - ~~:;~sts of the left and right.

Row 4 - Consists of the left and right

bleed air pressure readout

fuel pressure annunciation

SVO and IGN annunciation

analog TGT and HP scales.

Row 5 - Consists of the left oil pressure readout box, the SING RUDLIMIT annunciation box, and the right oil pressure readoutbox.

ROW 6 - Consists of the COMB, FLIGHT, UTIL, and AUX hydraulicpressure readouts.

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APWBLEEDAPU

BLEED AIR PRESS

E23

APU/BLEED System Page Failure IndicationsFigure 236

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ENGINE START.BLEED AIR 40.FUEL PRESS LOU

L SVO

I

700

600

Soo

40

300

200

100

L lGN

I

60

50

40

30

20

10

LTGTILHP

?BLEED AIR 40?FUEL PRESS LOU

? Svo

I700

600

500

40

300

200

100

50

40

J30

20

10

R TGT R HP

L OIL 35 SING RUD R OIL 35PRESS LIMIT PRESS

HYDRAULIC PRESSURE

COMB FLIGHT UTIL AUX3000 3000 3000 3000

Engine Start Page DisplayFigure 237

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2. G. (8) (a) Bleed Air Pressure Digital Readouts (RBLEED/LBLEED AIR) - TheRBLEED AIR and LBLEED AIR digital readouts comprise the firstrow of displays on the ENGINE START display page. The whitedigital readouts are readable to a 1 PSI resolution.

(b) Fuel Pressure Low Annunciations - The LFUEL PRESS LOWannunciation is located in the box directly below the leftbleed air pressure digital readout. The RFUEL PRESS LOWannunciation is located in the box directly below the rightbleed air pressure digital readout. These annunciations aredisplayed when a fuel pressure low condition is detected.

(c) Start Valve Open (SVO) Annunciation - The SVO annunciationsare located in the third row of data. The first box houses theL SVO annunciation while the third box houses the R SVOannunciation. The SVO annunciation changes color inrelationship to HP9ARPM as described in paragraph 2.E.(2)(primary engine instruments).

(d) Ignition (IGN) Annunciation - The green IGN annunciations arelocated in the third row of data. The second box houses theL IGN annunciation while the fourth box houses R IGNannunciation.

(e) Left Turbine Gas Temperature (L TGT) Display - The L TGT s~~~~is the first scale in the (fourth) row of analog scales. “display is a low end expansion of the TGT round dial gauge onthe primary engine instruments. The white scale is markedevery 100 “C and ranges between 100 and 700 “C. A set-pointtriangular mark is shown at 425 “C. A white label L TGTappears at the scale bottom. The data display pointer is amoving ladder type thermometer. As the TGT value increases,the thermometer grows up the scale. A ladder rung is left ateach labeled tickmark. The normally white thermometer turnsred for TGT values greater than 700 “C.

(f) Right Turbine Gas Temperature (R TGT) Display - The RTGT scaleis the third scale in the (fourth) row of analog scales. Thisdisplay is a low end expansion of the TGT round dial gauge onthe primary engine instruments. The white scale is markedevery 100 ‘C and ranges between 100 and 700 ‘C. A set-pointtriangular mark is shown at 425 “C. A white label RTGTappears at the scale bottom. The data display pointer is amoving ladder type thermometer. As the TGT value increases,

the thermometer grows up the scale. A ladder rung is left ateach labeled tickmark. The normally white thermometer turnsred for TGT values greater than 700 “C.

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2. G. (8) (g) Left High Pressure (L HP) Tach Display - The L HP scale is thesecond scale in the (fourth) row of analog scales. Thisdisplay is a low end expansion of the HP round dial gauge onthe primary engine instruments. The white scale is markedevery 10% RPM and ranges between O and 60% RPM. A set-pointtriangular mark is shown at 15%RPM. A white label L HPappears at the scale bottom. The white data display pointer isa moving ladder type thermometer. As the HP value increases,the thermometer grows up the scale. A ladder rung is left ateach labeled tickmark.

(h) Right High Pressure (R HP) Tach Display - The L HP scale is thefourth scale in the (fourth) row of analog scales. Thisdisplay is a low end expansion of the HP round dial gauge onthe primary engine instruments. The white scale is markedevery 10% RPM and ranges between O and 60% RPM. A set-pointtriangular mark is shown at 15% RPM. A white label R HPappears at the scale bottom. The white data display pointer isa moving ladder type thermometer. As the HP value increases,the thermometer grows up the scale. A ladder rung is left ateach labeled tickmark.

(i) Left Engine Oil Pressure (L OIL PRESS) Display - The L OILPRESS digital readout is located in the first box of the fifthrow of the ENGINE START format. The normally white digits arereadable to a 1 PSI resolution. The display digits changecolor as a function of oil pressure as defined below:

RED O PSI <OIL PRESS< 16 PSIAMBER 10 PSI gOIL PRESS < 30 PSIWHITE 30 PSI sOIL PRESS

(j) Single Rudder Limit (SING RUD LIMIT) Annunciation - The amberSING RUD LIMIT annunciation appears in the center box of thefifth row whenever this condition is active.

(k) Right Engine Oil Pressure (ROIL PRESS) Display - The ROILPRESS digital readout is located in the third box of the fifthrow of the ENGINE START format. The normally white digits arereadable to a 1 PSI resolution. The display digits changecolor as a function of oil pressure as defined below:

RED O PSI sOIL PRESS< 16 PSIAMBER 10 PSI <OIL PRESS < 30 PSIWHITE 30 PSI <OIL PRESS

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2. G. (8) (1) Hydraulic Pressure Readouts

The engine start display has digital readouts of the followinghydraulic pressures:

. Combined Hydraulic Pressure (COMB)

. Flight Hydraulic Pressure (FLIGHT)

. Utility Hydraulic Pressure (UTIL)c Auxiliary Hydraulic Pressure (AUX)

The operation of these displays is discussed in paragraph2.E.(3).

(9) Engine Start Page Failure Indications

DAU Failures - A DAU failure is indicated by removing the datadisplay thermometers and replacing all the digital readouts withamber dashes.

● The above indications are given on the LBLEED AIR, L TGT, L OILPRESS, COMB PRESSURE, and UTIL PRESSURE displays for a DAU 1failure. The LFUEL PRESS LOW, L SVO, L IGN, and SING RUD LIMITannunciations won’t activate in this condition. Figure 238 showsa DAU No. 1 failure indication.

● The above indications are given on the RBLEED AIR, RTGT, ROILPRESS, FLIGHT PRESSURE, and AUX PRESSURE displays for a DAU No. 2failure. The RFUEL PRESS LOW, R SVO, and R IGN annunciationswill not activate in this condition.

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Honeywell !f!Ii##&.cE

I ENGINE START

LBLEED AIR -- RBLEED AIR 40

RFUEL PRESS LOW

R SVO R IGN60

50

50 40

408 30 406 – – 30

– 208 a1 – 10

0lLTGTILHpl RTGTl RHp

L OIL ---- R OIL 35

PRESS PRESS

HYDRAULIC PRESSURE

I COMB FLIGHT UTIL AUX---- 3000 ---- 3000

Engine Start Page Failure IndicationsFigure 238

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2. G. (10) Engine/APU Exceedance (EXCEEDANCES)Display

The EXCEEDANCES system page display consists of the followingfunctions and annunciations (Figure 239).

. No exceedances recorded message

● Engine exceedance monitor with the following functions:

- Snapshot recording of maximum exceedance value for TGT,LP, and HP

- Recording of time the parameter was in exceedance band- Indication of which parameter tripped the exceedance monitor

● APU exceedance monitor with the following functions:

- Snapshot recording of maximum exceedance value for APU EGTand RPM

- Indication of which parameter tripped the exceedance monitor

Selecting the EXCEEDANCES system page prior to an exceedance causesthe NO EXCEEDANCE RECORDED message to be displayed in white acrossthe system page display field as shown in Figure 240.

When an engine exceedance is encountered, the blue ENGINE EXCEEDANCEEICAS message is displayed. Subsequent selection of the EXCEEDANCESsystem page will show the parameter that caused the engineexceedance on the TRIGGERED BY line. The highest value of all ~parameters is recorded on the display. The time clock is startedonly if the appropriate parameter is above the exceedance threshold.The exceedance thresholds are set in accordance with the G-IV flightmanual. The clock is stopped when the actual value drops belowthe lower hysteresis trip threshold, (LP< 95%, HP<97%, and TGT< 710 “c). Once all values drop below the exceedance thresholds,the display will hold the last exceedance data until the power isremoved or another exceedance is detected.

The APU exceedance monitor works in the same manner except that ablue APU EXCEEDANCE EICAS message is displayed and time is notrecorded.

All exceedances are held in the fault warning computer memory andcan be retrieved by maintenance personnel as needed.

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APu/Bl&ED UAYPT LIST

CHECKLISTL d

TG

LP

HP

EXCEEDANCES

ENGINE EXCEEDANCEMAX-TIME

800 0:49 750 0:0049.5 0:00 49.5 0:0080.3 0:00 80.3 0:00

TRIGGERED BY: L TGT

EGT ~

RPM

APU EXCEEDANCEMAX

0000000.0

TRIGGERED BY N/A

Engine/APU Exceedances PageFigure 239

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EXCEEDANCES

NO EXCEEDANCES RECORDED

No Exceedances Recorded FormatFigure 240

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Honeywell ##!b.cE2. G. (11) EXCEEDANCES System Page Failure Indications

(a) Fault Warning Failure - In the event that the FWC fails, amagenta EXEEDANCES UNAVAILABLE message is displayed in thesystem page display field as shown in Figure 241.

(b) Bus Controller Failures - In the event that the three buscontrollers fail, an EXCEEDANCES UNAVAILABLE message isdisplayed (Figure 241).

(12) CHECKLIST System Display

The CHECKLIST display (Figure 242) is comprised of the followingfunctions:

. 12 lines of text max per page

. 26 characters max per line● Normal/abnormal/emergency procedures display. Auto call-up of emergency procedures

The CHECKLIST is displayed when selected by either pilot on thedisplay controller. The data which makes up the checklist isProvided b.ythe displayed FWC. A typical CHECKLIST format is madeup

●

●

●

●

●

(a)

of the ~ollowing” parameters: ‘“

Normal/emergency/abnormal labelCHECKLIST cursorCHECKLIST procedure titleProcedure page indexProcedure step description

Normal CHECKLIST Procedures - The power-up default mode of theCHECKLIST is normal procedures. When selected on the displaycontroller, the normal procedure CHECKLIST index is displayed(Figure 242). In this mode, the normal procedure label andpage index are displayed in white at the top of the index list.The index contains those procedures which make up the normaloperation section of the flight manual. The items which makeup the normal procedures CHECKLIST are numerically listed, inblue, by procedure, in correspondence with the aircraft flightmanual.

When a particular procedure is selected, the CHECKLIST indexdisplay is replaced by the selected procedure list. The pageis labeled in white with the selected procedure located inplace of the normal procedure label already discussed. Eachstep of the procedure is numerically displayed in blue. Whenall steps of the procedure are completed, the displayautomatically returns to the normal procedure index with thecompleted procedure shown in green.

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EXCEEDANCES

EXCEEDANCES UNAVAILABLE

Exceedance Data Failure IndicationsFigure 241

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Honeywell !!!$!%H.CE

PROCEDURE PAGELABEL \ INDEX , CURSOR

\ c“”CKLIST

NORMAL

‘/PROCEDURES

1/2

3 STARTINGENGINES

4 AFTER STARTINGENGINES

5 J~[W~~FORE

6 LINE UP

Checklist System Page DisplayFigure 242

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2. G. (12) (b) Emergency/Abnormal CHECKLIST Procedures - The emergency andabnormal selections call up procedures which make up theemergency and abnormal sections of the aircraft flight manual.The CHECKLIST functions identically to the already described,normal procedures, except that the procedure labels and pageindex are displayed in amber.

(c) CHECKLIST Cursor - CHECKLIST cursor control is accomplishedthrough the DC-884 Display Controller as shown in Figure 243.The white cursor can be controlled forward (ADV) or backwards(BACK) by either PAGE or LINE as shown in the figure. Thecursor box surrounds the entire procedure under consideration.

(d) CHECKLIST Procedure Checked - To check off a 1ine item orselect a procedure from the index, the CHECKLIST cursor ismoved over the appropriate item an the ENTER key is selected.This causes the selected line item to change color from blue togreen giving the crew an indication that this procedure iscomplete.

(e) RECALL - Selection of this key will call up the first (bynumerical designation) CHECKLIST item skipped (i.e., notentered).

(f) CHECKLIST Mode Exit - To get out of the CHECKLIST mode, adifferent system page must be selected.

(g) Auto Callup Of Emergency Procedures - Emergency checklistprocedures are called up for display automatically whenever anyone of the following conditions have been detected:

c Engine hot. Engine fire● Reverser unlock. APU fire. Cabin pressure low

Additionally, the display controller checklist menu is selectedautomatically with the EMER mode active for any of the aboveconditions.

There are two priorities of auto callup of emergencychecklists. Engine fire is the highest priority. All otheremergency checklists are of equal priority and lower thanengine fire. From the normal checklist any auto callup ofemergency checklist will display that checklist. Engine firewill be displayed in place of a previously called up checklist.If a low priority auto callup checklist is displayed, anotherequal priority auto callup checklist will not be displayeduntil the current emergency checklist is completed.

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CHECKLIST

NORMAL 1/2PROCEDURES

1 DISCLAIMER[2---:~~i~~T1RT7~G– j-— --- ---—. --——

‘-3i

STARTINGENGINES

II

4 AFTER STARTINGENGINES

5 TAXI/BEFORETAKEOFF

6 LINE UP

II

—~3 LINE At tiANCES

Checklist System Cursor ControlFigure 243

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2. G. (13) CHECKLIST Failure Mode Indications

Fault Warning Computer Failure - In the event that the FWC failswhen displaying a checklist format, the CAS display shows a red “X”and the system page displays a magenta CHECKLIST UNAVAILABLE messagebelow the checklist system page title.

(14) Waypoint Listing

A waypoint listing of the active flight plan can be displayed whenselected on the display controller as shown in Figure 244. Thewaypoint list consists of 12 lines defined as follows:

. FMS Source Annunciation with IDENT, LAT, and LON Labels (Line 1)

. Yellow FROM Waypoint Data (Line 2)

. Magenta TO Waypoint Data (Line 3)c Next 8 Flight Plan Waypoints Data (Lines 4-11)c Present Position (PPOS) Data (Line 12)

Each waypoint is displayed with IDENT and corresponding latitude(LAT) and 1ongi tude (LON).

(15) Waypoint List Failure Mode Indications

FMS Failure, No Active Flight Plan - In the event that the pilot’sdisplayed FMS fails, all the data in the waypoint list is replacedwith amber dashes. The same indication is given for unused lines inthe waypoint list (Figure 244).

(16) System Page Declutter Mode

The declutter mode removes all system pages from the system displayarea as shown in Figure 245.

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/

WAYPT LIST

I DENT LAT LON FKSAV N 32 07.6 U 81 12SAV N 32 09.6 U 81 06DBN N 32 33.6 U 82 50MCN N3241.5 U8338ATL N 33 37.8 U 84 26KATL N 33 38.4 U 84 2STO+5 -—- -- - ---- -.TO+6 ---- -- - ---- --To+7 ---- -- - ---- --TO+8 ---- -- - ---- --PPOS N 32 07:8 U 81 12

L

Waypoint List Display PageFigure 244

FMSSOURCE

UNUSED‘ LINES

7RESENT3osITIoN

22-14-00Page 298.20Aug 15/91


f -

CABIN PRESSURE LOWBAGGAGE DOORL-R FUEL PRES LOWAOA HEAT 2 FAILL-R AIL HYD OFFL-R PITOT HT FAILBATT 1-2 CHGR FAILL-R AC POWER FAILICE DETECTEDTAT PROBE HT FAILTONE GEN FAILVOICE REC FAILFLIGHT REC FAILAC EXT POWERDC EXT POWERVHF COMM 1-2 FAILISOLATION VLV OPENSERVICE DOORSL OIL FILT BPASSR COWL A/IE BATT 1 FAILAPU ALT OFFCPL DATA INVALID

TIT $3$

L d

System Page Declutter ModeFigure 245

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2. H. Compacted EICAS Display

The compacted EICAS display format is only displayed in an engine or CASsystem reversionary mode. This paragraph addresses the display formatonly. Paragraph 2.1. discusses reversionary mode operation of thedisplay system in detail.

(1)

Analog EPR Scales with Digital ReadoutFMS (PZ) Computed EPR Limit/Target BugsFMS (PZ) Target Bug Digital ReadoutFMS Source/EPR Mode AnnunciationsManual EPR Limit Set CapabilityAnalog TGT Scales with Digital ReadoutAnalog LP Scales with Digital ReadoutLP SYNC AnnunciationLP Anti-Ice AnnunciationsDigital HP ScalesHP SYNC AnnunciationAnnunciations of Start Valve Open (SVO)Annunciations of Ignition (IGN)Digital Readouts of FFFuel Flow Valve Status IndicationsDigital Readouts of Oil PressureOil Pressure Warning IndicationsDigital Readouts of Oil TemperatureOil Temperature Warning IndicationsCrew Alerting System Messages

Compacted EICAS - General

The compacted EICAS display is formed by combining primary enginedata parameters with the CAS system on one display format. Theengine instruments (ENG) comprise the left two-thirds of thisformat. The crew alerting system (CAS) utilizes the right one-thirdof the display format. The ENG and CAS display description isprovided in paragraphs 2.E. and 2.F., respectively.

The engine instrument area is divided into seven rows of enginedisplays as follows:

● EPR Analog Gauges● TGT Analog Gauges● LP Tach Analog Gauges. HP Tach Digital Readout. FF Digital Readout“ Oil Pressure (P) Digital Readout. Oil Temperature (T) Digital Readout

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Each engine gauge is a 225-degree arc with moving pointer and fixedcurrent value window located above the pointer rotation point. Eachengine gauge type is labeled in white between the left and rightgauges as shown in Figure 246.

2. H. (2) Engine Instruments (ENG)

(a) Engine Pressure Ratio (EPR) Display - The EPR instruments(left/right) are the top row of scales on the engine instrumentside of the display format (Figure 246). The details of theEPR scale operation are discussed in paragraph 2.E.(2)(a).

(b) EPR Limit/Target Readout and Bug, Source and Mode Annunciations- Operation of this function is identical to the EPR commanddescription provided in paragraph 2.E.(2)(b).

(c) Turbine Gas Temperature (TGT) Display - The TGT instrumentsform the second row of scales on the engine instrument side ofthe display format. The operational description of the TGTfunction is provided in paragraph 2.E.(2)(g).

(d) Low Pressure (LP) Tach Display - The LP instruments comprisethe third row of gauges on the primary engine display. Theoperational description of the LP tach function is provided inparagraph 2.E.(2)(h).

(e) Anti-Ice (A/1) Annunciator - This annunciation is displayed ingreen above the LP digital readout windows. The annunciationlabel isA/I.

(f) LP Synchronization (SYNC) Annunciator - If the crew chooses tosync the engines using LP, a SYNC annunciation is displayed ingreen directly below the LP scale label.

(9) High pressure (HP) Tach Display - The HP instruments form thefourth row of engine instruments on the format. This displaydiffers from the engine display format in that only a digitalreadout of HP is provided. The HP digital readout colors aredependant on actual HP values as follows:

RED O%RPMS HP < 46.7% RPMWHITE 46.7% RPMs HP < 97.5% RPMAMBER 97.5% RPM~ HP < 99.7% RPMRED 99.7% RPMs HP < 110% RPM

The HP digital readout is readable to a 0.1% RPM resolution.

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.

<Ml FMS1EPRMCT

TGT

LP

CAB IN PRESSURE LOWBAGGAGE DOOR

L-R FUEL PRESS LOWAOA HEAT 2 FAILL-R AIL HYD OFFL-R PITOT HT FAILBATT 1-2 CHGR FAIL

L-R AC POWER FAIL

ICE DETECTEDTAT PROBE HT FAIL

TONE GEN FAILVOICE REC FAIL

FLIGHT REC FAILAC EXT POWER

DC EXT POWERVHF COMM 1-2 FAIL

ISOLATION VLV OPENSERVICE DOORS

L OIL FILT BPASSR COWL A/IE BATT 1 FAILAPU ALT OFF

CPL DATA INVALID

‘rI’r &3&

Compacted EICAS Display FormatFigure 246

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2. H. (2)

(3)

(4)

(h)

(i)

(j)

(k)

(1)

(m)

Start Valve Open (SVO) Annunciation - The SVO annunciation islocated adjacent (outboard side) to the HP digital readout(Figure 246). When the start valve is open, the SVO labelcolor is determined by the following logic:

Condition Color

HP < 42% B1ueHP > 42% Flashing Amber

Ignition (IGN) Annunciation - The green IGN annunciation islocated positionally below the SVO annunciation.

HP Synchronization (SYNC) Annunciator - If the crew chooses tosync the engines using HP, a SYNC annunciation is displayed ingreen directly below the HP scale label.

Fuel Flow (FF) Display - The FF instruments form the fifthrow of engine instruments on the format. This display operatesidentically to the FF digital readout discussed in paragraph2. G.(4)(a).

Engine Oil Pressure (P) Display - The engine oil pressure (P)digital readouts comprise the sixth row of the engineinstrument display format area. The operational description isprovided in paragraph 2.E.(3)(a).

Engine Oil_Temperature (T) Display - Theengine_oil temperature(T) digital readouts comprise the seventh row of the engineinstrument display format area. The operational description isprovided in paragraph 2.E.(3)(b).

Crew Alerting System (CAS) Display

The operational description for this system is provided in paragraph2.F. The only difference between the CAS display on this format andthe normal CAS display (paragraph 2.F.) is the positional locationof the display. The normal CAS display is located on the left one-third of the display format. The compacted EICAS-CAS display islocated on the right one-third of the display format.

Compacted EICAS Failure Mode Indications

The failure mode indications for the compacted EICAS, shown inFigure 247, represent a DAU No. 1 and FWC failure. The failureindication descriptions for each of the above parameters arecontained in paragraph 2.E.(4) for ENG and paragraph 2.F.(4) forCAS.

22-14-00Page 298.25Aug 15/91


-----

LJ0EPR

TGT

LP

Compacted EICAS Failure Mode IndicationsFigure 247

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2. I. EFIS/EICAS Reversionary Modes

This section goes into detail on the various reversionary modes availableto the crew in the EFIS and EICAS systems.

(1) Display System Reversionary Panel

The display system reversionary panel is located in the overheadbetween the pilot and copilot. The reversionary panel layout isshown in Figure 248. The display system reversionary (REV) modesare divided into display unit (DU) REV modes and symbol generator(SG) REV modes. These modes are discussed in more detail in thefollowing paragraphs.

(2) DU and SG Failures - General

(a) DU Failure Mode Indications - The failure mode indication for aDU failure is a blank display. This may be caused by a coolingair failure or an internal DU failure. A cooling air failuregives a clue that the display is going down. When a DUoverheats, the raster (PFD - ADI sky blue, ADI ground brown,and altitude and speed gray shading/ND - WX returns) isremoved. The next indication is a blank display.

(b) SG/Interface Mode Indications - An SG or interface failure isindicated by displaying a red “X” over the entire DU displayarea as shown in Figure 249.

(3) DU Reversionary Modes

(a) Pilot’s DU Reversionary Modes - The pilot’s PFD reversion is totransfer the PFD format onto the NAV display. This is done bymoving the pilot’s DU reversionary switch to the PFD XFERposition. The display configuration for this reversionary modeis shown in Figure 250.

In the event the pilot’s ND fails, the reversionary mode is toturn the DU reversionary switch to the ND OFF position.

(b) Copilot’s DU Reversionary Modes - The copi1ot’s DU reversionarymodes (Figure 251) are identical to those discussed in theprevious paragraph except the copilot uses the right-most setof DU REV switches.

22-14-00Page 298.27Aug 15/91


Honeywell fl!~.c’

➤ �� ✍ �� ✎�� 11 II II Ii ~— DISPLAY SWITCHING -1 Ii II PILOT E I CAS COPI LOT II I

I II II ~SYMBOL GENERATOR CONTROL- II II I

I II IL ——-— ———— ———— ———— ———— ———— ———— ———— -1

Display System Reversionary PanelFigure 248

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Symbol Generator Failure Mode IndicationFigure 249

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❑F COPILOTPFD

Pilot’s PFD Reversionary ModeFigure 250

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Use or disclosure of information on this page is subject to the restrictions on the title page of thw document.

MAINTENANCE

Honeywell %W%f!hw

PILOT’S DC

(q-l

COPILOT’S DC

m~

ENGINE

CASSYSTEM

Copilot’s PFD Reversionary ModeFigure 251

22-14-00Page 298.31Aug 15/91


2. I. (3) (c) EICAS DU Reversionary Modes

In the event that the ENGINE DU fails, the compacted EICASformat (paragraph 2.H.) is selected for display on the CAS DUas shown in Figure 252. This is achieved by selecting theEICAS DU REV switches to the BOT CMPT position.

In the event that the CAS DU fails, the compacted EICAS format(paragraph 2.H.) is displayed on the ENGINE DU as shown inFigure 253. This is achieved by selecting the EICAS DU REVswitches to the TOP CMPT position.

(4) Symbol Generator Reversionary Modes

(a) Symbol Generator 1 (Pilot’s Normal) Failure - As previouslynoted, a failure of SG No. 1 would be shown in one of thefollowing manners:

● Red “X” on the pilot’s PFD. Red ‘lX”on the pilot’s ND. Red “X” on both the pilot’s PFD and ND

The reversionary mode for this failure is to select the pilotSG reversionary knob from the NORM to the ALT position. The SGdriving the EICAS at that moment, normally SG No. 3, would thendrive the pilot’s PFD and ND as well as the EICAS displays. IfSG No. 3 control was selected to the alternate position priorto the SG No. 1 failure, then SG No. 2 would be in use by allsix DUS and a multiple SG failure condition would exist (referto paragraph 2.1.(4)(d)).

(b) Symbol Generator 2 (Copilot’s Normal) Failure - A failure of SGNo. 2 would be shown in one of the following manners:

. Red “X” on the copilot’s PFD

. Red “X” on the copilot’s NDc Red “X” on both the copilot’s PFD and ND

The reversionary mode for this failure is to select the copilotSG reversionary knob from the NORM to the ALT position. The SGdriving the EICAS at that moment, normally SG No. 3, would thendrive the copilot’s PFD and ND as well as the EICAS displays.If SG No. 3 control was selected to the alternate positionprior to the SG No. 2 failure, then SG No. 1 would be in use byall six DUS and a multiple SG failure condition would exist(refer to paragraph 2.1.(4)(d)). For this condition, theinitial SG No. 2 failure mode indication would show the red “X”on the EICAS display(s) as well as the indications listedabove.

22-14-00Page 298.32Aug 15/91


2. I. (4) (c) Symbol Generator 3 (EICAS Normal) Fai 1ure

A failure of SG No. 3 would show itself in one of the followingmanners:

. Red X on the ENGINE display● Red X on the CAS display. Red X on both the ENGINE and CAS displays

The reversionary mode for this failure is to select the EICASSG reversionary knob from the NORM to the ALT position. The SGdriving the copilot’s PFD and ND at that moment, normally SGNo. 2, would then drive the EICAS displays as well as thecopilot’s PFD and ND. If SG No. 2 control was selected to thealternate position prior to the SG No. 3 failure, then SG No. 1would be in use by all six DUS and a multiple SG failurecondition would exist (refer to paragraph 2.I.(4)(d)).

(d) Multiple SG Failures - A single SG is capable of driving thesix DUS in the G-IV EFIS/EICAS system in the event that any twoSGS fail. The pilot’s and copilot’s PFD and NDwould beidentical in this case. All same source indications (ambersource annunciations) would be on each display. The sinqle SGreverts to the pilot~s DC for display command;which SG is driving the six DUS. If the pilotsingle SG reverts to the copilot’s DC for disp”

(5) Display Controller Failures

regardles~ ofs DC fails, theay commands.

(a) Pilot’s DC Fails - In the event that the pilot’s DC fails, theSG driving the pilot’s displays automatically revert to thefollowing display source selections:

~ @

IRS 1 MAP ModeDADC 1 FMS 1NAV 1 ID WAYPTRADALT 1RAD ALT SET - 200BRG - AUTOBARO - INFD CMD - SC

22-14-00Page 298.33Aug 15/91


2. I. (5) (b) Copilot’s DC Fails

In the event that the copilot’s DC fails, the SG drivingcopilot’s displays automatically revert to the followingdisplay source selections:

(C) Both

~ ~

IRS 2 MAP ModeDADC 2 FMS 2NAV 2 ID WAYPTRADALT 2RAD ALT SET - 200BRG - AUTOBARO - INFD CMD - SC

Pilot’s and Copilot’s DC Fails

the

In the event that both the pilot’s and copilot’s DC fails, theSG driving the ENGINE and CAS displays will automatically useDAU 1 channel A, DAU 2 channel B,-and FWC 1 sources. Thepilot/copilot PFD and ND sources for this condition are thesame as previously described in paragraphs 2.I.(5)(a) and (b).Dual DC failures have no effect on the FGC source other thanthe source cannot be manually selected via the DC.

22-14-00Page 298.34Aug 15/91


PILOT’S DC

/m]

COPILOT’S DC

-j

EIFlCOMPEICAS

COPJ;OT

II

Engine Display Failure Reversionary ModeFigure 252

22-14-00Page 298.35

Aug 15/91Use or disclosureofmformatlonon Ihlspage !ssubjecttotherestrictionson thetitlepage ofthisdocument

Honeywell !!!!!!!!)!.c’

PILOT’S DC

zmlm~

EllzlCOMPEICAS •1COP;&OT 1COPILOT

PFD

CAS Failure Reversionary ModeFigure 253

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Aug 15/91Use or disclosure Of information On this page is subject to the restrictions on the title page of this document.

2. J. Traffic Alert and Collision Avoidance System (TCAS) Displays

(1) TCAS General

The TCAS system determines the range, altitude and bearing of otheraircraft equipped with mode S or ATCRBS transponders with respect tothe location of own aircraft. The system monitors the trajectory ofthese target aircraft for the purpose of determining if any of themconstitute a potential collision hazard. The system is responsiblefor estimating the separation at closest approach and determining ifa potential conflict exists. If so, the system displays aresolution advisory on the vertical speed tape of the pilots andcopilots map mode o the navigation displays. These displays arecrew selectable through the DC-884 Display Controllers. The TCASsystem page will automatically callup when the TCAS detects anytraffic advisory or resolution advisory target.

The TCAS system is composed of the TCAS computer with inputs fromradio altimeters, mode S computers, and through the mode Scomputers, the air data computer. In addition, the TCAS has twodirectional antennas for range and bearing information. The TCASoutput information is sent to the symbol generators through two highspeed 429 buses. Mode of operation is selected with the TCASadapter panel, but is dependent on selected mode S controller statusas well as internal and discrete logic, such as current radioaltitude, gear up/down, weight-on-wheels, and the ground proximitywarning system inhibit.

TCAS FAIL is displayed on the CAS in eventof failure, TCAS mode ofoperation is displayed beside the PFD vertical speed tape, and anextended test page is available on the ground on the TCAS systempage.

(2) TCAS Display Pages

The TCAS information is displayed in three primary areas: a TCASsystem page on DU 4, the map mode of the navigation display onDUS 2 and 5, and the vertical speed tape on the PFD. The CAS isused for a blue TCAS FAIL message.

(a) System Page Display (Figure 253.1) - The system page on DU 4 isthe primary traffic display and is selected by the crew througheither the DC-884 Display Controller or is called up by theTCAS in the event of either a Resolution Advisory or a trafficadvisory. Like the ND, the system page is centered on an ownaircraft symbol and provides a fixed range fore and aft of six-miles and side-to-side of five-miles. A two-mile rangering-of-dots surrounds the aircraft symbol with a dot at eachclock hour position. Targets that exceed the fixed range ofthe display are shown as off-scale targets around the edges ofthe display, and up to two lines of no bearing target

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2

information are displayed in the lowerwhen needed. All displays are limited

portion of the screento sixteen targets. The

display mode, above/normal/below, is annunciated in the upperright corner of the display.

J. (2) (b) Navigation Display (Figure 253.2) - The ND is the only rangeselectable display for TCAS. TCAS information is overlayedonto the normal map display information. Only targets withinthe selected range are displayed and off-scale targets are notused. This display does not use a two-mile range ring and nobearing targets are not displayed on this page. This displayis selected for on-side display through the pilot and copilotDC-884 Display Controllers. Symbols used are identical tothose used on the system page. Targets are subject to theA/N/B selection annunciated on the TCAS system page.

(c) Primary F1ight Display (Figure 253.3) - The PFD is used forresolution advisory guidance information, as well as displayingthe current TCAS operational mode. The vertical speed scale isexpanded to t6000 fpm. Recommended performance is presented onthe left edge of the vertical speed tape with acceptable andunacceptable vertical speeds shown in green and redrespectively. The vertical speed indicator digits are coloredto correspond with the color band for the vertical speed theyoccupy; so that if TCAS is displaying a red guarded band from-6000 fpm to -500 fpm, and a green target band from -500 fpm toO fpm, the vertical speed indicator digits are red from -6000to -500, green from -500 to O, and white above O to +6000. TheTCAS mode is displayed on the PFD, beside the vertical speedtape opposite the O fpm mark. TCAS modes are STBY, TEST, TAand TA/RA.

(3) TCAS Traffic Display Symbology (see Figure 253.1)

(a) Above/Normal/Below (ANB) - This selection is available on the. .TCAS adapter panel and is annunciated on the system page only.The ANB function determines the targets to be displayed byrelative altitude from own aircraft. The normal mode is notannunciated and displays all altitude reporting traffic from2700 ft below the aircraft to 2700 ft above. Above and belowexpand the normal display to 7000 ft relative in the chosendirection, while retaining the 2700 ft selection in the otherdirection. Non-altitude reporting targets are alwaysdisplayed.

(b Resolution Advisory (RA) Traffic Symbol - The RA traffic symbolis a solid red square. Targets not reporting altitude,(transponder not on, or not transponder equipped) cannotgenerate resolution advisories.

22-14-00Page 298.36.2


2. J (3) (c) Traffic Advisory (TA) Traffic Symbol - The TA traffic symbo”a solid amber circle. TAs may be non-altitude reportingaircraft, if non-altitude reporting they will always beselected for display.

is

(d) Proximate Traffic (PT) Symbol - The PT symbol is a solid cyan(pale blue) diamond. A PT is any altitude reporting trafficwithin t1200 ft of own aircraft altitude and within 6 miles.

(e) Other Traffic (OT) Symbol - The OT symbol is a hollow cyandiamond. OT aircraft are altitude reporting targets outsideilZOO ft of own aircraft, but within the ANB limits of either2700 or 7000 ft as selected, of own aircraft altitude.

(f) Data Tag - A data tag is composed of two items, the relativealtitude display (RAD), and the intruder vertical speed arrow(IVSA). These items are attached to each altitude reportingtarget. The RAD represents the altitude separation between thetarget’s altitude and own aircraft in hundreds of feet, roundedto the nearest hundred. Targets above have their data tagabove their symbol, targets below have their data tag belowtheir symbol A t symbol is also attached to each RAD asconfirmation. Data tags are the same color as their associatedtarget. The IVSA is an arrow appended to the RAD to indicatetarget vertical speed direction, if it exceeds 499 fpm. Thisarrow points up for climbing and down for descending traffic.

(9) off-scale Symbol - On DU 4 only, targets that are being trackedbut are not within the range of the display are shown asoff-scale targets. These are regular target symbols cut inhalf and against the edge of the display on the correct bearingto the target. A data tag is attached to each symbol.

(h) No Bearing Targets - An area below the target display on DU 4is reserved for no bearing targets. These are targets whichare altitude reporting but have no resolved bearing. This isusually a temporary condition and this area is reserved for RAand TA targets. For these reasons the header of NO BEARINGTARGETS is only displayed when active. No bearing targetinformation is displayed in the color corresponding to threatlevel of the target; RAs are written in red and TAs are writtenin amber. All information usually provided on targets ispresented including threat level, distance of separation andall data tag information.

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Apr 15/93Useor disclosure of information on this page is subject to the restrictions on the title page of this document.

2. J. (3) (i) Voice Advisories - Voice advisory is provided through theaircraft audio system and consists of advisories and commandsas needed:

●

●

●

●

●

●

●

TCAS TEST is announced at the beginning of test.

TCAS PASS indicates a successful test completion.

TCAS FAIL indicates an unsuccessful test.

TRAFFIC, TRAFFIC, TRAFFIC indicates a target hastransitioned to traffic advisory threat level.

CLIMB, CLIMB, CLIMB or DESCEND, DESCEND, DESCEND is aresolution advisory alert that accompanies PFD verticalguidance.

INCREASE CLIMB or INCREASE DESCENT are alerts to furtheraction needed to resolve conflict. These accompany revisedPFD vertical guidance.

CLEAR OF CONFLICT is announced at the resolution of threat.

Other advisories are provided for and a complete list can befound in the TCAS Pilot Manual.

(4) TCAS Test (Figure 253.4 and 253.5)

TCAS test is initiated from the test button of the ACTIVE mode Scontroller and lasts approximately 8 seconds. TCAS test consists ofthe following:

(a) The TCAS system page will appear if not already selected(Figure 253.4).

(b) TCAS TEST will be announced over voice advisory.

(c) TCAS FAIL will be displayed on CAS (Figure 253.4), and TESTwill be displayed on PFD as TCAS mode (Figure 253.4).

(d) Four targets will be generated by the TCAS computer anddisplayed on any active TCAS display (Figure 253.4).

● RA at 2 NM and level at +200 feet.

● TA at 2 NM and climbing at -300 feet.

. PT at 3.6 NM and descending at -1100 feet.

● OT at 3.6 NM and level at +2000 feet.

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Apr 15/93Useordisclosure of information onthispage issubject totherestrictions on the title page of this document.

2. J. (4) (e) A resolution advisory will be generated and displayed on thePFD as shown below (Figure 253.4).

. Green from O to +250 feet (Green target band).

. Do Not Climb >+2000 feet (Upper red guarded band).● Do Not Descend (Lower red guarded band).

(f) Test result will be announced over voice advisory, either TCAS

PASS or TCAS FAIL.

(5) TCAS Extended Test (Figure 253.6)

TCAS extended test required weight-on-wheels, transponder active andTCAS in standby. Prior to exiting the test mode, TCAS looks at thetest button to see if it is still active. If so, TCAS enters theextended reporting mode for maintenance information. Thisinformation duplicates some of the information on the front of theTCAS computer, but is current and active as long as the test buttonis held. If, for example, the transponder No. 2 circuit breaker ispulled while viewing the test page, the Mode S Bus 2 will changefrom active to inactive until the breaker is reset.

NOTE: Attempting extended test with both TCAS and transponders instandby will lockup the TCAS in test mode. Activating theNo. 1 transponder will release it.

(6) TCAS Computer Test

A test function is available from the front of the TCAS unit itself.This test is of the current status of the TCAS and associatedequipment. Failures are annunciated through indicators on the frontof the CU. Past failure data is also available by pressing the testbutton again within eight seconds. This will display any failureslogged to memory on the previous flight, a flight is defined as aWOW transition. This can be done repeatedly, stepping back oneflight each time test is pushed. The end of flight memory issignalled by flashing all indicators. Ten flight legs may beavailable.

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Apr 15/93Useor disclosure of information onthispage issubject to the restrictionsonthetitle page of thisdocument

OFF

SCALESYMBOL

TRAFFICADVISORYTRAFFICSYMBOL

2NMRANGERING

9 ‘“-p ‘N “v0000

‘:z =$

ENG START tiv

SENSOR FLT REF TRS S?SISM FUEL EXCEEDANCES

Ooom APU/BLEED WAYPTLISTsn

,15Y3 5m CHECKLIST

oa (2

DATATAG\

n 1!

w \

II I NO BEARING TARGETSII

1.1 NM +02T~ 2.0 NM -23?

+

/ ABOVE,NORMAL,OR BELOWANNUNCIATOR

~ PROXIMATETRAFFICSYMBOL

- RESOLUTIONADVISORYTRAFFIC

SYMBOL

- DATA TAG

- OWNAIRCRA~SYMBOL

- OTHERTRAFFICSYMBOL

- NO BEARINGTARGETS

AD-34393@

TCAS System Page DisplayFigure 253.1 22-14-00

Page 298.36.6Apr 15/93

Use or disclosure of informationon thispage issubject to the restrictionson the titlepage of thisdocument.

9~oJ”’~””Gzl Ooo

SENSOR FLTRIIf ‘IRs SYSTEM

0000

-J35 5

Sfr

TCASTARGETS%

TCAS Targets on theNavigation Display

Figure 253.2

Aoasal@

22-14-00Page 298.36.7


HoneyweU !!!!!!!~;;’

0.22M

.$

CRS

1. 010

O.io

.2 HDG050

0

(3

ITCASMODEANNUNCIATOR

TCAS Resolution Advisory onthe Primary Flight Display

Figure 253.3

/EXPANDEDVERTICALSPEEDSCALE

OVERTICALSPEEDGREENTARGETBAND

‘VERTICAL SPEEDREDGUARDEDBAND

Ao-s4s92@

22-14-00Page 298.36.8


Ak34S9S@

TCAS Test on the System Page DisplayFigure 253.4

22-14-00Page 298.36.9

Apr 15/93IJse or disclosure of information on this page is subject to the restrictions on the title page of this document,

7136 I LOC I I GS I E16000

Bs8180 AP 1 6000,

190 2Ll— 20

2000 5800

1o— 100

;’2❑@ 55%

23 0 5480

240REF 10==10 o

250 520020— 20

260 1500

rIzz#l so ;:000.22M

4

:;:1.

E

.\\’j’\ ‘J ‘II,, !ivdME i ~

\

P“/, 1

~. TEST ~, o

0.50$&\

o

f

.0+-

—

.2 HOG050

00 OK\ ‘Owl: :

\

/TCASMODEANNUNCIATOR

TCAS Resolution Advisory Teston the Primary Flight Display

Figure 253.5

<GREENTARG~BAND

‘LOWERREDGUARDEDBAND

Ao-34394@

22-14-00Page 298.36.10


TCAS TEST

COMPUTER UNIT PASSUPPER ANTENNA PASSLOWER ANTENNA PASSRADIO ALT BUS 1 ACTIVERADIO ALT BUS 2 ACTIVEMODE S BUS 1 ACTIVEMODE S BUS 2 ACTIVE

~

@

@,

--------Au-mq

TCAS Extended Test onSystem Page Display

Figure 253.6

22-14-00Page 298.36.11

Apr 15193


2. K. Microwave Landing System (MLS)

(1) General

The ML-850 MLS Receiver is designed for use with the C-band TimeReferenced Scanning Beam MLS as defined in 1986 by ICAO. Groundstations conforming to the previous ICAO definition (predominantlyEuropean) are not compatible with this receiver as they did notprovide for magnetic heading selection of runway centerline.

The MLS system provides 200 channels between 5031.0 and 5090.7 MHz.The signal format is time multiplexed with each function (azimuth,elevation, basic data, auxiliary data, and back azimuth) transmittedsequentially on a single frequency. Each function is identified byan encoded preamble, followed by TO and FRO scanning beam signals ormore digital data depending on the function.

Basic data from the ground is used to determine runway length forazimuth scaling, proportional coverage limits, minimum glidepath,runway heading and station identification. The receiver alsoprocesses and outputs auxiliary data which pertains to the ground

station for use by other systems such as EDS, AFCS, FMS< or RNAVequipment.

The ML-850 receiver system provides guidance to the azimuth andglidepath angles selected on the control unit or transmitted fromthe ground station. Guidance is output from the receiver in theform of digital deviation signals intended to drive conventionalcourse deviation indicator (CDI) displays. The MLS receiver scalesand biases these ARINC 429 labels to the corresponding ILS mV perdots of deviation allowing integration to the autopilot and displayon the EDS.

The receiver computes the centers of the received TO and FRO scans,calculates the aircraft position angle for each scan, and subtractsthe selected angle to derive deviations.

Approach azimuth angles may be selected from the runway centerlineout to the limits of the proportional coverage area of the groundstation. This angle is entered as the approach magnetic heading.Azimuth deviations are scaled as a function of runway lengths.

Glidepath angles may be selected from the minimum safe angle for thedesired runway heading (as transmitted from the ground station) tothe maximum allowable glidepath angle of 4 degrees for the G-IV.Glidepath deviations are scaled as a function of the selectedglidepath angle.

22-14-00Page 298.36.12


2 K. (2)

The MLS receiver transmits ARINC 429 labels for the purpose of

flight guidance, crew display, and internal monitoring. Once these

labels are transmitted by the MLS receiver, they are switched intothe SG-884 Symbol Generator by an external relay. The DC-884Display Controller energizes this relay based on flight crewselection of MLS as the active or preview nav source.

A Morse code station identifier is decoded by the receiver andoutput as an audio signal, a discrete signal and digitally on bothbuses.

The MLS system in the G-IV is configured as a dual receiver and dualcontrol head system. Each MLS receiver tunes a DME in a sharingscheme with the other equipment on the aircraft. MLS only controls

DME when NAV is not selected. As MLS uses the DME it will be normalto see “F---” in the NAV control head window. MLS No. 1 tunes No. 1

DME receiver, MLS No. 2 tunes No. 2 DME receiver. This DME

information is then sent to the symbol generators with the paired

MLS receiver as selected by the display controller.

MLS Displays

MLS is selected on each DC-884 Display Controller with both onsideand offside sources available.

MLS 1 FAIL or MLS 2 FAIL is displayed in the CAS as appropriate inthe G-IV dual installation. MLS FAIL is also supported in theFC-880 Fault Warning Computer, but is for single installations.

(a) Primary Flight Display (Figure 253.7)

Annunciation for lateral and vertical modes with MLS are AZ andGP respectively, and are displayed at the top of the PFD aswith other nav sources. AZ and GP are displayed in white whenarmed, boxed in green for five seconds when captured, and greenwith no box after five seconds.

The vertical deviation indicator beside the attitude sphere isgreen and is labelled P for path in glidepath. Lateraldeviation on the compass portion of the display is conventionaland also green.

The NAV source annunciator displays the active source as eitherMLS 1 or MLS 2 as appropriate. This is displayed in whiteunless the pilot and copilot select the same NAV source whenthey are in amber. Directly below it is the collocated DMEdisplay. The distance display will be dashed if DME isunavailable.

As the G-IV MLS is configured for front azimuthonly, there is no need for a TO/FROM indicator.

approaches

22-14-00Page 298.36.13

Apr 15/93Use or disclosure of informationon this page issubject to the restrictions on the title page ofthis document,

COMPARISONMONITORANNUNCIATOR ~

LATERAL VERTICALMODE MODEANNUNCIATOR

\ANNUNCIATOR

/\ //

136 I Yzl I GPJI Lii6000

E z ; k “

180 AP 1 6000

190 20— 20

000 5800

21lo— 10 0

22 A. ❑

MLS55T

23 0 541%

240 REF 10:10 0

250 52002~20

260 1500

- 5000

0.22M 1770M 30.021N

4

CRS1. 010 F\\\t;lm~~tiM46

\\

\

C“/,1

:$ 0° ~$TA’R !—

+

\

—

.2 gig ~“ ~\ OADFI ~

/ \COURSE COURSEDEVIATIONBAR POINTER

- VERTICALDEVIATIONINDICATOR

— NAVSOURCEANNUNCIATOR

= DMEDISPIAY

AX4.U9S@

MLS Displays on the Primary Flight DisplayFigure 253.7

22-14-00Page 298.36.14


2. K (2) (b) Navigation Display

MLS has available both active and preview modes, andselected on the display controller NAV page.

~ Active Mode (Figure 253.8)

s

When displayed as the active nav source MLS will be familiarfrom ILS usage. AZ will be displayed in white at the topleft of the ND, with the numeric display below it in green.Unlike ILS or VOR which use the course knob to set desiredcourse, the desired azimuth is set on the MLS control heador is uplinked from the ground station. The course pointerand deviation bar will also be in green. As with the PFDannunciation, the active nav source displayed in the topriqht hand corner is white unless the pilot and copilotseiect the same source then it will be”amber. “

The vertical deviation indicator functions as with ~is green and is labeled P for path in glidepath.

DME is in the top right corner and is displayed withThe distance readout will be dashed if DME is unavai”Due to DME sharinq scheme the DME can be unavailable

LS but

MLS.able.to the

MLS . As the G-IV-is configured for front azimuth approachesonly there is no need for a TO/FROM indicator.

Preview Mode (Figure 253.9)

With MLS, preview mode works very similiary to ILS preview

with appropriate annunciation changes. MLS azimuth, DME,

and source are displayed in a box in the lower portion ofthe display. Deviation information will be displayed in thepreview ghost display in addition to the active source.Again DME will be unavailable if NAV is the active source.

22-14-00Page 298.36.15


COURSE NAVAZIMUTH COURSE DEVIATION SOURCEDISPLAY POINTER BAR ANNUNCIATOR

\ I!

/42 Y1050 MAGI MLS1020 17.2 DM

HDG025

WxVARGCR

I 0 ADF1

SAT–52TAS300

GSPD325

.

,

~DMEDISPU4Y

/VERTICALDEVIATIONINDICATOR

Ao-s4397@

MLS Active Mode Displays onthe Navigation Display

Figure 253.8

22-14-00Page 298.36.16


PREVIEWPREVIEW COURSECOURSE DEVIATIONPOINTER BAR

\ 1

DTRK020

HDG025

M .t-f050 MAGI FMS165.0 DME

IRS2fl

\\l J~q,

\

//\’ 3

+.+ P’5,1;

\

\

//o ,/

,; +=>2 -----

+

-.:._---->...------+;. ..-/------

‘-’’--’-’=&=

.-~..- --m.0/0 0‘/ 1;‘/ 4f

+’

%/,\\Lz \\

h;; n\\\’I

O ADF1II

Wxr

——— ——— .—— . .PREVIEW I

SATVAR

I–52

GCR TAS

‘ 58

MLS1

17.2 DME I g:

‘4 0(——— ——— .—— 325 II,

I \,!I r \

\\

/ \ \PREVIEW PREVIEW PRE-VIEWA21MUTH DME NAVDISPLAY DISPLAY SOURCE

ANNUNCIATORAD-343ea@

MLS Preview Mode Displayson the Navigation Display

Figure 253.9

22-14-00Page 298.36. 17/298.36.18


3. DFZ-820 Dual Fliaht Guidance Svstem

A. System Performance/Operating Limits

The system performance/operating limits for the System are presented inTable 205. All limits have a tolerance of f10 percent.

Mode Control or Sensor Parameter Value

A/P A/P Engage

Basic A/P TCS Switch

Turn Knob

Pitch Wheel

ElectricPitch Trim

Heading GP-820Select Heading Knob

and HDG SELButton

VOR/VORAPP GP-820LNAV Button,CRS Knob, andNAV Receiver

Engage Limit

Roll Angle Limit

Pitch Angle Limit

Roll Angle Limit

Roll Rate Limit

Pitch Angle Limit

Pitch “g”Command Limit

Up Trim LimitDown Trim Limit

Roll Angle Limit

Roll Rate Limit

VOR/VOR APP Capture:Beam InterceptAngle (HDG SEL)

Lateral BeamSensor (LBS)Capture Point

Rol1: Up to *75 degPitch: Up to t50 deg

Up to t35 deg maxt6 deg min

20 deg up, 15 deg down

*3O deg Roll

*5.5 deg/sec


+o.2g; -0.1 g

20.52 0.5deg6.()~ ().sdeg

~za deg120 deg low bankswitched on GP-820

4.0 deg/sec2 deg/sec low bank

Up to t90 deg

Function of BeamError, BeamClosure Rate,and Course ErrorMin Trip Point,*20 mv dc

Max Trip Point,~180 mv dc

System Performance/Operating LimitsTable 205 22-14-00

Page 298.37Aug 15/91




VOR/VOR APP Roll Angle Limit(cent)

NOTES:

I

LOC or BC

I

Roll Rate Limit

Course Cut LimitDuring Capture

VOR Track:Roll Angle Limit

Roll Rate Limit

CrosswindCorrection

Over Station:Course Change

Roll Angle Limit

Roll Rate Limit

1. VOR APP mode is automatic and engagesflaps other than zero.

t24 deg VORt30 deg VOR APP

5.5 deg/sec VOR7.0 deg/sec VOR/APP

~45 deg Course Error

314 deg

4.0 deg/sec

Up to t45 degCourse Error

Up to t90 deg

~14 deg

7.0 deg/sec

when TAS s 180 knots or

2. VOR mode cannot be armed or engaged unless IRS TRUE TRACK isvalid (IRS TRUE TRACK is valid only if groundspeed is >20 kts).

GP-820 LOC Capture:LNAV Button (LOC), Beam Intercept Angle Up to t90 degBC Button (BC), (HDG SEL)CRS Knob, andNAV Receiver Capture Point Function of Beam,

Beam Closure Rate,and Course ErrorMin Trip Point,

~35 mv dcMax Trip Point,

~200°mv dc(~175 mv d~-906/907 FGC)

System Performance/Operating Limits 22-14-00Table 205 (cent) Page 298.38

Apr 15/93Useor disclosure of information on this page issubject to the restrictions on the title page of this document.

MAINTENANCEMANUALGULFSTRE4M IV


LOC or BC(cent)

APR

GA

I

GP-820APR Button,CRS Knob,and NAVReceiver

Control Switcheson Throttles(DisengageA/P)

Roll Angle Limit

Roll Rate Limit

Course Cut LimitDuring Capture

LOC Track:Roll Angle Limit

Roll Rate Limit

CrosswindCorrection

LOC Gain Programming

Glideslope Capture:Capture Point

Pitch Command Limit

Glideslope Damping

Pitch Rate Limit

GS Gain Programming

Fixed Flight DirectorPitch-Up Command;Wings Level in Roll

~30 deg

7.0 deg/sec

*45 deg Course

224 deg

5.5 deg/sec

Up to t45 degCourse Error

Function of RadioAltitude, TAS andGS Deviation

Function of Beamand Beam ClosureMin Trip Point,

20 mV dcMax Trip Point,37.5 mVdc

+10 deg, -15 deg

ErrorRate

Vertical Acceleration

Preset 0.2 g maximum

Function of RadioAltitude, TAS andVertical Speed

15.0 deg noseup(12.0 deg noseup-906/907 FGC)

System Performance/Operating LimitsTable 205 (cent) 22-14-00

Page 298.39Apr 15/93

Useor disclosure of information on this page issubject totherestrictions on the title page of this document.


ALT Hold DADC

VS Hold DADC

FLCH(IAS/MACH)

DADC

ALT DADC andPreselect ALT SEL ControlARM (ASEL)

ALT HoldEngage Range

ALT HoldEngage Error

Pitch Limit

Pitch Rate Limit

VERT Speed Engage

VERT Speed HoldEngage Error

Pitch Limit

Pitch Rate Limit

Mach Engage Range

Mach Hold Error

Pitch Limit

Pitch Rate Limit

IAS Engage Range

IAS HoldEngage Error

Pitch Limit

Pitch Rate Limit

VS Required forAutomatic ASEL ARM

O to 60,000 ft

~40 ft


Preset 0.1 g maximum

O to +6000 ft/min-8000 ft/min

*1OO ft/min


0.1 g maximum

0.4 to 0.88 Mach

~0.01 Mach


0.2 g maximum

80 to 340 KN CAS

&5 KN


0.2 g maximum

60 ft/min for3 seconds in thedirection of the ALTpreselect value


Page 298,40Aug 15/91



ALT DADC and PreselectPreselect ALT SEL Control Capture Range

CAP (ASEL)Maximum VerticalSpeed for Capture

Capture ManeuverDamping

Pitch Limit

Pitch Rate Limitat Capture

Maximum AltitudeCapture Error

VALT Hold FMS

VALT FMSPreselectARM (VASEL)

VALT FMSPreselectCAP (VASEL)

ALT Hold Engage Range

ALT Hold Engage Error

Pitch Limit

Pitch Rate Limit

VS Required forAutomatic VASEL ARM

PreselectCapture Range

Maximum VerticalSpeed for Capture

CaptureManeuver Damping

Pitch Limit

O to 60,000 ft

*20,000 ft/min

ComplementedVert Acceleration


0.05 g with VSK 10,000 ft/min

0.08 gwith VS> 10,000 ft/min

325 ft

O to 60,000 ft

*4O ft.

20 deg up,15 deg down

0.1 g maximum

60 ft/min for3 seconds in thedirection of theALT preselect value

O to 60,000 ft

t20,000 ft/min

Complemented verticalacceleration


System Performance/OperatingLimitsTable 205 (cent) 22-14-00

Page 298.41Rug 15/91

Use or dkclosure of mformatlon on tfVs page is subject to the restrictions on the title page of this document.


VALT Pitch Rate LimitPreselect At Capture;~;n~~SEL)

Maximum AltitudeCapture Error

VIAS FMS Mach Engage RangeVMACH

Mach Hold Error

Pitch Limit

Pitch Rate Limit

IAS Engage Range

IAS Hold Engage Error

Pitch Limit

VPATH FMS

Pitch Rate Limit

Altitude Range

Angle Range

Bias Range

Pitch Limit

Pitch Rate Limit

0.05g forVS < 10,000 ft/min0.08g forVS > 10,000 ft/min

?25 ft

0.4 to 0.88 Mach

&O.01 Mach


0.2 g maximum

80 to 340 KN

&5 KN


0.2 g maximum

O to 60,000 ft

O to -6 deg

f (FMS waypoint)


0.1 g maximum


Page 298.42Aug 15/91

Use or disclosure of mformatlon on thm page is subpct to the restrictions on the tttle page of thw document.

3. B. Flight Director/Autopilot Functional Description

(1) General

Paragraph 3.B.(2) discusses conditions and functions that arereferred to in the text accompanying each mode of operation inparagraphs 3.B.(4), (5), and (6). These paragraphs discuss thesignal flow through the flight guidance computer for each flight-path mode and the associated roll, pitch, or yaw flight controlaxis. Figures 259, 260, and 261 are simplified diagrams that showthe signal flow and interconnect wiring for the applicable selectedflight director mode and autopilot axis. Figures 254, 255, and 256are AP, YD, Mach Trim, and PFD-CMD select, mode select, and APengage logic diagrams that are used in conjunction with Figures 259,260, and 261 to aid in understanding the system operation.

(2) Control Functions

(a) Lateral Beam Sensor (LBS)

When flying to intercept the VOR or LOC beam, the LBS will betripped as a function of beam deviation, course error, TAS andDME. In the LOC mode, the course error is compared with thebeam deviation signal and rate of crossing the beam todetermine the LBS trip point.

When the LBS trips, the flight director commands a turn towardthe desired VOR radial or runway at the optimum point for asmooth capture of the beam. If the intercept angle to the beamcenter is very shallow, the LBS will not trip until theaircraft is near beam center. For this reason, an override onthe LBS occurs when the beam deviation reaches a specifiedminimum. The minimum beam sensor trip point for the VOR modeis ~20 mv. In the LOC mode, the minimum trip point is f35 mV.The maximum LBS trip point is t175 mV in VOR and t200 mV (*175mV -906/907 FGC) in LOC.

(b) Vertical Beam Sensor (VBS)

The VBS determines the point of glideslope capture utilizing anumber of inputs. The VBS is armed when the NAV radio is tunedto a LOC frequency, the LOC receiver is valid, and the LBS istripped. The VBS trips as a function of vertical speed, TAS,and glideslope deviation. The VBS will trip when verticaldeviation is less than 150 mV and acapture sensor issatisfied. The capture sensor combines airspeed, rate ofchange of beam deviation, and acceleration to determine theoptimum capture point. In the event the aircraft isparalleling the beam, i.e., no beam closure rate, the VBS willtrip at a vertical deviation less than 20 mV. This will resetthe previously selected pitch mode and change aircraft attitudeto smoothly capture the glideslope beam.

22-14-00Page 298.43Apr 15/93

Useor disclosure of information onthispage issubject to the restrictions onthetitle page of this document.

3. B. (2) (c)

(d)

(e)

(f)

HoneywellIn the back course (BC)selected.


mode, the VBS is locked out when BC is

When capturing from above the beam, the aircraft must bedescending at a rate that will create a suitable interceptangle.

Glideslope Gain Programming

Gain programming starts after the VBS trips. The gain isprogrammed as a function of radio altitude and vertical speed.If the radio altimeter is invalid, gain progrannningoccurs atGS capture and is controlled by a runway height estimator. Thevalue estimated is a function of GS capture, GS track, andmiddle marker. At GS capture, the height is estimated at 1500feet. At GS track and middle marker not passed, the height is250 feet. At GS track and middle marker passed, the height is100 feet. If the DADC is not valid, vertical speed is apreprogrammed fixed rate.

True Airspeed (TAS) Gain Programmer

TAS gain programming is used to program heading select error,course select error, pitch wheel command, air data commands,and glideslope deviation to achieve approximately the sameaircraft response regardless of the aircraft’s airspeed andaltitude. The TAS computation is derived from airspeed,altitude, and outside air temperature.

PFD Command Bars or Cue

When a command signal is applied to the bar or cue input, thebar will move left or right (roll) or up or down (pitch) andthe cue will move CW orCCW (roll) or up or down (pitch). Thisprovides the required visual command to allow the pilot tomaneuver the aircraft in the proper direction to reach thedesired flightpath.

If the information required to fly the desired flightpathbecomes invalid, the command bar or cue is biased from view.

VOR 0SS

The over station sensor (0SS) is used to detect the erraticradio signals encountered in the area above the VORtransmitter. When these radio signals reach a certain level ofdeviation, they no longer are useful and the 0SS eliminatesthem from the control signal.

22-14-00Page 298.44Aug 15/91


The VOR over station sensor trips whenare satisfied:

● VOR track has

. Either of the

the following conditions

occurred plus 3 seconds of elapsed time.

following conditions occurs:

- Distance to the station less than one fourth of thebarometric altitude and DME present.

- Lateral deviation is greater than 75 mV and the rate ofdeviation is greater than 8 mV per second and the DME isnot present.

3. B. (2) (g) GS CAP

The following conditions are necessary for glideslope capture:

. Glideslope mode is armed plus 3 seconds.

. The localizer mode is captured or in the track phase.

. Glideslope deviation is less than 150 mV.

. Either of the following conditions

- The VBS tripped.

- GS deviation less than 20 mV.

(h) GS Track

is satisfied:

Glideslope track occurs after the aircraft has captured theglideslope and is now tracking the beam. The track phaseprovides for tighter flying of the beam. The followingconditions are necessary for the track mode to be satisfied:

. GS capture plus 15 seconds.

. Localizer has gone into track 1 or track 2.

● GS deviation must be less than 37.5 mV.

● The vertical deviation must be changing at a rate of lessthan 10 ft~sec.

22-14-00Page 298.45Aug 15/91


Honeywell #i[!r.cE3. B. (2) (i) VOR AOSS 1 and VOR AOSS 2

When the aircraft is flying in the 0SS state, beam deviation iscontinually monitored to determine when it is again useful toinclude in the command signal. The AOSS monitors beamdeviation. When certain conditions are satisfied regardingbeam deviation, the AOSS will trip. There are two stages tothe AOSS. The first stage is AOSS 1, and AOSS 2 is the secondstage. AOSS 2 will not trip until AOSS 1 has tripped. Thesesensors ensure that when beam deviation is again included backinto the control signal it will indeed be useable information.

VORAOSS 1 will occur when the following conditions are allsatisfied:

. VOR 0SS has occurred dependent on the active lateral mode.

. A calculated period of time has elapsed since the lastto/from transition on the EHSI in order for AOSS 1 to trip.The period of time elapsed is calculated using true airspeedand altitude. The higher the altitude, the longer it takesto get through the cone of erratic radio information;therefore the longer the time period must be. Likewise, thelower the aircraft altitude, the smaller the cone of erraticradio information, and the shorter the time period must beto trip AOSS 1. The required elapsed time period is alsoaffected by the aircraft’s true airspeed. The faster theairspeed, the quicker the aircraft will be through the cone.’The slower the airspeed, the longer it will take to passthrough the cone, and a longer time period is required totrip AOSS 1.

VOR AOSS 2 will occur when the followina conditions are allsatisfied:

. VORAOSS 1 has tripped

. Beam deviation is less

● Beam rate is less than

plus 3 seconds.

than 75 mV.

25 feet per second.

Once VORAOSS 2 trips, beam deviation will again be part of thecontrol signal.

(j) VOR CAP

VOR capture will occur when the following conditions aresatisfied:

● The VOR mode has been armed plus 3 seconds of elapsed time.

● The lateral beam sensor (LBS) has tripped.

22-14-00Page 298.46Aug 15/91

Use or disclosureof intorrnation on this page is subject to the restrictions on the title page of this document.

3. B. (2) (k) VOR Track

VOR track will occur as the aircraft is established on beamcenter and the following conditions are all satisfied:

● The VOR mode has been captured or AOSS 2 has occurred.

. Thirty seconds of

● Lateral deviationthe aircraft bank

elapsed time from capture.

rate is less than 50 feet per second andangle is less than 6 degrees.

At this time course error is eliminated from the commandsignal, leaving beam deviation and lateral acceleration fromthe IRS to maintain the aircraft on beam center. There is novisual indication in the cockpit that the VOR track submode hasoccurred.

(1) LOC CAP 1 and BC CAP 1

Localizer and back course capture 1 are the initial capturephases of their respective modes. Localizer capture 1 and backcourse capture 1 will occur when the following conditions areall satisfied:

● LOC armed plus 3 seconds.

. Either of the following occurs:

- Lateral beam sensor trips.- Beam deviation less than 35 mV.

(m) LOC CAP 2 and BC CAP 2

Local izer and back course capture 2 are capture phases whichindicate the aircraft is now flying closer to the center of thebeam. The capture 2 phase will occur for each mode when thefollowing conditions are all satisfied:

. LOC CAP 1 plus 3 seconds.

. Course error less than 35 degrees.

. Beam deviation less than 175mV (165 mV -906/907 FGC).

22-14-00Page 298.47Apr 15/93

Useor disclosure of information on this page issubject totheresttictions onthetitle page of this document.

Honeywell !%!!!?~.c’3. B. (2) (n) LOC Track 1 and BC Track 1

Localizer track 1 and back course track 1 signify the aircraftbeing on beam center and the roll rate limit being decreasedfrom 7.5 deglsec during the capture phase down to 5.5 degjsecin the track submode. At the track submode occurrence, thecourse error is eliminated from the control signal, leavingbeam deviation and lateral acceleration from the IRS tomaintain the aircraft on beam center. The track 1 phase willoccur when the following conditions are all satisfied:

● LOC CAP 2 plus 30 seconds.

● Estimated beam rate less than 30 feet/second.

● Localizer beam deviation less than 20 mV.

● Aircraft bank angle less than 6 degrees.

There is no visual indication in the cockpit that the LOCtrack 1 submode has occurred.

(o) LOC Track 2 and BC Track 2

The track 2 submode will occur only after track 1 has beensatisfied. There is no visual indication to the pilot that thetrack 2 mode has been activated. Radio altitude, distance tothe transmitter, and a vertical velocity indicating theaircraft is descending are the factors involved in determiningthe track 2 condition. When these conditions reach certainlevels, track 2 is tripped so as to provide tighter controlduring the final stages of an approach.

The track 2 phase will occur when the following conditions areall satisfied:

● LOC track 1 has been tripped.

● The aircraft is descending at a vertical speed which wouldindicate a runway approach.

c Either of the following conditions has occurred:

- ~i;;~a~e to the transmitter isless than approximately..

- Radio altitude less than 1200 feet with the radioaltimeter valid.

22-14-00Page 298.48Aug 15/91


Honeywell ##@R.cE3. B. (3) Flight Director Mode Selection (See Figure 255.)

There are nine mode select pushbutton switches located on the GP-820Flight Guidance Controller as shown on sheet 1. When one of theseswitches is pushed, a ground (PB ARM) is provided at llJ1-38 to theFZ-820 to interrupt the “A” processor. Also, when a switch ispushed, the ground to the GP-820 PISO is removed and 5 volts isapplied to the parallel input of the PISO. The clock and strobe forthe PISO is received from the FZ-820. The pushbutton serial dataoutput from the PISO device is routed to the SIPO device of theFZ-820 which provides a parallel output to the “A” processor. Asignal representing the mode selected is provided on the ASCB to thesymbol generator to annunciate the mode selected on the PFD.

Figure 255, sheets 2 through 15, show the logic conditions requiredto annunciate each mode and also the conditions that will reset orclear a selected mode. For example, VOR ARM (sheet 2) isannunciated on the PFD when the LNAV switch is pushed on the GP-820and the following conditions are met:

. Tuned to VOR frequency (NAV VOR Select)

. Not tuned to localizer (~)

. Lateral beam sensor has not tripped (LBS TRIP)

. All inputs to the OR gate to clear (CLR) or reset (R) theflip-flop are not present.

If VOR is annunciated in white indicating the mode is armed, themode will be reset if any of the following inputs are present to the“OR” gate to reset (R) the flip-flop:

;tR CAP 1PFD CMD SELNAV SOURCE CHANGETKOD ● AP ENG

Also, the mode will be cleared if any of the following inputs arepresent the OR gate to clear (CLR) the flip-flop:

. Flight guidance computer valid not present (FGC VALID)

. Inertial reference system miscompare and inertial referencesystem No. 3 valid not present (IRS MISCOMPARE● IRS 3 VALID)

. Selected di ital air data computer valid not present(SDADC VALI;)

. Tuned to a localizer frequency (TTL)

● CPL valid not present for two seconds (CPL VALID ● 2 SEC)

22-14-00Page 298.49/298.50

Aug 15/91Use or disclosure of information on this page is subjecf 10 the restrictions on the title page of this document,

10

0,

0

%

B

)6

N

A

x

Is

a

x

B1

10

u

I IIJ1

-——. ——— ——— —.— .—— ——— —.— .—— _

1FZ-S20 FLIGHT GUIDANCE COMPUTER (PILOT’S)

1

2

3

4

5

6

M

rT

71

66

78

m

rs

9

10

15

16

w

46

49

iIIIIIIIIIIIIIIIIIIIIIIIIII

I

IIL

l-+(H)

(c)

SERIALDATA IN

{H)

(cl

CLOCK(Llm&r4

I

SERIAL lN-PARALLEL OUTOEVICE (SIPO)

\ANN

~ LATCHEWDRIvERS

A/P ON

PFO-CMO

m

<

(H)

(c)}

P

* LAMPTEST I

} *

LAMP

!

7 9

LAMP TESTTEST 1

ANN VLOORIVE I

ENO AROUNO

( I

* CLOCK

- STR08EPARALLEL

*INPUTS

( : - -, ,

WRALLEL lN-SERIAL OUT * PFD-CMO

~ cL~R DEVICE (PISO) LEFt/RIGliT

JLD

-: 1-LOGIC 1

SERIAL +tl -

WRALLELOATA OUT

5V INPUTS

}

STROBE PARALLEL

STROBE OUTPUTS

TPILOT’S

I r-

LAMPCOPILOT S TEST

AP AP

82LOCK

c-

*

INOTES1. THE GP-620 PILOT’S SIDE (CH ! ) IS

I SHOWN, THE COPILOT’S SIDE (CH2) ISSHOWN ON SHEET 2.

, 2.

13

:?

LEVELSHIFTERS(NOTE 2)T

SERIAL lN-PARALLEL OUT A

OEVICE(SIPO)

THE INPUT IS LEVEL SHIFTED TOPROVIDE A 5 VDC PARALLEL INPUTTO THE PISO

ALL CIRCLED LETTERS DENOTE TIEPOINTS ON StiEET 2,

..w

PROCESSOR wow -9

28 VOC ANNUNCIATOR VALID9

11

r

6

K3

Is

{

(H)

(c)

—L %“%-’ J5V —PBARMMO 1

●I

I8 AP

I ‘-

I -

E!!K- DISENGAGE

HORN

1’ - &ENGAGEANNo-

apII

(III

@

@

aPFD-CMD

+

@

-@

1 I120 Voc

YD OISENG -

MACH TRIMDISENG -L 1 I

(

3ASCBOUTPUTTO SYMBOL wGENERATORS

O-5vDC *

6PUSHBUTTONSERIALOATA

=/EDGE

LIGHTING

I-(H -—— ——— — ——— ——— ——— ——— ——— —.— _

—— ——— ——— —.

AP, YD, MACH TRIM, and PFD-CMDSelect Diagram

Figure 254 (Sheet 1)22-14-00Page 298.51/298.52

Aug 15/91Use or disclosure of mtormatlon on this page is subject to the reslricllons on the title page of this document.


r—————————.

PZ420 PLIGHT GUIDANCE COMPUTER (COPILOT’S)-——— ———— ———— ———— —,— __ __>_— __GP-82UPIJGHT GUIDANCECONTROLLER (COPILors CHANNEL) 1!B 1lJ:

1

2

3

4

5

6

Ed

n

?1

X

?[

7$

H

9

la

15

16

iI

(k

(c

(H

(c

}

SERIAL DATASERIALOATA IN

}

CLOCKSERIAL IN. I

PARALLEL OUTDEVICE (Sllw)

I

PARALLEL

STROBE OUTPUTS ANN ILATCHESDRIVERS

I

CLOCK I

c=

t

wTEST 2-

IAMP TEST

ANN VLLI- ORIVE 2 I

END AROUNO

1b :)0 IF

I

++ CLOCK

w STROBE

e -o:$P( z)@

IPARALLEL IN. I

-SERIAL OUT

* C*R DEVICE (PISO)

LEVEL 4 V-O

m

*:: : 1

‘b

11J2

SHIFTERS(NOTE 2) SERIAL (H) 7

DATA WT

5VPARALLEL (c) 8INPUTS

5V

I

I

*I

IIIIII

}* LAMP

TEST2

‘“”7

4 (h

(c }

STROBE

CLOCK(u~sEFl P-LAMP

TEST

—J ●

NOTES1, THE GP.620 COPILOT’S SIOE (CH2) IS

SHOWN THE PILOTS SIDE[CH1) IS SHOWN ON SHEET 1

2. tHE iNPUT IS LEVEL SHIFTEO TOPROVIOE A S VOC PARALLEL INPUT TOTHE PISO“A”

PROCESSOR 3. ALE LE7TERS CIRCLED DENOTETIE POINTS ON SHEET 1,

wow --

26 VOC ANNUNICATOR VAIJOm

~

SERIAL lN-PARALLEL 04JT

OEVICE(SIP(3)

PSAJ3MN0.2

--03

+@

28 Voc f-l.ANN PWR —NO. 2

6

28vOCYD OISENG -

MACH TRIMOISENG -

>

ASCB OUTPUTTO SYMBOLGENERATORS

.—— — ——— — ——— ——— ——— ——— ——— ——— -1J(’ PUSH BUITON SERIAL OATA

PFD-CMD

2)

I——— ——— ——— _.

AP, YD, MACH TRIM, andSelect Diagram

Figure 254 (Sheet22-14-00Page 298.53/298.54

Aug 15/91Use or dwclosure of Intormatlon On Ihm page is subject 10 the restrictions on the Mle page 01 thm document.

Honeywell !!!?!”fr.c’

r—————————.FZ-820 FLIGHT GUIDANCE COMPUTER (PILOT’S).—— — ———— ———— ———— ———— —.—.. — ,——,

1llJ! 3P-s2O FUGHT GUIDANCE CONTROLLER (PILOTS CHANNEL 1)

1

2

3

4

5

6

64

w

72

74

2s

9

1(

1!

1(

Y

4

4

L

1+(H)

{c]

SERIALDATA IN I

IIIIIIIIIII

IIIIIIIIIIIIIIIIIIIIIIIIIIIIIIII

l/2@ANKSWI

ON@MODE SWI

ANNLATCHES/DRIvERS

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(c,

CLOCK(u::gFr

t

SERIALIN.PAFIALLELOUT

DEvICE(SIPO)

L

CLOCKSERIAL lN-

PARALLEL OUTDEVICE (SIPO) NOTES:

Y’=1

2

3

4

5

6

THE NINE MODE SELECTSWITCHES ARE IABELED,

HOG SEL ALT HLD FLCHAPR Sc VNAVBANK Vs LNAV

EACH SWITCH GOES TO ASEPARATE INPUT ON THE PISODEVICE. LOGIC FOR EACH MODESELECTEO IS SHOWN ON THEFOLLOWING SHEET5

STROBE

c-

LAMPTEST

PARALLELSTROBE OUTPUTS

* LAMPTEST 1

CLOCK

CLEAR 7 “’lLJ+=STANN VLCI fEND A170UND I

+ CLOCK

I ) - STROBE

MRALLEL lN-SERIAL OUT

DEVICE (PISO)ANN VLO A

“ )

FORALL MOOES, A SECONOACTIVATION OF THE SWITCH WILLRESET THE MODE.

THE INPUT IS LEVEL SHIFTED TOPROVIDE A 5 VDC PARALLEL INPUTTO THE PISD

THE GPw21J PILOTS SIOE (CHI ) ISSHOWN. THE COPILOTS SIOE (CH2) ISSHOWN ON SHEET 11

..A

PROCESSOR?8VOC ANNuNCIATOR VALIO

ML CIRCLEO LETTERS OENOTE TIEPOINTS ON SHEET 1 I

llJ1

{~

(H) 7

ICI 6

PILOT-STOGASW

LEvELSWTERS(NOTE 4)l——=—o

* PILOTSTCS SW

m

r“

●

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IIII

P9 ARM NO 1—‘1 SINGLE-POLE00LJSLE-THRDW

SWITCHES(NOTE1)

—

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I

6I

KAS(X OUTPUT

=3ggggJ0

PFO

2s VDCANN PWR —n~NO. 1

1-

~ EDGE

~ LIGHTING

PUSH BuTTON SERIAL OATA 6

AD-12462 @1- ——F1ight Director

Mode Select DiagramFigure 255 (Sheet 1) 22-14-00

Page 298.55/298.56Aug 15/91

Use or disclosure of mlormatlon on this pege E subject 10 the reslrtcllons on the Illle page of thm document.

HoneywellMAINTENANCEMANUALGULFSTREAMIV

r——————————

FZ-820 FLIGHTGUILIAHCECOMPUTER(COPILOT’S)

IIIIIIIIIIIIIIIIIIIIIIIIIIIIIIIIL .——

“A”PROCESSOR

v

——— ——— —.

1

SERIAL DATA

CLOCK

STROSE

c“LAMPTEST

2aVDCANNUNCIATORVALID●

COPILOTSTO(M SW

* COPILOTSTCS SW

r’”●

IPS ARM NO. 2

/$

1 “PUSH BuTTON SERIAL DATA

28 VOCANN PWR —n ●

No. 2 r*

.——. ———— ———— ———— ——. — ———— ——W-820 FLIGHT GUIDANCECONTROLLER(COPILOT%CHANNEL2) 1P-+-

{

LEVELSHIFTERS(NOTE 1}

SERIALDATA IN

SERIALlN-PARALLELCiJTOEVICEF31PO)

PARALLELSTR09E OUTPUTS

CLOCK

cm

II

\II

+— > - ANN ILATCHESl

- LAMPORIVERS

TEST 2

I

( P # I

+

wI

NOTESLAMP TEST

TEST2+ 1 THE INPUTIS LEVELSHIFTEDTOANNVLD

+ ORIVE2 PROVIDEA5 VDCPARALLELINPUT

I TO THE PISOEND AROUND 2 THE GP-820 COPILOTS SIOE (CH2) IS

SHOWN THE COPILOTS SIDE (CH1 )I IS SHOWN ON SHEET I

3. ALL CIRCLED LETTERS DENOTE TIE

-9 CLOCK

I

POINTS ON SHEET 1.

- STROBE

+

PARALLEL IN- 1

* CLEARSERIAL OUT

OEVICE (PISO)ANN VLD

.; 1‘b

11J2

SERIAL(H) 7

PARALLEL DATA OUTINPUTS (C) 8

5V

I

(5o IIIIIIII

I 1 IPULSE WIDTHCONVERTER k I

T A

IIIII

JAo12462~

F1 ight DirectorMode Select Diagram

Figure 255 (Sheet 1.1)22-14-00

Page 298.57/298.58Aug 15/91

Use or dwclosure of Intormallon on this page IS subject 10 the restrictions on Ihe htle page of thm document.

HDG SELSWITCHINPUT(REFSH1)

10J2E

rlJ17

7 a Ioi

8 + 10’

3s - 1of

2

llJ17

7 * la

LNAV SWITCH ~ 4 10INPUT(REF SHl)

Sa -9 la

2

L

r—.—— .—— ————— ———.FZ-820FLIGHT GUIDANCE COMPUTER 1

I

h CLR

ANY OTHER LATERAL MODE CAP

GA

PFD CMDSELECT

TKOD ● AP ENG H

FGC =

IRS MISCOMPARE ● lRS3~D

SDADC G

SG VALID FOR COUPLED PFD *

NAV SELECT_ t 1

d-r’VOR CAP

PFD CMD SEL

NAV SOURCE CHANGE

ITKOO ● AP ENG

+

I FGC~

IRS MISCOMPARE ● IRS 3~

I SDADC=

TTL

SG VALID FOR COUPLED PFD +

HDG●

b-VOR

Q ARM

III

III

VOR ARM ● 3 SEC

LBS TRIPSs

ovOR CAP

m

GAR

HDG SEL

LNAV SELCLR

vOR TRKL

PFO CMD SEL

NAV SOURCE CHANGE

ITKOD ● AP ENG j I

.

3J,

3Jj

,

IIIIIIIIIIII

I - 1%’B’’0J2B

rIIIIII

I

I FGC ~D

IRS MISCOMPARE ● IRS 3 ~

I SDAOCVALID

SEL LAT GUIOANCE VALID .5 SEC

I CPLVALID.2SEC=-wI TTL

I.93 _ FOR COUPLED PFD

4NOTE:vOR APPROACH MODE SELECTED AUTOMATICALLY WHEN i

IIIII

LTAS = 1S0 KNOTS OR FLAPS >0 DEGREES.

——— —.— ——— ——— ——— — -1

LATERAL/ROLL MODES AD.13662@R5

F1 ight DirectorMode Select DiaqramFigure 255 (She& 2) 22-14-00

Page 298.59Apr 15/93


-——— ———— ———— ———— ———— —-I FZ-S20FLIGNT GUIDANCE COMPUTER

IIIIIIIIIIIIIIIIIIIIIIIIIIIIIII

30sEcfi

vOR CAP+VOR AOSS 2

LATERAL OEVIATION RATE ~< 50/SEC ● BANK ANGLE< 6° I

COURSE CHANGE > 3 DEG ● =

GA vOR TRACK

HDG SEL

LNAV SEL

0ss

PFD CMD SEL CLR

NAV SOURCE CHANGE

TKOD ● AP ENG

SG VALID FOR COUPLED PFD

FGC ~

IRS MISCOMPARE ● IRS 3 VALID-3

SDADC VALID

SEL LAT GUIDANCE VALID ● 5 SEC

CPL=. 2 SEC

TTL

‘-T

‘OR‘RACK“ ‘SEC--l--& I

~~~;~?i!!-vs a ‘oRoss>75 mV . RATE

>8 mV/SEC.

w

R

DME VALIDGA

3

CLR

+ DME HOLDHDG SEL

LNAV SEL

vOR AOSS 1

PfD CMD SEL

NAV SOURCE CHANGE

TKOD ● AP SNG Yld

%Jr+R

GA CLR

HDG SEL

LNAV SEL

vOR AOSS 2

PFD CMD SEL

NAV SOURCE CHANGE

TKOD ● AP ENG 3

‘ORAO:M:=-tFGA

Y

h

CLR

HDG SEL

LNAV SEL

VOR TRACK

PFD CMD SEL

NAV SOURCE CHANGE

TKOD ● AP ENG

vOR AOSS 2D

L ——— ——— —————— ———LATERAL/ROLL MODES

F1iqht DirectorMode select DiagramFigure 255 (Sheet 3)

3J;

.—— — -1AO.13e63 @) -R3

22-14-00

I

IIIIIIIIIIII

d10J 1W1OJ2B

1ASCB OUTPUT

“A2 TO SYMBOL

PROCESSORGENERATOR TO

4 DISPLAY SELECTEDMODE ON PFD

l-’5

IIIIIIIIIIIIII

Page 298.60Aug 15/91

Use or disclosure of information on this page is subject to the restrictions on the title page of Ihts document.

LNAV OR APR SWITCHINPUT(REFSH1) !

llJ1

7 10(

8 101

38 106

I

III

IIIIIIIIIIIII

.—— ——— ——— ——— .—— ——.FZ-820FLIGNT GUIDANCE COMPUTER

NAV1 OR NAV2

w

LOC CAP1

BC SEL

GA

PFD CMIY SEL

t.lAV SOURCE CHANGE

TKOD ● AP ENG 6

SG VALID FOR COUPLED PFD~

FGC VALIDIRS MISCOMPARE*lRS3~ _

SDAOC VALIO

CPL VALIO ● 2 SEC

--+ K

LOC ARM ● 3 SEC

LOC CAP2

HDG SEL

LNAV SEL

GA

PFD CMD SEL L

ro LOC ARM

-tR

c

— 1rso LOC CAP 1

R

CLR

LOC CAP 1 ● 3 SEC

COURSE ERROR <35”

v

s

DEV <175 mvLOC CAP 2

0(NOTE)

LOC TRACK1

3

R

HDG SEL CLR

LNAV SEL

‘CS5-J L t

LATERAL/ROLL MODES

PFD CMD SEL

NAV SOURCE CHANGE

TKOD ● AP ENG T

-. .——

F1ight DirectorMode Select DiagramFigure 255 (Sheet 4)

IIIIIIII

III

loJIB/loJ2B-1

bASCB OUTPUT2 TO SYMBOL

GENERATOR TO4 DISPLAY SELECTED

MODE ON PFD

NOTE: DEV f 165 mvFOR -936/-907FGC.

AD-13662 ~-R5

22-14-00Page 298.61Apr 15/93


——. — ———— .—— — ——. — .—

I

I

IIII wR

LOC TRACK 2CLR

HDG SEL

LNAV SEL

BC SEL

GA

PFD CMD SEL

NAV SOURCE CHANGE

TKOD . AP ENG

l—–SG VALID FOR COUPLED PFD

FGC vALID

I IRS MISCOMPARE ● IRS 3 VALID

SDADCVALID

l– CPL VALID . 2SEC

“wF

I “-4SEL LAT GuIDANCE =.5 SEC I

OC TRACK 1-

(NOTE)

LOC TRACK 1

vERT VEL <-2° (TAS/57.3) LOC TRACK 2

r-ii(NOTE)

R

ClE?.-lHDG SEL ~ II 4

.NAV SOURCE CHANGE

10J2B TKOD . AP ENG 3r

IJ1 NAV. LOC SEL

s7 + 100 e SERIAL IN

BC ARMo

PARALLEL OUT T8 + 101 + DEVICE

~ R

TTL CLR

3s + 106

2 BC CAP1 4 !

BC SWITCHINPUT(REF SH1)

YAPP SEL

LNAV SEL

GA

I

PFD CMD SEL

NAV SOURCE CHANGE

TKOD ● AP ENG

I

P

I FGC VALID

IRS MISCOMPARE* lRS3~

I SDADC=

CPL = ● 2 SEC

1 K‘w

F1iqht Director

IIIIIIIIII

a,,w loJIB/loJ2B

PROCESSOR

1

Ascs ouTPuT,,A,, 2 TO SYMBOL

PROCESSOR GENERATOR TO4 DISFIAY SELECTED

MODE ON PFD5

IIIIIIIII1IIIIII

L—————— ——————— ——— -1LATERAUROLL MODES

Mode ~elect DiagramFigure 255 (Sheet 5)

NOTE:IF BOTH DADCS ARE INVALID,THE FGC WILL USE DEFAULTAIR DATA VALUES WHEN INAPPROACH TRACK MODE. THEDEFAULT VALUES ARE:

DYNAMICPRESSURE 75 La/F?

TAS 160 KNOTS

AD-136S2 @ -R3

22-14-00Page 298.62Aug 15/91

use or disclosure of information on this page is subject to the restrictions on the title page of this document.

I

I

Honeywell

r

—.— .—— ——— ———

FZ-S20 FLIGNT GUIOANCE COMPUTER

IIIIIIIIIIIIIIIIIIIIIIIIII

III

LNAV SEL q I

NAV SOURCE CHANGE

TKOD ● AP ENG1

SG VALID FOR COUPLED PFD

FGC =0

IRS MISCOMPARE* IRS3VALI0-3

=cmb--CPL VALID. 2 SEC

SEL LAT GUIOANCE VALID ● 5 SEC

m <

EC CAP 1 ● 3 SEC

COURSE ERROR <35°DEV <175 mv(NOTE) @

BC TRACK 1

APP SEL

HDG SEL

LNAV SEL

GA

PFD CMD SEL

NAV SOURCE CHANGE

TKOD ● AP ENGP

BC CAP 2 ● 30 SEC

BEAM RATE<30 FT/SEC

LO. DEV <20 mv

+BANK ANGLE <6°

EC TRACK 2

APP SEL

HDG SEL

LNAV SEL

GA

PFD CMD SEL

NAV SOURCE CHANGE

TKOD ● AP ENGP

‘cTRAcK’kVERT VEL <-2° (TAS/57.3)

–=m-LONG DIST < 26K’ ● RAVAL

RAO ALT <1200 * RAVAL

APP SEL

LNAV SEL

HDG SEL

GA

PFD CMD SEL

NAV SOURCE CHANGE

TKOD ● AP ENG +

I


——— ——— ——1

BC CAP 1●

rs BC CAP 2Q

R

CLR

Fs

Q

R

CLR

EC TRACK 1●

I

rs BC TRACK 20

R

CLR

IIII

IIIII

w10JIB/10J2B

1ASCO OUTPUT

,.A.. 2 TO SYMBOL


4 DISPLAY SELECTEDMODE ON PFD

5

1-————————————————————J

rIIIII

NOTE: DEV. 165 mv

FOR -906/-s07FGC.

I

IIIIIIII

LATERAUROLL MODES F1ight Director AD-1 3342 @ -R3

Mode Select DiagramFigure 255 (Sheet 6) 22-14-00

Page 298.63Apr 15/93

USe or disclosure of information on thispage issubject to the restrictionson the titlepage of thisdocument.

LNAV SWITCIINPuT(REF SH1)

e———— ————— ————— ———w

I FZ-820FLIGHTGUIDANCE COMPUTER

I

I FMS VALID—1OJ2B

lJ1 SQFMS ARM

*

7 + 100 * SERIAL IN

BPARALLEL T

+ 101 - OUT DEVICE

- RCLR

FMS SELECTED— )

J-’-IG=JG=J

I MOVING AWAYFROM DES TRK● WITHIN

I2.5 MILES OFDES TRK

T

l;:~~D~::.

I

DIFFERENCE BETWEENACTUAL A C TRK ●

I

DES AC TRK)

IIIIIIIIIII

LNAVARMJ

LNAV SEL

HDG SEL

APP CAP

BC CAP L

-J

rSQ FMS CAP

R

CLR

GA

PFD CMD SEL

TKOD ● AP ENG Y

5J3 a FORCOUPLEO PFD

FMS VALID

-u

FGC ~DIRS MISCOMPARE ●

lRS3~ _SDADCVALID

CPL VALID. 2 SEC 4

ISELLAT GUIDANCE I~.5SEC

iIIIII

=%

“A”PROCESSOR I10J1B/l OJ2B

1ASCB OUTPUT

2 TO SYMBOL

PRO&SOR GENERATOR TO4 DISPLAY SELECTED5 MODE ON PFD

IIIIIIIIIIIIIIIII—.—— ———— ———— ———— A

Flight DirectorMode Select DiagramFigure 255 (Sheet 7)

[email protected]

22-14-00Page 298.64Aug 15/91


llJ1

B

7FLCH SWITCHINPUT 8(REFSH1)

se

llJ1

B

7

VS SWITCHINPUT 8(REF SH1)

3s

llJ1

E

7

ALT HLDSWITCH 8INPUT(REFSH1)

3s

r——. —————— ——— ——— _.

FZ-820 FLIGNl GUIDANCE COMPUTER

I

I =LF’DMODE’EL--LJANY vERTICAL F/D MODE SEL

;G VALID FOR COUPLED PFDA

: - -FPFD CMD SEL

GA

ANY OTHER VERTICALF/D MODE ON +CAP

P

?

SDADC VALIDIRS M~MPARE ●

IRS 3 VALIDFGC ~

;G VALID FOR COUPLED PFD

PARALLEL OUT

- ‘%-JANY OTHER VERT FD

MODE ON+ CAP +

ALT SEL CAP

h

ALTERROR<25’

ALT RATE<5 FT/SEC

PW MOTION

ALT SEL KNOB MOTION

-\

I GA

I PWMOTION

ANY OTHER VERT P

IF/D MODE ON +CAP

!Ts Vs

oT

RCLR

il--S ALT

0T

R

CLR

A

1II

IIIIIII

- I-Y’”’z’

w1ASCBOUTFUT

“A 2 TO SYMBOL

PROCESSOR GENERATOR TO4 DISPLAY SELECTED

MODE ON PFD5

IIIIIIIIIIIII

1- ———— ———— ———— ——(——, JVERTICAL/PITCH MODES AD-t3e62 @ -R3


22-14-00Page 298.65Aug 15/91

Use or disclosure of mformatlon on this page is subject to the restrictions on the title page of this document,

IIIII!IIIIIIIIIIIIIIIIIIIII

r FZ4ZDPLIGNT GUIDANCE COMPUTER——. — .—— — —.—— —.. — 1

1ALT SEL cAP 1

DECREASINGALTITUDE ERROR

CMD ALT RATE > VS

vs > 1,3 FT/SEC

3 SECS

AP ENG + FD STBY d

GS CAP + TRKd

— Au=lAOM--“-- c.. . ..!

‘LT=sr’1:“tl ~.r-f

RCLR

$-pALT SEL CAP

ALT

GS CA

VNAV

GA

PFO CMD SEL

GS TRACK

1ALT SEL ARM

Q ,AL, SEL CAPDETECTOR TRIPS

ALT sET KNOB - RIN MOTION CLR

PW IN MOTION

ANY OTHER VERTMODE SEL + CAP #)

PFD CMD SEL

GA

*

FGC ~IRS MISCOMPARE ●

IRS3WDSOADC V~

SG = FORCOWLED PFD

.us i

1III1WI BH OJZB

1

ASEL CAP B 2

- us 4

15

d

hIIIII1II

b———————— ———————— —A

VERTICAL/PITCH MODES


AD.13es2 @ -w

22-14-00Page 298.66

Aug 15/91Use or disclosure of information on this page is subject to the restrictions on the Mle page of this document.

APR SWITCHlNI%tT(REF SH1 )

MAINTENANCE

Honeywell H%!#&M.

f--- “-PZ-62aFUGHT GUIDANCECOMPUTERlaJ2B

——— ——— ——— ——— ——— ——. —1

17q“:‘=~lJ1 NAV ● LOC SEL

s7 100 SERIAL IN a

PARALLEL OUT T

8 101 DEVICER

lTL..A CLR

3s 1% PROCESSORGA ~

-1IIIIIIIIIIIIIIIIIIIIIII1

GS CAP

PFD CMD S=P

vFGC ~

IRS MLSCOMPARE* lRS3~

SDAOC m

SG m FOR COUPLEO pm

GSARM ● 3 SEC—

LOC CAP + TRK —

VES TRIPS

GSDEV<20mv

~ s

GAo

GS TRACK

ANY OTHER VSRT MODE CAP R

PFOCMOSEL CLR

GS WPm

S0 ~ FOR COUPLEOPFDSOURCEVERTGUID- ● 5 SEC

LOCCAP+ TRKFGC~

IRSMISCOMPARE● IRS3VALIDSDADCW

-w

W--l--YvERT DEV RATE <<10 m/SEC

RGA

CLRPPDCMOSEL

MY OTHER VERT MODE ON +CAP1 I 1

‘EMOTEGAsw’TcH-nl-===ANY OTHER VERT MODE ON +

AJP 0?4

PFD CMO S5

-D

R

TCS CLR

SG VALID FOR COUPLED PFOFGC ~

IRS MISCOMPARE ● IRS3 ~

soArx=i15

VERTfCAIJPITCH MODES


IIIIIIIIIIII

+

‘A”PROCESSOR

10J1B/10J2B

1ASCB OUTPln

‘A-2 TO .WMSOL

PROCESSORGEWRATOR TO

4 DISPIAY SSLECTEOMODE ON PFO

5

IIIIIIIIIIII

——— — J0126s2 @j .R3

22-14-00Page 298.67Aug 15/91


I

7 ———— ———— ———— ——

I Fz.8~ FLIGHT GUIDANCE ~MpuTER

IllJ1

PW MOTION10J2B

T7 ~ 100 +

VNAVSERIAL IN

PARALLEL OUTSWITCH

INPUT 8+ 101

(REF SH 1)

w

s’10’ -

.

IVNAV Pushbutton

PFO CMD SEL.

IGS+GP CAP

GA

I

VALT

VPTH

I

VFLC

VASEL

I

FLCH

Vs

I

ALT

SFMS SWITCH OVER

I

SDADC SWITCH OVERY

IIIIII

sa

T

R

CLR

FGC VALID

SDADC V=‘3

SG VALI~+COUPLED PFD

VNAV

ARM

+

IIIII

‘A” PROCESSOR

t1

ASCB OUTPUT

“A. 2 TO SYMBOL

PROCESSOR GENERATOR TO4 DISPLAY SELECTED

5MODE ON PFD

IIIIIIIII

L——————————————— J Awl 3632@ -R1@


22-14-00Page 298.68Apr 15/93


FLCH

SWITCH

INPUT

(REF SH 1)

F ———— ———— ———— ——q

I %?-S20FLIGHT GUIDANCE COMPUTER

I f (ACTIVE vERTICAL

I

FLIGHT PLAN)

llJ1 10J2BVNAV AFIM

>7 SERIAL IN

PARALLEL OUT

8 * 101 - DEVICE

I PFD CMD SEL J

J

I ALT

Vs

I

GA

Gs+GP cAp

I FLCH Pushbutton

VASEL CAP

I

VPTH

VNAV PuSHBUTTON

I

AUTOMATIC SELECTION

OF ANY VNAV MODE

I

FLCH?

s 1-VFLCH

Q

T

-1-JR

c1

vNAV TARGET VALI D

SDABC VALID

IRS MISCOMPARE ●

IRS 3 VALIDFGC VALID

SFMS VALID-=s

SG VALID FOR ~

COUPLED PFD

.

IIIIII

‘N PROCESSOR

+1

ASCB OUTPUT

‘A” 2 TO SYMBOL

PRCCESSOR GENERATOR TO4 DISPIAY SELECTED

5MODE ON PFD

IIIIIII

L J AD-13E62 ~-Rl@———— ———— ———— ———

F1ight DirectorMode Select Diagram

Figure 255 (Sheet 12)

22-14-00Page 298.69Apr 15/93


r FZ.820 FLIGNT GUIDANCE COMPUTER———— ————

IIIII[IIIIIII1IIIIIIIIIIIII

———— —.. — 1

DECREASINGALTITUDE ERROR

CMD ALT RATE > VS

2vs > 1,3 FT/SEC

3 SECS

AP ENG + FD STBY

ALT HOLD e

3ALT SEL CAP

VNAV ENG

GS CAP + TRK

VFLCH + VPATH

VASEL~ ARM

+ s b

~ RCLR

VASEL CAP

VALT d

PFD CMD SEL-j

zaa3--DETECTOR TRIPS

ALT SET KNOB

IN MOTION

PW IN MOTION

ANY OTHER VERT

P

MODE SEL + CAP

PFD CMD SEL

GA

FGC VALI

‘+

IRS MISCOMPARE .lRS3~

SOADC =

SG = FORCOUPLED PFD

OVERSPEED1

L———0—

VERTICAIJPITCH MODES

P 6AR0 ALT =

VASEL

F

~ CAPs

RCLR

IIIII1IIIIIIII

40J1 B/l 0J2B

+

1ASCB OUTWT

2“A”

TO SYMBOL


4 OISPIAY SELECTEDMOOE ON PFD

5

III

———— ———— ——— d

F1 ight DirectorMode Select Diagram

Figure 255 (Sheet 13) 22-14-00Page 298.70Aug 15/91


ALT

SWITCH

INPUT

(REF SH 1)

E-wiiFmHRuFANcE——— ——. ——— —.

mCOMPUTERI VASEL CAP

I

ALT ERROR< 25 FT

ALT RATE< 5 FT/SEC

I PWMOTION

ALT SEL KNOBMOTION P

IllJ1 1W2B ANY VNAV MODE

7 r7 w 100 + SERIAL IN

PARALLEL OUT

L >L

T

8 + 101 + DEVICE

m - 106 “A” PROCESSOR rR

III GA

I

PW MOTION

ANY OTHER VERT b

I F/D MODE ON+ CAP ‘

III

SDADC ~IRS MISCOMPARE ●

IRS 3 Vm 4

I FGCVALID ~I SFMS =

IVNAV TARGET V=

I

SG VALID FORCOUPLED PFD

BAROALT V=

IVNAV PUSHBUTTON-r

QUALT

I

IIIIIIII

“A”PROCESSOR 10JIB/1 W2B

t 7 ASCB OLJTPUT““A 2 TO SYMBOL

PROCESSOR GENERATOR TO4 DISPIAY SELECTED

5 MODE ON PFD

IL——— ——— ——— ——— ——— J

IIIIIIIII

AD-15662 ~-Rl@


22-14-00Page 298.71Apr 15/93


r-——— ——— —.— .—— — ——— ——. —] FZ-820 FLIGHT GUIDANCE

I COMPUTER

IIIIIIIIIIIIIIIIIII

TARGET ALT>150 ~ 4BELOW JVC ALT

FMS MODE ENGAGED

TOP OF DESCENT DEFINED

BOlXX4 OF DESCENT DEFINED

PATH ANGL= 6°b

VALT* PATH CAPTURE ZCNE

VFLCH* PATH CAPTURE ZONE

GA

PFD CMD

Gs CAP

VFLCH

VALT

OVERSPEED

?

sc

T

SFMS=

VERT TARGET VALID

SG =FOR C~PLED PFD

VNAV PUSHBUTT(34 GF *IRS MISGOMPARE ●

+IRS 3~D

VPTH

IIIII

‘A- PRCCE.SXR >1

10J1SW3J2B

T1

ASC8 OUTPUT

“A” 2 TO SYMSC)L

PRCY2ESSCRGENERATCR TO

4 DISPIAY SELECTEDMOOE CN PFD

5

rIIIIIIIII

L ——— ——— ——— ——— ..— ——— —— J AD-1 3662 (&Rl @


22-14-00Page 298.72Apr 15/93


Honeywell !!i$!!!!i.cE3. B. (4) Autopi1ot Engage Logic

The autopilot engage logic diagram is shown in Figure 256. Servoengagement is under software control, however, both hardware andsoftware monitors can either prevent engagement or cause adisengagement if a failure occurs. All logic gates shown inFigure 256 are actually software within the FGC.

(a) Pushbutton Switch Discrete Inputs

● GP-820 (Push On/Off)

- AP pushbutton

● Aircraft Panel (Push On/Off)

- YD pushbutton- Trim pushbutton

● Manual disconnect pushbuttons

- AP quick-disconnect- YD quick-disconnect (not used, wired to ground)- Trim quick-disconnect- Go-around- TCS- Manual trim switches (disconnectAP only)

. Miscellaneous Discrete Inputs

- Weight-on-wheels- Stall warning

(b) Servo Engage Control

Relays K2, K3, and K4 are used for servo engage control. Thepurpose of the servo engage control is to engage the servoclutches and activate servo brakes as well as completing themotor drive circuits. The servo engage control provides engagestatus information for use by the servo switching monitors andfor transmission on the ASCB. When relay K2 is energized, theaileron and elevator servo clutches are engaged and the aileronand elevator servo commands are allowed to drive the servo.The cross-side servo’s brakes are applied as well to preventthe cross-side from engaging. When relay K3 is energized, therudder actuator brake excitation is applied and the ruddercommand is allowed to drive the actuator. When relay K4 isenergized, the trim servo clutch is engaged and the trim servocommands are allowed to drive the servo. The cross-side brakeis applied as well to prevent the cross-side from engaging.Both fault warning computers receive AP and YD clutch inputswhich are used for computing the EICAS messages and the aurelalerts.

22-14-00Page 298.73Aug 15/91


Relay K4 is energized by the trim engage logic from ANDgate “C”. The processor valid and servo power enable inputsare partitioned between the A and B processors through ANDgates “D” and “E”. The trim enable is output from theA processor only through AND gate “B”. The trim engage commandinput can be from the M/T pushbutton or autopilot engagement.Trim engagement returns to the engage status that existed priorto AP engagement. AP servo power is also input to AND gate“C” through the K1 relay.

Relay K3 is energized by the YD engage logic from AND gate “G”.The processor valid and servo power enable inputs arepartitioned between the A and B processors. AND gate “D”derives the servo power enable input from the A and Bprocessor. Servo power enable and the processor valid outputfrom AND gate “E” is derived from the internal power supplymonitors, the A and B heartbeat monitors, and processor valids.The YD drive enable is output from the B processor only. YDengage can be reset by deselecting YD or by the pilot orcopilot YD disconnect inputs to the FGC.

Relay K2 is energized by the output from AND gate “J”. Theautopilot cannot be engaged unless the yaw damper is engaged,thus AND gate “J” needs the engage logic from AND gate “G” (yawdamper) and gate “F” (autopilot). Again, servo power enableand processor valid are partitioned between both processorsthru AND gates “D” and “E”. The AP drive enable is output fromthe B processor only. The autopilot cannot be engaged if theweight-on-wheels or stick shaker discrete inputs into the FGCare active. AP engage can be reset by deselecting AP, use ofthe pilot or copilot take-off/go-around (TOGA) switches, thepilot or copilot AP disconnect switches, or the pilot orcopilot electric trim switches.

3. B. (4) (c) Servo Power Enable Control

The purpose of the servo power control relay (relay Kl) is toprovide servo clutch power to the servo engage relays(relays K2, K3, and K4) to provide an alternate method of servodisengagement, and to provide cross-channel status validationlogic (relay-offcrossfeeds).

The servo clutch power is controlled by both processors viatheir respective servo power enable discrete outputs. Thehardware heartbeat and power supply must also be valid beforethe servo clutch power can be enabled.

Relay K1 must be energized before the autopilot, yaw damper, ortrim can be engaged via relays K2, K3, and K4. Relays K2, K3,and K4 are used for all normal engagements and disengagementsof the AP, YD, or trim. Relay K1 is used for all abnormaldisconnects. Whenever any failures are detected, relay K1 is

22-14-00Page 298.74Aug 15/91

Use or disclosure of information on this page is subject to the restrictions on the tnle page of thm document.

de-energized. The servo switching monitor in the FGC validatesdifferent engaged states against allowed states. Anytime anunallowed state is detected, such as AP engaged without YD(relay K2 energized but not K3), the servo switching monitorwill cause the servo power control relay to de-energize,causing the system to disengage and the FGC to become invalid.The servo switching monitor is discussed in detail insection 4.

3. B. (4) (d) Cross-Channel Discretes

The status of one FGC is conveyed to the cross-channel FGC viathe cross-channel discretes. The cross-side servo power offoutput from OR gate “A” will be active whenever the servo powercontrol relay is open, thus removing power from all clutches,and power to the clutches is switched off. Whenever thisoutput is active, the FGC is invalid and will not engage thesystem should the cross-channel FGC fail. The servo disconnectoutput from AND gate “H” will be active whenever the FGC doesnot have AP, YD, and trim engaged. As long as this output isnot active, this FGC is the priority channel and AP, YD, ortrim is engaged. The FGC also outputs the AP and YD powervalids to the cross-channel FGC.

(e) ASCB Engage Logic Data

The cross-channel engage status data listed below is used forengage synchronization and channel prioritization. All ASCBoutput data is controlled and formatted by the “A” processor.

● AP/YD/Trim Engage Status (Word 2)

- Off/Invalid- Off/Standby- Selected/Standby- Selected/Engaged

. Channel Priority Status (Word 1)

. AFCS Status to Fault Warning Computer (Word 37)

- TCS Status- AP/YD/Trim Status- Channel Priority

22-14-00Page 298.75/298.76

Aig 15/91Use or dmclosure of information on this page is subject to the restrictions on the title page of this document.

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——— ——— ——— .— --lAutopilot Engage Logic DiagramFigure 256 (Sheet 1)

AD.14619@R2

22-14-00Page 298.77/298.78

Apr 15/93


———— ——COPILOrS FZ-&!OFLIGHTGUIDANCE COMPUTER 1AP CROSS PWF! FDBK IYD CROSS PWR Ft)13K

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Autopi 1ot Engage Lcgic I)iagramFigure 256 (Sheet 2) 22-14-00

Page 298.79/298.80Aug 15/91. .

use or alsclosure 01 mlormallon on Ilws page IS subject 10 the reslrlctlons on the title pege of this document

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Autopi1ot Engage Logic DiagramFigure 256 (Sheet 3) 22-14-00

Page 298.81Aug 15/91

Use or disclosure of information on thw page is subject to the restrictions on the title page of this document.

3. B. (5) Monitors

The flight guidance computers incorporate internal monitors that aredesigned to disconnect the system (if only one FGC is valid) orautomatically switch to the second FGC (if two FGCS are valid) whena fault is detected. When a fault is detected, the bad FGC willautomatically force itself invalid (EICAS will annunciate FGC 1 FAILor FGC 2 FAIL).

After an FGC has detected a fault, it goes invalid. It will thenautomatically initiate a self-test. If this self-test passes, the

(5.1)

FGC will becbme valid again (the EICASautomatically clear). This process ofself-test after a detected fault is ca”time required for the monitor recoveryif the cross-channel FGC is valid, or ~cross-channel FGC is invalid (-905 FGC

FGC FAIL message” would-automatically initiating aled monitor recovery. Theto be performed is 7 seconds5 seconds if theonly).

Monitor recovery is initiated automatically except when a fault isdetected with only one FGC valid. In this case, the pilot mustinitiate the monitor recovery action by depressing either the pilotor copilot quick disconnect switch.

Prior to an FGC forcing itself invalid, the FGC writes the faultdata into memory. This data may be retrieved by accessing theflight fault summary section of the AFGCS test when performing theground maintenance test. The instructions for the groundmaintenance test are provided in section 4.

Monitor recovery is enabled in all modes except VOR/BC/LOC trackmodes. If an FGC detects a fault when in one of these modes,monitor recovery will not occur until after the track mode iscancelled. As long as both FGCS are valid when a fault is detected,there will be no loss of the active mode because the remaining FGCwould automatically perform all the functions that were beingperformed by the failed FGC.

If a particular fault results in an FGC performing monitor recoverymore than once in a 45-second period, the FGC will force itselfinvalid and will not perform additional monitor recovery attempts.

Power Interruptions

Figures 257 and 258 illustrate the engage timing requirements forvarious power interruptions and AFCS fault conditions.

22-14-00Page 298.82Apr 15/93


Honeywell &~$&.cE

WARM START COLD START—

AFCS WILL COME UP IN PREVIOUSLY ACTIVE MOOESALL AFCS MODES WILLHAVE TO BE MANUALLY RE-ENGAGEO

SERVOS WILL BE RE-ENGAGEDFROM THE PREVIOUSLY SERVOS WILL BE RE-ENGAGED

ENGAGED CHANNELFROM FIRST-UP CHANNEL

DISENGAGE ALL SERVOS

i INTERRUPTION

I I Ib OURATION

AO-10157

Power Interruptionof Both ChannelsFigure 257

22-14-00page 298.83


MAINTENANCE

Honeywell !Y!K&!#h.

I AFCS WILL COME UP IN PREVIOUSLY ENGAGED MODES I

I I THE INTERRUPTED CHANNEL

I WILL INITIATE COLD START

IAND WILL BE SYNCHRONIZED

SERVOS WILL RE-ENGAGE ITO THE MODES NOW ACTIVE

SERVOS WILL BE RE-ENGAGEDFROM FORMERLY

IN THE FORMER STANDBYFROM FORMER STANDBY I CHANNEL

ENGAGED CHANNEL I CHANNEL WITHIN 100 mSECI

I I

I I I, INTERRUPTION

I I I

w DuRATION

AD-101 58

Power Interruption of Engaged Channel OnlyFigure 258

22-14-00Page 298.84Aug 15/91


3. B. (6) Comparators

The FGC uses dual sensors and averages theinputs from each sensor are compared.

(a) DADC Comparator

data. The nonaveraged

The DADC comparator is not enabled unless at least one TASinput is greater than 50 kt. The different parameters comparedby the FGC are shown below.

Parameter Allowed Error Threshold

Dynamic Pressure 10 lb/ft2+qAvg/16

True Airspeed 10 kt + TAS Avg/32

Indicated Airspeed 10 kt + IAS Avg/32

Pressure Altitude 100 ft +ALTAvg/256

Mach 0.02 Mach (if both Mach >0.7)

A miscompare condition occurs when one of the above parametersdeviate greater than the allowed threshold. When a miscompareis detected in cruise modes or approach modes prior to theapproach track, the AP and YD will disengage and the DADCmismatch will be annunciated on EICAS, however, flight directormodes will be retained. A miscompare in approach track mode,the flight director mode will remain in approach track (APPT),however, reversion to constant speed gain programming willoccur in order to retain operational AP and YD.

(b) IRS Comparator

The different parameters compared by the FGC are shown below.


Pitch/Roll 6.

Pitch Rate 0.6°/sec +Avg Rate/16

Yaw Rate 0.6°/sec +Avg Rate/16

Roll Rate 3.0°/sec +Avg Rate /16

Normal Accel 0.2 g +Avg Normal Accel /16

Lateral Accel 0.1 g +Avg Delta Accel /16Fore-AFT Accel

Heading 6,

MAG Hdg 60

True Hdg 6*

True Track 6“

Gnd Speed 10 kt + Vgnd Avg/32

22-14-00Page 298.85Aug 15/91


A miscompare condition occurs when one of the above parametersdeviate greater than the allowed threshold. When a miscompareis detected, the FGC reversion will be dependent upon whetherthe IRS monitor resident in the AHRS or IRS 3 is valid. Withthe IRS monitor invalid, all FD modes will be dropped, the APand YD will disengage, and the IRS miscompare will beannunciated. A heading miscompare with the IRS monitor invalidwill cause the aircraft to revert to wings level from HDG holdand annunciate the IRS miscompare.

With the IRS monitor valid, a miscompare in the FGC will causethe FGC to look at the miscompare discretes from the IRSmonitor. The IRS monitor compares the parameters listed belowbetween IRS 1 and IRS 2.


Pitch/Roll 3°

Pitch RateYaw Rate 0.3°/sec +Avg Rate /32Roll Rate

Normal AccelLateral Accel 0.05 g + Avg Delta Accel /32Fore-Aft Accel

Heading N/A

A miscompare condition in the IRS monitor occurs whenever theabove parameters deviate greater than the allowed thresholdbetween the IRS monitor and IRS 1 or the IRS monitor and IRS 2.The IRS monitor will set the appropriate discrete to the FGC(either IRS 1 miscompare or IRS 2 miscompare).

The FGC does not use the miscompare discretes from the IRSmonitor until the FGC has detected a miscompare. At that pointthe FGC uses the IRS monitor as a voter by looking at themiscompare discretes. The FGC will stop using data from theIRS with which the IRS monitor has miscompared and EICAS willannunciate “FGC NOT USING IRS (X)”.

3. B. (6) (c) Radio Altitude Comparator

The parameter compared by the FGC is shown below.


Radio Alt 5 ft + ALT Avg /8

22-14-00Page 298.86Apr 15/93

USe or disclosure of information on this page is subject to the restrictions on the title page of this document,

3. B. (6) (d)

(7) Roll

(a)

Lateral and Vertical Approach Deviation Comparator

The parameters compared by the FGC are shown below. Thiscomparator is enabled only if in dual approach mode.


LOC/AZ Deviation 38 mV (1/2 dot)*GS/GP Deviation

*Miscompare detected only if opposed polarities and smallerdeviation is more than 10 mV. Miscompare retained unlessdelta between deviations is less than 20 mV.

Channel Functional Operation

Heading Select (HDG) Mode (See Figure 259, sheet 1.)

The heading select mode is used to intercept and maintain amagnetic heading. Activation of the HDG SEL pushbutton on theGP-820 Flight Guidance Controller selects the heading selectmode and overrides all roll active F/D modes. The HDG SELpushbutton annunciates ON. Selection of heading is made by aknob on the GP-820 and is displayed on the PFD and also on theND if the full compass is displayed. The heading select modeis annunciated on the PFD by a green HDG.

In the heading select mode, all armed roll F/D modes areallowed, but the capture of any armed roll mode will overridethe heading select mode. To allow initiation and continuationof the mode, the selected PFD data must be valid and the turnknob must be in detent.

To operate the mode, the heading bug on the PFD and ND ispositioned around the compass card to the heading the pilotdesires to intercept, using the heading knob on the GP-820Flight Guidance Controller. The heading select signal from theGP-820 to the FZ-820 Flight Guidance Computer represents thedesired aircraft heading. The signal is routed from theFZ-820 on the ASCB to the SG-884 Symbol Generator. In thesymbol qenerator, the desired aircraft headinq is com~aredagainst-actual aircraft heading, anderror signal is routed to the FZ-820through the ASCB.

In the flight guidance computer, theTAS (true airspeed) gain programmed.performed on the heading error signal to achieve approximatelythe same aircraft response, regardless of the aircraft’sairspeed and altitude. The TAS computation is derived fromairspeed and barometric altitude information provided from theAZ-81O Digital Air Data Computer, through the ASCB.

the resuitant headingFlight Guidance Computer

heading error signal isTAS gain programming is

22-14-00Page 298.87Apr 15/93


I

3. B. (7) (b)

From the TAS gain programmer, the heading select command isrouted to Figure 259, sheet 4, and is processed as discussed inparagraph 3.B.(7)(g).

The BANK pushbutton on the GP-820 allows manual control of thebank angle limit in the heading select mode only. Alternateactivation of the BANK pushbutton causes alternate selection ofa high bank angle limit (28 degrees) and a low bank angle limit(20 degrees). Power-up state is high bank angle unlessaltitude is greater than 20,000 feet at power-up. Once a bankangle limit is selected, it is retained in memory for theheading select duration of the flight; that is, reselection ofheading select mode will not deselect the bank angle valuestored in memory. When the BANK pushbutton is pressed andannunciates ON, low bank is selected.

Climbing through 29,500 feet with wings level willautomatically select the low bank angle limit, if it is notalready selected. The low bank angle limit can be reselectedby pressing the BANK pushbutton.

Descending through 28,500 feet with wings level willautomatically select the high bank angle limit if it is notalready selected. The low bank angle limit can be reselectedby pressing the BANK pushbutton.

VOR (LNAV) Mode (See Figure 259, sheet 2.)

The VOR mode provides for automatic intercept, capture andtracking of a selected VOR radial, utilizing the selectednavigation source displayed on the active PFD and ND. Thenavigation source displayed on the PFD and ND is a function ofthe NAV source selected on the DC-884 Display Controller.Prior to engaging the mode, the pilot would perform thefollowing:

. Tune the navigation receiver to the desired VOR frequency.

● If the station is listed as a VORTAC (combination of VOR andTACAN), then tune the DME receiver to the station frequencyto obtain distance to the station information.

. Set the course pointer for the desired course to be flowntoward or away from the station.

● Set the heading bug for the desired heading intercept forthe selected course, since the heading select mode is usedto achieve the VOR intercept.

● Select NAV 1 or NAV 2 as the navigation source on the DC-884Display Controller.

22-14-00Page 298.88

Apr 15/93Use or disclosure of informationon this page is subject to the restrictions on the title page of this document,

With the aircraft outside of the normal capture range of theVOR signal (typically the CDI on the PFD and ND is greater thantwo dots), the pilot selects the LNAV button on the GP-820Flight Guidance Controller. At this time, the PFD willannunciate HDG in green and VOR in white. The FZ-820 FlightGuidance Computer is now armed to capture the VOR signal and isgenerating a roll command to fly the heading select mode aspreviously discussed in paragraph 3.B. (7)(a).

When reaching the lateral beam sensor (LBS) trip point, thesystem automatically drops the heading select mode andswitches to the VOR capture phase. The following is observedon the PFD:

. The white VOR annunciator extinguishes.

. The green HDG annunciator extinguishes.

. A green VOR is annunciated and is boxed for 5 seconds toemphasize the capture phase of operation.

The FZ-820 now generates the proper roll command to bank theaircraft to capture and track the selected VOR radial. Whenthe course select pointer was set on the ND using the courseknob on the GP-820 Flight Guidance Controller, the courseselect error signal was established. This signal representsthe difference between the actual aircraft heading and thedesired aircraft course. The course error signal is then sentfrom the DC-884 Display Controller to the symbol generator andto the FZ-820 through the avionics standard communications bus(ASCB). Next, the course error signal is TAS (true airspeed)gain programmed. TAS gain programming of the course errorsignal is performed to achieve approximately the same aircraftresponse for a given command, regardless of the aircraft’sairspeed and altitude. The TAS computation is derived fromairspeed and barometric altitude information provided from theAZ-81O Digital Air Data Computer through the ASCB. From theTAS gain programmer, the course error signal is summed withradio deviation.

The radio deviation signal is routed from the navigationreceiver to the symbol generator on the ARINC 429 bus. Fromthe symbol generator, the radio deviation signal is routed tothe FZ-820 through the ASCB, where the signal is lateral gainprogrammed.

The lateral gain programming is performed as a function ofDMEdistance to the station and barometric altitude. This gainprogramming adjusts for the aircraft either coming toward ormoving away from the VOR station. The DME compensation circuitapproximates ground range to the station for more accurate gainprogramming and to help calculate over station sensing (0SS).

22-14-00Page 298,89Aug 15/91

Use or disclosure Of Information on this page is subject to the restrictions on the title page of this document.

I

From the lateral gain programmer, the radio signal is filteredand summed with the course error signal. The sum of courseerror and radio deviation is then sent to a course cut limiter.

The course cut limiter functions primarily when approaching thedesired VOR radial at an intercept angle greater than 45degrees and at high speed. Its function is to limit steeringcommands to 45 degrees which forces a flightpath to get on theselected radial sooner to prevent overshooting beam center.Typically, the roll command will make an initial headingchange, then level out and fly toward the beam, then make asecond heading change to get lined up on the center of theselected radial.

When the aircraft meets the VOR ON TRACK criteria, the (VORTRK . CE < 45”) switch opens. This removes course error fromthe roll command, leaving radio deviation, roll attitude andlateral acceleration from the IRS, to compensate for beamstandoff in the presence of a crosswind.

As the aircraft approaches the VOR station, it will enter azone of unstable radio signal. This zone of confusion radiatesupward from the station in the shape of a truncated cone. Inthis area, the radio signal becomes highly erratic and it isdesirable to remove it from the roll command. The over stationsensor monitors for entry into the zone of confusion and opensthe 0SS switch, removing radio deviation from the roll command.

From the course cut limiter, the VOR SEL command is routedto Figure 259, sheet 4, and is processed as discussed inparagraph 3.B. (7)(g).

Should the pilot select FMS as his navigation source on theDC-884 Display Controller, the system flys desired track frompresent position to the next waypoint, with the followingdifferences:

. Instead of using course error and radio deviation from thesymbol generator, a composite lateral steering command isutilized from the NZ-920 Navigation Computer.

. This lateral steering command is lateral gain programmedin the NZ-920 and therefore is not gain programmed again inthe FZ-820.

. When FMS is selected as the navigation source on the DC-884to perform the VOR intercept, the mode annunciation on thePFD will be FMS.

22-14-00Page 298.90


3. B. (7) (c)

● FMS will be displayed in white on the PFD during the armphase of operation. At LNAV capture, FMS will be displayedon the PFD in green and is boxed for 5 seconds to emphasizethe capture mode.

Another option the pilot has when flying a VOR intercept, is tofly a zero deviation approach to the VOR station. TO do this,th~ pilot would do the-following:

. Establish the VOR intercept as previously

. On the DC-884 Display Controller, set theto the same NAV source being used for the

discussed.

bearing selectorVOR intercept.

The PFD and ND will now display the following:

● The course pointer will display the pilot selected course tothe VOR station.

. The bearing pointer will display a zero deviation course tothe VOR station.

Should the pilot desire to fly the zero deviation course to thestation, he would push the SYNC button on the GP-820 CRS knob.This will cause the course select pointer to align with thebearing pointer and provide for a zero deviation course to beflown to the VOR station.

Localizer Mode (See Figure 259, sheet 3.)

The localizer mode provides for automatic intercept, capture,and tracking of the front course localizer beam, to line up onthe centerline of the runway in use. Prior to mode engagement,the pilot would perform the following:

● Tune the navigation receiver to the published front courselocalizer frequency for the runway in use.

● Set the course pointer for the inbound runway heading.

. Set the heading bug for the desired heading to perform acourse intercept.

● Select NAV 1 or NAV 2 as the navigation source on theDisplay Controller.

The PFD and ND now display the relative position of the

DC-884

aircraft to the center-of-the localizer beam and the desiredinbound course. With the heading bug set for course intercept,the heading select mode is used to perform the intercept.

22-14-00Page 298.91Apr 15/93

Useor disclosureof informationonthispage issubject to the restrictionsonthetitle page of thisdocument.

Honeywell &!#%.cEOutside the normal capture range of the localizer signal(between one and two dots), pressing the LNAV button on theGP-820 Flight Guidance Controller will cause the PFD toannunciate:

● LOC in white.

● HDG in green.

The aircraft is now flying the desired heading intercept andthe system is armed for automatic localizer beam capture.

With the aircraft approaching the selected course intercept,the lateral beam sensor (LBS) is monitoring localizer beamdeviation, beam rate, and TAS. At the computed time, the LBSwill trip and capture the localizer signal. The flightguidance computer now drops the heading select mode andgenerates the proper roll command to bank the aircraft towardlocalizer beam center. When the LBS trips, the PFDwilldisplay LOC in green and it is boxed for 5 seconds to emphasizethat the capture phase has occurred.

As the aircraft continues toward localizer beam center, thecomputer enters the LOC CAP 2 submode. With the aircraftalmost lined up on localizer beam center, the computer willautomatically change to the LOC TRACK 1 and the LOC TRACK 2submodes. The LOC CAP 2 and LOC TRACK submodes provide tightercontrol law programming on the localizer signal to bettermaintain a truer flightpath along the localizer beam. Thereare no visual indications in the cockpit that these submodeshave occurred.

When the course select pointer was set on the PFD using thecourse knob on the GP-820 Flight Guidance Controller, thecourse select error signal was established. This signalrepresents the difference between actual aircraft heading anddesired aircraft course.

The course select error signal is routed from the DC-884Display Controller to the FZ-820 Flight Guidance Computer andthe SG-884 Symbol Generator through the avionics standardcommunications bus (ASCB).

Next, the course error signal is TAS.(true airspeed) gainprogrammed. TAS gain programming is performed to achieveapproximately the same aircraft response for a given command,regardless of the aircraft’s altitude and airspeed. The TAScomputation is derived from airspeed and barometric altitudeinformation provided by the AZ-81O Digital Air Data Computerthrough the ASCB. After being TAS gain programmed, the courseerror signal is summed with radio deviation information.

22-14-00Page 298.92Aug 15/91

Use or disclosure of information on this page is subject to the restrititons on the title page of this document.

The radio deviation signal is routed from the navigationreceiver to the symbol generator on the ARINC 429 bus. Fromthe symbol generator, the radio deviation signal is routed tothe FZ-820 through the ASCB, where the signal is lateral gainprogrammed.

Lateral gain programming is required to adjust the gain appliedto the localizer signal due to the aircraft approaching thelocalizer transmitter and beam convergence caused by thedirectional qualities of the Iocalizer transmitter. Thelateral gain programmer is controlled by a distance fromtransmitter estimator. The distance estimator is actually alow pass filter and rate limiter with two modes of operation:

● A calculated range mode

. An estimated range mode

If both radio altitude and glideslope deviation are valid, thendistance is calculated using radio altitude and glideslopedeviation data. If only radio altitude is valid, distance isfirst estimated for capture and then, when in the final track 2mode, it is assumed that an approach to the runway is beingmade without glideslope, and distance is calculated based onradio altitude only.

If radio altitude information is not valid, then distance isestimated as a function of glideslope deviation and TAS. Ifneither radio altitude nor glideslope data is valid, thendistance is estimated as a function of TAS and time.

From the lateral gain programmer, the localizer signal isfiltered, amplified and summed with the course error signal.The resultant localizer command signal is then course cutlimited. The course cut limiter functions primarily whenapproaching localizer beam center at an intercept angle greaterthan 45 degrees and at high speed.

The course cut limiter’s function is to limit steering commandsto 45 degrees, which forces a flightpath to get onto localizerbeam center sooner. This is done to prevent overshootinglocalizer beam center during the capture phase of operation.Typically, the roll command will make an initial headingchange, then level out and fly toward the beam. At thecomputed time, the roll command will initiate a second headingchange to line up on the localizer beam.

When the aircraft meets the LOC track conditions (LOC TRK + BCTRK ● CE < 450), the switch opens. This removes course errorfrom the roll command, leaving radio deviation, roll attitude,and lateral acceleration to compensate for beam standoff in thepresence of a crosswind.

22-14-00Page 298.93Aug 15/91

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3. B. (7) (d)

From the course cut limiter, the localizer command is routed toFigure 259, sheet 4, and is processed as discussed in paragraph3.B.(7)(g).

Localizer Approach Mode (See Figure 259, sheet 3.)

The localizer approach mode provides for automatic intercept,capture, and tracking of the front course localizer and glide-slope signals. This allows the pilot to fly a fully coupledILS approach. The mode is set up and flown exactly like thelocalizer mode, with the following differences:

● On the GP-820 Flight Guidance Controller, the APR button isselected.

With the aircraft outside the normal localizer capture limits,the PFD will annunciate the following modes at this time:

. HDG in green.

● LOC in white.

. GS in white.

As with the localizer mode, heading select is usedthe localizer approach intercept.

Any other vertical mode in use at this time will a-annunciated on the PFD. At localizer capture, theannunciate:

to initiate

so bePFD wi11

. LOC in green and is boxed for 5 seconds to emphasize thecapture-phase.

● GS in white.

. Any other vertical mode in use at the time.

The flight guidance computer now generates a roll command tosmoothly capture and track the localizer signal. With thelocalizer signal captured, the aircraft proceeds inbound and,at the computed time, will automatically capture and track theglideslope signal. The GS capture overrides all vertical modeswhich were previously engaged. At this time, the PFD willannunciate:

c LOC in green.

. GS in green and is boxed for 5 seconds to emphasize thecapture phase.

The aircraft is now flying a fully coupled ILS approach.

22-14-00Page 298.94

Aug 15/91Use or dleclosure of information on this page is subject to the restridlons on the tiile page of this document.

3. B. (7) (e) Back Course (BC) Mode (See Figure 259, sheet 3.)

The back course mode provides for automatic intercept, capture,and tracking of the back course localizer signal. When flyinga back course localizer approach, glideslope capture isautomatically inhibited. The back course mode is set up andflown exactly like a front course localizer approach, with thefollowing differences:

. On the GP-820 Flight Guidance Controller, the BC button isselected.

With the aircraft outside the normal localizer capture limits,the PFD will annunciate:

●

●

At

●

HDG in green.

BC in white.

localizer capture, the PFD will annunciate:

BC in green and is boxed for 5 seconds to emphasize that thecaptur~ phase has occurred.

tlhenthe back course mode was selected on the GP-820, logicin the flight guidance computer was established to internallyreverse the polarity of the course error and localizer signals.Additionally, a gain change takes place in the computer whenBC is selected, since the aircraft will now be closer to thelocalizer transmitter by the length of the runway plus1000 feet.

At back course capture, the flight guidance computer willgenerate a roll command to smoothly capture and track the backcourse localizer signal. In addition, the previous roll modeis cancelled. To allow initiation or continuation of the mode,the selected PFD and NAV data must be valid.

(f) Category 2 - Optional

A category 2 approach is flown the same as an ILS approach.Refer to paragraph 3.B.(7)(d).

(g) Roll Autopilot Mode Flow (See Figure 259, sheet 4.)

The roll autopilot diagram shows two signal paths for thelateral steering command. The first path is with the autopilotdisengaged and routes the lateral steering command to thesymbol generator only. This path is discussed in paragraph3.B.(7)(g)~. The second path is with the autopilot engaged,and routes the lateral steering command to both symbolgenerators and to the aileron servo drive motor. This path isdiscussed in paragraph 3.B.(7)(g)~.

22-14-00Page 298.95Aug 15/91

Use or disclosure Of information On this page is subject to the restrictions on the title page of this document.

Honeywell !i!gr.c’

3. B. (7) (g)

When the autopilot is engaged and nomode is selected, the autopilot willwinas level attitude. When the bank

lateral flight directordrive the aircraft to aanale is less than

6 d;grees, the autopilot will fly headi~g hold. Heading holdis discussed in paragraph 3.B.(7)(g)a. The roll hold mode ofoperation is discussed in paragraph 3.B.(7)(g)~, turn knob isdiscussed in paragraph 3.B.(7)(g)s, and the go-around mode isdiscussed in paragraph 3.B.(7)(g)~.

~ Lateral Steering Command with Autopilot Disengaged

With the autopilot disengaged, the selected flight directorsteering command is routed through the following:

● Rate Limiter. Bank Angle Limit● Roll Hold Switch. Summing Point● AP Engage Switch. Roll Bar Bias Switch

The steering command is then routed to the symbol generatorthrough the ASCB to drive the flight director lateralsteering cue on the PFD.

As the pilot banks the aircraft to follow the steeringcommand, roll attitude information is provided to thesummation point from the IRS through the AP ENG switch.As the steering command is satisfied, the flight directorlateral steering cue is centered on the PFD, and theaircraft is now flying the flight director command.

As the aircraft approaches the selected heading or radiobeam center, the flight director command diminishes in sizeand the roll attitude signal predominates. This causes theflight director lateral steering cue to move out of centerin the opposite direction to the original command. As thepilot flys the aircraft to satisfy the command, thefollowing is taking place:

. The aircraft is rolling back to a wings level attitude.

. The flight director command is satisfied.● The roll attitude signal is going to zero.● The flight director lateral steering cue is returning to

center.

Roll bar bias (RBB) is a fixed voltage level used to biasthe flight director lateral steering cue out of view toprevent the pilot from flying invalid data. The lateralsteering cue is biased out of view if:

c No flight director mode is selected.. The flight director is not valid.

22-14-00Page 298.96Aug 15/91

Use or disclosure of information on this page is subpct to the restrictions on the title page of this document.

Flight director valid is comprised

. CPU “A” heartbeat monitor valid

of the following:

which Drovides validity

logic to engage/disengage the autopilot functions. Th~heartbeat monitor is hardware independent from the CPU,so that no single fault can disable both the CPU and themonitor. The heartbeat monitor output is provided as aninterrupt to the nonmonitored processor.

● Flight director flag/annunciator valid which is a directdiscreet output from the “A” CPU to drive the FD flag onthe PFD and enable the annunciator drivers on the sameside guidance controller channel.

. FGC power supply valid which monitors the internal powersupply voltages for proper operating levels.

3. B. (7) (g) Z Lateral Steering Command with Autopi1ot Engaged

With the autopilot engaged, the selected flight directorsteering command is routed through the:

. Rate Limiter

. Bank Angle Limiters Roll Hold Switch● First Summation Point. Second Summation Point

From the second summation point, the steering command israte limited again, acceleration and bank angle limited, andthen split into two paths. The first path goes up throughthe activated AP ENG switch, the roll bar bias switch andthen to the symbol generator.

The second path routes the steering command to a M degreesposition command limiter, is gain adjusted and applied to apulse width command limiter and output servo amplifier.

The pulse width command limiter serves two functions.First, it is the D/A converter for the servo amplifier.Second, as a motor driver it issues a continuous string of28 V dc pulses, at a rate of one pulse approximately every2 milliseconds. The pulse width of the pulses is determinedby the lateral command. The pulse width command limiter hasits current limits established by a software programcontained in the “B” CPU.

The output of the servo amplifier is sent to the SM-600aileron servo drive motor and to the “A” processor currentmonitor to check for servo runaway current. The SM-600 is apermanent magnet dc motor that utilizes a dc tachometer forrate feedback. It has but does not use a position feedbacksynchro. As the servo motor drives to position the aileron,

22-14-00Page 298.97Aug 15/91


it also drives the dc tach generator through mechanicalcoupling (represented by a dotted line). The tach generatorprovides a rate feedback signal that serves two functions.First, it acts as a damping term when summed with thelateral command input to the pulse width command limiter.This helps to stabilize aileron position and minimizeexcessive aileron travel.

Second, the rate feedback signal is integrated to obtainposition feedback, gain adjusted and summed with thesteering command. When these signals are equal, the aileronis in the proper position to satisfy the steering command.As the aircraft responds, roll attitude and roll rateinformation provided by the IRS, are summed with thesteering command. This allows the feedback position signalto drive the aileron back to its original position.

As the steering command is satisfied and diminishes in size,the roll attitude signal becomes dominant and provides acommand to move the aileron in the opposite direction, toreturn the aircraft to a wings level attitude. The servoloop follow up would be identical to that just discussed.

If the summation of command and roll attitude are notexactly equal, the difference between the two signals issent to the command rate taker. The signal is changed torate, and summed with tach generator rate feedback. Thesumming of these two signals is then integrated to obtainposition data and summed with the steering command. Thisboost helps eliminate flightpath or attitude standoffs.

3. B. (7) (g) ~ Wings Level and Heading Hold Mode

If the autopilot is engaged and no flight director mode hasbeen selected, then a zero roll command becomes the desiredsteering command. This zero command is routed through thefollowing:

. Rate Limiter

. Bank Angle Limiter

. Roll Hold Switch

. First Summation Point

. Second Summation Point

At the second summing point, roll attitude provided by theIRS is added to the command. Since the command is zero, thesumming point’s output is an attitude displacement signalthat is rate limited, acceleration limited, and summed withroll rate, provided by the IRS.

This attitude signal is now routed to the aileron servodrive as discussed in paragraph 3.B.(7)(g)~.

22-14-00Page 298.98Aug 15/91


Honeywell !g!!gH#h---

3. B. (7) (9) 4

As the aircraft rolls to a wings level attitude, the systemwill automatically switch to the heading hold mode. Headinghold is defined as:

● No lateral flight director mode selected.● Bank angle less than 3 degrees and 10 seconds.

When in HDG HOLD (refer to Figure 259, sheet 1), headinginformation from the IRS is routed to the heading holdreference synchronizer. When the system recognizes theheading hold criteria as being true, the HDG HOLD switchopens, locking the reference heading in the synchronizer.Actual heading informatifheading. Any differenceprogrammed and routed tohold error signal is nowas previously described ~

Roll Hold Mode

The autopilot recognizesoperational when:

n is now compared to the referencebetween the two is TAS gainFigure 259, sheet 4. The headingrouted to the aileron servo drive,n paragraph 3.B.(7)(g)~.

the roll hold mode as being

● No lateral fliaht director mode is selected.● The aircraft b~nk angle is greater than 6 degrees.. Touch Control Steering (TCS) was used to initiate the

bank maneuver.. Turn knob is not active.

The roll hold mode can be used by the pilot to maneuver theaircraft into a bank and utilize the autopilot to hold thebank angle.

With the roll hold criteria being met, roll attitudeinformation from the IRS is entered into the roll holdreference block. With the ROLL HOLD switch activated tothe up position, desired roll attitude is compared againstactual roll attitude at the second summation point. Sincethese signals are equal and opposite, no command is issuedto the aileron servo drive, and the autopilot maintains thedesired bank angle.

The roll hold reference block is a synchronizer that locksits output as a function of releasing TCS with theaircraft’s bank angle greater than 6 degrees. Should atransient wind gust occur, the desired bank angle remainsunchanged, while the actual bank angle changes with aircraftmovement. The difference between the two becomes an errorsignal routed to the aileron servo drive, as described inparagraph 3.B.(7)(g)~, to fly the aircraft back to thereference bank angle.

22-14-00Page 298.99Aug 15/91

Use or disclosure Of Information on this page IS subject to the restrictions on the title page of this document.

3. B. (7) (g) 5 Turn Knob

When the TURN knob is rotated out of detent, with theautopilot engaged, any previously selected flight directorroll mode is cancelled and a turn knob signal is applied tothe FZ-820 Flight Guidance Computer to change the aircraft’sroll attitude reference (refer to Figure 259, sheet 1). Theturn knob command is limited to t30 degrees by knobexcitation and is routed to Figure 259, sheet 4. The turnknob command is then routed to the aileron servo drive, aspreviously described in paragraph 3.B.(7)(g)~.

Returning the TURN knob to detent will return the airDlaneto a win~s-level condition andmode. Other lateral modes canknob is returned to the detent

~ Go-Around (Wings Level)

The go-around mode is normally

then engages the headi;g-holdbe selected after the TURNposition.

used to transition from anILS approach to a climb out condition when a missed approachhas occurred. The pilot selects go-around by pressing theTOGA button located on either outboard throttle handle.With go-around selected, all armed and active flightdirector modes are cancelled, and the autopilot isdisengaged. Laterally, the pilot sees a wings level commandon the PFD.

After go-around mode selection, HDG SEL mode can be selectedto cancel the wings level roll command.

The go-around mode is cancelled by selecting another pitchmode, engaging TCS, or engaging the autopilot.

The go-around mode is annunciated on the PFD by a green GA.

22-14-00Page 298.100


-——— ——— --—— —— ——— —— ——— ——— ——— ———FZ-SZOPLIGNTGUIDANCE COMPUTER17&ooTuRNPITcHc0N7R0LLER 1 w

——— ———10.ll B/10J2B

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1.l’HE FOLLOWING ASCB “A AND “W DATA (SH1THRU 4) IS AVERAGED IN THE FZ-820

● ROLL ATTITUDE● ROLL RATE

TO SH 1

r———

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III

———

(ASCB “K DATA ‘415

ASCB “W DATA

(

12

13

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● HEADING ERROR● COURSE ERROR

(NOTE 1)

● HEADING● LATERAL STEERING CMO● DME. ALTITuDE● RADIO OEVIATION

IF ONE OF THE SIGNALS IS INVALIO AND AVERAG.ING IS NOT POSSIBLE. THEN THE SYSTEMSWITCHES TO SINGLE SIDE OPERATION USINGTHE SIGNAL FROM THE VALID SENSOR.

2 THE SWITCH NOMENCLATURE CAUSES THESWITCH TO CHANGE FROM STATE SHOWN

P=sY=OL=N~TZ – 65J

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——— ——— —.. — ——— ——— ——— ——— ——— ——— ——AD 13663 @ R4

HDG SELECT AND HDG HOLO MODES

F1ight Director/AutopilotRoll Channel Mode Flow Diagram

Figure 259 (Sheet 1) 22-14-00Page 298.101/298.102

Aug 15/91Use or dwclosure of mformallon on Ihls page KSsub]ect 10 the restrlcllons on the title page of this document

.——. ———— ———— ———— ———— —.. . ———— ———— —— __ ____ _Z-ESOfl.2QEIT2sU10ANCECOMPUTER

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Apr 15/93


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Figure 259 (Sheet 3) 22-14-00Page 298. 105/298. 106

Apr 15/93


Honeywell #.#!r.cE3. B. (8) Pitch Channel Functional Operation

(a) ~~li~~tlD~ rector Pitch Attitude Hold Mode (See Figure 260,.

The pitch attitude hold mode is the basic vertical flightdirector mode. It is activated when a flight director rollmode is selected without an accompanying pitch mode and is notannunciated on the PFD. The pitch command on the PFD providesthe pilot with a pitch reference corresponding to the pitchattitude existing at the moment the roll mode was selected.This pitch reference may be changed with the TCS button locatedon the pilot’s and copilot’s control wheel.

The reference pitch attitude may also be changed as a functionof the pitch wheel on the PC-880 Turn Pitch Controller when theautopilot is engaged. (Refer to discussion in paragraph3.B.(8)(i)~.) Prior to the mode being operative, IRS pitchattitude information is applied to a summation pfit and thenrouted through a closed [PITCH HOLD + (AP ENG ● TCS ● NO VERTF/D MODE)] switch to the input of a synchronizer. The outputof the synchronizer is of opposite polarity to the pitchattitude signal, and therefore the two signals cancel eachother. This results in a zero signal out of the summationpoint.

When only a lateral fligmdirector mode is selected, the[PITCH HOLD + (AP ENG ● TCS ● NO VERT F/D MODE)] switch opens.This clamps the synchronizer output as a reference for thepitch hold mode. As long as the pitch attitude of the aircraftremains unchanged, there will be no command to drive the pitchcue on the PFD. If the aircraft deviates from the referenceattitude established at mode engagement, an error signalcorresponding to the difference between the actual aircraftattitude and the reference attitude will drive the pitch flightdirector cue in the proper direction to fly the aircraft backto the pitch reference attitude.

As the aircraft responds to the command, the error signaldiminishes and the flight director pitch cue on the PFD returnsto center.

The pitch reference signal is limited to +20, -15 degrees andthen routed to Figure 260, sheet 7, and is processed asdiscussed in paragraph 3.6.(8)(i).

Lift compensation is provided to hold the aircraft’s nose up ina turn commanded through the autopilot. Roll attitude from theIRS is routed to a block where the cosine of the bank angle issubtracted from one. The resultant output will be routed tothe pitch autopilot to keep the nose of the aircraft up duringthe turn maneuver. The lift compensation function is inhibitedduring the FLCH mode of operation.

22-14-00Page 298.109


Lift compensation is also provided as a function of flapposition and transitions. This is accomplished to minimize

Iaircraft pitch attitude changes due to the aerodynamic liftchanges induced through the deployment of flaps.

3. B. (8) (b) Vertical Speed (VS) Hold Mode (See Figure 260, sheet 2.)

Activation of the VS pushbutton on the GP-820 selects theVS hold mode and overrides all active pitch F/D modes. Thevertical speed hold mode is annunciated on the PFD by agreen VS.

In the VS hold mode, all armed pitch F/D modes are allowed, butthe capture of any armed pitch mode will override the mode. Toallow the initiation or the continuation of the mode, theselected DADC must be valid.

The vertical speed hold mode is used to automatically maintainthe aircraft at a pilot selected vertical speed reference. Thevalue set with the VERT SPEED knob on the GP-820 FlightGuidance Controller is displayed on the GP-820 and PFD. Toinitiate the mode, the pilot would maneuver the aircraft to thedesired climb or descent attitude, establish the vertical speedreference, and engage the mode. The reference vertical speedmay be changed by use of the VERT SPEED knob or by pressing theTCS button on the control wheel and maneuvering the aircraft toa new vertical speed reference and then releasing the TCSbutton.

Prior to mode engagement, altitude rate (VS) information,provided by the AZ-81O DADC through the ASCB, is fed to asumming point where it is compared to selected vertical speed.Should a difference between these signals occur, the differenceis TAS gain programmed and routed as a VS command signal toFigure 260, sheet 7. TAS gain programming accurately adjuststhe VS command signal as a function of the aircraft’s currentspeed and barometric altitude. The VS command signal willdrive the flight director pitch command cue on the PFD in theproper direction to fly the aircraft back to the pilotselected vertical speed. As the aircraft returns to thereference vertical speed, the VS command will decrease towardszero. The aircraft has now returned to the selected verticalspeed reference. For discussion of the vertical speed commandsignal as it is processed on Figure 260, sheet 7, refer toparagraph 3.B.(8)(i).

22-14-00Page 298.110

Apr 15/93

Useor disclosure of information on this page issubject totherestrictions on the title page of this document.

3. B. (8) (c) F1ight Level Change (FLCH) Mode and VNAV F1ight Level Change(VFLCH) Mode (See Figure 260, sheet 3.)

Activation of the FLCH pushbutton on the GP-820 Flight GuidanceController selects the flight level change mode and overridesall active pitch F/D modes. The FLCH mode will fly to thespeed target which is displayed on the PFD and in the GP-820speed window.

Activation of the VNAV arm and FLCH pushbuttons on the GP-820Flight Guidance Controller selects the VNAV flight level changemode and overrides all active pitch F/D modes. The VFLCH modemay also be selected automatically by the FMS. The VFLCH modewill fly to the speed target which is displayed on the PFD andin the GP-820 speed window.

The speed target is GP-820 switch selectable by the pilot to beeither IAS or Mach and automatic or manual. Automatic speed issupplied by the FMS and manual speed is set by the pilot withthe GP-820 speed knob.

The airspeed target display on the GP-820 Flight GuidanceController functions in automatic and manual (speedintervention)modes as follows:

. Automatic Mode - The displayed speed target is input fromthe selected FMS which also controls the IAS/Mach displayformat.

. Manual Mode - Manual mode (speed intervention) is selectablevia the SPD pushbutton on the GP-820. The IAS/Mach displayformat will not change when this mode is selected, but thespeed target will be synched to present speed. In thismode, the SPD button will be annunciated (MAN), and thetarget display format will toggle selectable via theIAS/Mach changeover pushbutton. Alterations of the targetare accomplished via the speed-set knob inputs. Continuityof speed shall be maintained when toggling between IAS andMach targets, via the changeover button.

The PFD annunciates Mach or IAS (FLCH), orVMACH or VIAS(VFLCH), and the GP-820 SPD switch annunciates MAN when in themanual mode. Before using the FLCH or VFLCH mode, the desiredaltitude is set with a knob on the GP-820 and it is displayedin the GP-820 altitude window.

When autothrottles are engaged, they provide proper powermanagement for the desired altitude and airspeed change.Without autothrottles, the pilot provides the power control.

22-14-00Page 298.111

Aug 15/91Use or disclosure of reformation cm thw page is subpct to the restrictions on the title page of this document.

The FLCH mode is annunciated on the PFD by a green IAS or MACH.The VFLCH mode is annunciated on the PFD by a green VIAS orVMACH . In FLCH and VFLCH, all armed pitch flight directormodes are allowed, but the capture of any armed pitch mode willoverride the mode.

The FLCH and VFLCH mode is set up to change level from presentaltitude to the preselected altitude. It will try to maintainthe speed reference over the long term and allow vertical speedto change, as a function of power setting. For example,throttle retard in a climb will cause the system to track thespeed reference while bleeding off vertical speed.

However, vertical speed will not be allowed to go below zero.If throttle retard in a climb was enough to cause verticalspeed to go to zero, vertical speed would be held to zero andthe aircraft would decelerate. At this time, the flightdirector command cue on the PFD would be out of center,indicating that power should be applied to track the FLCH orVFLCH speed reference.

In the FLCH or VFLCH mode, the AFCS should fly to the newpreselect altitude at the target speed from the FMS or asmanually selected on the GP-820, when aircraft thrust is setappropriately for climb or descent. When the power is not setappropriately, then the AFCS should maintain zero verticalspeed in order to not fly away from the preselected altitude.If the target speed is changed from IAS to MACH (or viceversa), the FLCH or VFLCH mode should remain engaged and fly tothe appropriate new speed target.

The DAFCS pitch guidance will @ generate commands to exceedv or M~Q when in FLCH or VFLCH mode. This speed limitingf~~ction Includes anticipation to account for flight envelopev and/or MH

?discontinuities. The vertical guidance will not

c&mand more han tO.2 g in FLCH and VFLCH modes.

Once in the FLCH mode the target speed can be changed via theGP-820 SPEED knob if using manual speed.

NOTE : Depressing the GP-820 FLCH button will automaticallyselect either IAS or MACH modes as function of aircraftpressure altitude. At altitudes greater than 27,850feet, Mach will automatically be selected. If FLCH isselected at or below 27,850 feet pressure altitude, theIAS mode will automatically be selected. The pilot canchange this automatic selection at any time by pushingthe GP-820 IAS/MACH button.

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The pilot can manually select IAS or MACH mode at any time bypushing the GP-820 IAS/MACH changeover button. If the FLCHmode is engaged, selecting IAS or MACH mode also causes thepilot and copilot PFDs speed tapes to change (i.e., MACH modeprovides a math tape and IAS mode provides a CAS tape). EFISinhibits the Mach tape below 25,000 feet but this will notprevent the selection of the FLCH MACH mode below this altitude(in this case the FGCwill hold a Mach value but an equivalentCAS target bug is displayed on the PFD CAS tape).

The FLCH mode will automatically switch from IAS to MACHmodeduring a climb when the actual aircraft speed is at or abovethe FMS initialized climb Mach (the PFDs will alsoautomatically switch from a CAS to a MACH tape). Conversely,the FLCH mode will also automatically switch from MACH to IASmode during a descent when the actual aircraft speed is at orabove the FMS initialized descent CAS (the PFDs will alsoautomatically switch from a MACH to a CAS tape). The pilot canchange this automatic selection at any time by pushing theGP-820 IAS/MACH button.

At mode engagement, with the speed bug on the PFD synchronizedto existing aircraft speed, the pilot has two options to flythe mode:

c Retard the speed bug reference (climb). Advance throttle settings (climb)

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Honeywell !!t!~f.c’

3. B. (8) (d)

In the first instance, with the pilot retarding the speed bugreference, selected IAS/Mach is compared against actualIAS/Mach. The difference is changed to a rate term and thenconverted to an altitude rate term. This signal is routed

through a zero vertical speed limiter, pitch rate limiter, TASgain programmer, and sent to Figure 260, sheet 7 as the FLCH orVFLCH command. This signal determines what is the aircraft’scommanded vertical speed as the system maintains the IAS/Machreference. As the aircraft flies to the commanded verticalspeed, actual altitude rate is summed against the command tocenter the flight director command cue. Actual airspeed orMach is routed through a rate taker and is gain changed to actas a damping term on the command signal.

In the second instance, with the mode engaged and the speed bugsynchronized to existing aircraft speed, the pilot advances thethrottles to maintain the speed reference during the FLCH orVFLCH maneuver. Initially, the aircraft starts to accelerate.The increase in TAS and longitudinal acceleration is changed toa potential speed rate, with normal acceleration added as adamping term. This potential speed rate is changed to analtitude rate signal and the commanded vertical speed signal isprocessed as previously discussed and is routed to Figure 260,sheet 7 and is processed as discussed in paragraph 3.B.(8)(i).

The FMS (and air data) information is used from the same sideFMS and DADC as are selected for display on the in-command sidePFD. If FMS is not selected for display, the DAFCS selects theFMS on the same side as the PFD in-command.

VNAV Path (VPATH) Mode (See Figure 260, sheet 3.)

The VPATH mode is used to descend to a new flight level at aprescribed angle (1 to 6 degrees). Activation of VPATH mode isautomatic from the coupled FMS. The FMS calculates a top ofdescent (TOD) based on the altitude constraints entered intothe FMS flight plan and the desired or calculated path angle.The VNAV path mode is annunciated on the PFD by a green VPATH.

Commands to maintain the aircraft on the prescribed path arebased on the target vertical speed received from the coupledFMS over ASCB. Note that this-target is not displayedGP-820 Flight Guidance Controller. The vertical speedon the GP-820 will be blanked except in vertical speedwhich is discussed in paragraph 3.B.(8)(b).

Prior to mode engagement or use of TCS, altitude rate

on thedisplaymode,

information provided by the AZ-81O DADC through ASCB is fed toa summing point where it is compared to altitude rate. Thesignals cancel so that there is no error signal. Upon modeengagement or release of TCS, the aircraft altitude rate iscompared with the FMS commanded vertical speed. Should adifference between these signals occur, the difference is TAS

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Honeywell ###!b.cEgain programmed and routed as a VPATH command (Figure 260,sheet 7). TAS gain programming accurately adjusts the VPATHcommand signal as a function of the aircraft’s current speedand barometric altitude. The VPATH command signal will drivethe flight director pitch command cue on the PFD in the properdirection to fly the aircraft back to the prescribed path. Asthe aircraft returns to the prescribed path, the VPATH commandwill decrease towards zero. The aircraft has now returned tothe prescribed path. For discussion of the VPATH commandsignal as it is processed (as shown in Figure 260, sheet 7,refer to paragraph 3.B.(8)(i).

3. B. (8) (e) ~~~~;u~e)Hold and VNAV Altitude Hold Modes (See Figure 260,.

Activation of the ALT HLD pushbutton on the GP-820 FlightGuidance Controller selects the altitude hold mode andoverrides all active pitch F/D modes. The altitude hold modeis annunciated on the PFD by a green ALT.

Activation of the VNAV ARM and ALT HOLD pushbuttons on theGP-820 Flight Guidance Controller selects the VNAV altitudehold mode and overrides all active pitch F/D modes. The VNAValtitude hold mode is annunciated on the PFD by a green VALT.

Activation ofALT or VALT modes is automatic after either analtitude preselect capture (ASEL) or VNAV altitude preselectcapture (VASEL).

In the ALT or VALT mode, all armed pitch F/D modes are allowed,but a capture of any armed pitch mode will override the ALT orVALT mode. To allow the initiation or the continuation of themode, the selected DADC must be valid.

ALT and VALT modes are vertical axis flight director modes usedto maintain a barometric altitude reference. The vertical axisof the flight director will maintain the barometric altitude atthe time of mode engagement. The reference altitude may bechanged by using TCS to maneuver to a new altitude and thenreleasing the TCS button. Using the pitch wheel on the PC-880cancels the ALT mode.

Prior to mode engagement, barometric altitude informationprovided by the selected DADC is routed through a summingjunction and a closed ALT HOLD + TCS switch to the input of thealtitude hold reference synchronizer. The synchronizerdevelops an output equal in amplitude but opposite in polarityto its input. The synchronizer output will sum with and cancelthe actual altitude information resulting in the altitudereference signal continually being synchronized to zero priorto mode engagement.

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Honeywell

When the altitude hold


mode is engaged, the ALT HOLD + TCSswitch opens and clamps the synchronizer output as thereference for the mode. With the aircraft on the desiredaltitude, the barometric altitude and integrator signals matcheach other at the summation point, and no error signal results.If the aircraft departs from the desired altitude, thesynchronizer output remains unchanged as a reference. Thechange in altitude is detected by the air data computer andcompared with the reference altitude clamped in thesynchronizer.

The difference in signals will generate a displacement errorsignal to fly the aircraft back to the selected altitude. Theerror signal is summed with washed out pitch attitude. Longterm pitch attitude is washed out, so that the aircraft will

maintain the pilot desired altitude.

The combined altitude reference and pitch attitude washoutsignals are filtered, rate limited and summed withinstantaneous vertical velocity (IVV).

The IVV signal is a control term that helps to stop the

aircraft from departing the desired altitude, and as it comes

back to the desired altitude, the IVV term controls how quickly

the aircraft comes back to reduce overshoot.

The combined altitude command and IVV signals are then TAS gainprogrammed to achieve approximately the same aircraft response,regardless of the aircraft’s airspeed and altitude. Thealtitude command signal is routed to Figure 260, sheet 7 and isprocessed as discussed in paragraph 3.B.(8)(i).

3. B. (8) (f) Altitude Preselect (ASEL) and VNAV Altitude Preselect (VASEL)Modes (Figure 260, sheet 5.)

The ASEL and VASEL modes are automatically armed when theaircraft is flying toward the preselected altitude or analtitude constraint. The preselected altitude is input to theAZ-81O Digital Air Data Computer via the altitude select knobon the GP-820 Flight Guidance Controller. The master DADC thentransmits the preselected altitude over ASCB to the FZ-820Flight Guidance Computer and the SG-884 Symbol Generator fordisplay. The altitude constraint is entered into the FMSflight plan and sent over ASCB to the FGC by the NZ-920Navigation Computer. VSHOLD, FLCH, PITCH HOLD, or VFLCH can beused to fly to the selected altitude. The arm modes areannunciated on the PFD by a white ASEL or VASEL.

When the bracket altitude is reached, the system automaticallyswitches to altitude capture mode. When the mode captures,ASEL or VASEL in green will be displayed on the PFD. The modeannunciation will be boxed for 5 seconds to indicate the

22-14-00Page 298.115


Honeywell

transition from arm to


capture. When in VASEL. the bracketaltitude will be either the preselected altitude or thealtitude constraint, whichever has the smallest altitude error.When the desired altitude is reached, the ASEL VASEL CAP modeis automatically cancelled, and ALT or VALT hold mode isautomatically selected.

The ALT hold mode is annunciated on the PFD by a green ALT.The VALT hold mode is annunciated on the PFD by a green VALT.The ALT or VALT hold mode will be dropped following pitch wheelmovement.

During the three phases (ASEL or VASEL ARM, ASEL or VASEL CAP,ALT or VALT HOLD), a GS capture will override the altitudemode.

If the selected DADC is not valid, none of the three altitudemodes (ASEL or VASEL ARM, ASEL or VASEL CAP, ALT or VALT HOLD)can be initiated or continued.

In the FZ-820 Flight Guidance Computer, actual barometricaltitude is compared against preselected altitude or analtitude constraint. The altitude signals are provided by theAZ-81O Digital Air Data Computer and/or the NZ-920 NavigationComputer through the ASCB. The resultant altitude errorcommands a changing altitude rate from a lookup table in theFZ-820. The lookup table is a predetermined software program.Commanded altitude rate is then compared against actualaltitude rate to determine the altitude capture point. For thecapture detector to trip the following must be true:

● Selected DADC valid.. ASEL or VASEL mode is armed.● The ALT set knob is not being turned (SLEW).● The pitch wheel is not in motion (PW MOTION).

In the capture detector, if commanded altitude rate is greaterthan actual altitude rate, the detector does not trip. Whencommanded altitude rate is less than actual altitude rate, thedetector trips and the aircraft will now initiate a flaremaneuver to capture the desired altitude. The capture pointis a nonlinear function dependent on commanded altitude rate.As an example, with a commanded altitude rate of 24 ft/sec(1440 ft/min), the capture point is 231 feet from the selectedaltitude. With a commanded altitude rate of 100 ft/sec(6000 ft/min), the capture point will be about 3000 feet fromthe selected altitude.

After mode capture, the altitude error signal is run throughthe flare computation circuit and G limiter before being summedwith washed out pitch attitude. The flare computation circuitis designed to gain adjust the altitude error signal to providea constant vertical acceleration capture of the selected

22-14-00Page 298.116


3. B. (8) (g)

Honeywell

altitude. The G limiteracceleration is held to


acts to ensurea maximum of 8

that the rate offt/sec2.

After the G limiter, the altitude error signal is summed withwashed out pitch attitude. Long term pitch attitude is washedout so that the aircraft will maintain the pilot desiredaltitude. From the summing junction, the ASEL or VASEL CMDsignal is filtered, rate limited and summed with IVV(instantaneous vertical velocity).

IVV is a combination of vertical (normal) acceleration andaltitude rate. This signal is used as a damping term and issummed with the ASEL or VASEL CMD signal to enhance thesmoothness of the flare maneuver. The aircraft will remain inthe ASEL or VASEL capture mode until the following conditionsexist simultaneously:

● ASEL or VASEL CAP● ALT error is less than 25 feet

. ALT rate is less than 5 ft/sec

At this time, the ASEL or VASEL mode is dropped and theaircraft is automatically placed in the altitude hold mode.After being summed with IVV, the ASEL or VASEL CMD signal isTAS gain programmed and routed to Figure 260, sheet 7, and isprocessed as discussed in paragraph 3.B.(8)(i).

Glideslope (APP) Mode (See Figure 260, sheet 6.) The glideslopemode is used for the automatic intercept, capture and trackingof the glideslope beam. The beam is used to guide the aircraftdown to the runway in a linear descent. Typical glideslopebeam angles vary between 2 and 3 degrees, dependent on localterrain. When the glideslope mode is used as the verticalportion of the localizer approach mode, it allows the pilot tofly a fully coupled ILS approach. The mode is interlocked, sothat glideslope capture is inhibited until localizer capturehas occurred as previously discussed in paragraph 3.B.(7)(c).

The glideslope mode is set up and flown as follows:

. The navigation receiver is tuned to the published ILSfrequency for the runway in use.

● The course pointer and heading bug are set on the PFD forlocalizer intercept.

● The APR mode is selected on the GP-820 Flight GuidanceController.

● The NAV/ILS navigation source is selected on the DC-884Display Controller.

22-14-00Page 298.117


With the local izer captured, and outside the normal glideslopecapture limits, the PFD will annunciate the following modes atthis time:

● LOC in green.● GS in white.

Any other vertical mode in use at this time will also bedisplayed.

As the aircraft approaches the glideslope beam, the verticalbeam sensor (VBS) monitors TAS, vertical speed, and glideslopedeviation in determining the correct capture point. Atglideslope capture, the computer drops any other vertical modethat was in use, and automatically generates a pitch command tosmoothly track the glideslope beam.

At this time, the PFDwill annunciate:

. LOC in green

. A green GS that is boxed for 5 seconds to emphasize that thecapture phase has occurred.

The glideslope deviation signal is routed to the symbolgenerator from the navigation receiver on the ARINC 429 bus.From the symbol generator, the signal is routed to the FZ-820Flight Guidance Computer through the ASCB.

Gain programming is performed on the glideslope signal tocompensate for the aircraft closing on the glideslopetransmitter, and beam convergence caused by the directionalproperties of the glideslope antenna. Glideslope programmingis normally accomplished as a function of radio altitude andvertical speed. The radio altitude signal is rate limited,summed with vertical speed, and limited again before gainprogramming the glideslope signal. If the radio altimeter isnot valid, then GS gain programming is accomplished as afunction of preset height above runway estimates and run downas a function of true airspeed. From the GS gain programmingblock, the glideslope signal is filtered, rate limited andsummed with estimated vertical deviation rate.

Estimated vertical deviation rate is used as a damping term tohelp maintain a truer track of the glideslope beam. Theestimator utilizes normal acceleration provided from the IRS,along with glideslope deviation, to provide an inertiallyderived vertical rate, with long term glideslope deviationcorrection.

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Honeywell #&!!r.cEThe summation of glideslope deviation and vertical deviationrate is then TAS gain programmed. TAS gain programming allowsfor better glideslope tracking qualities over a given range ofapproach speeds. After being TAS gain programmed, theglideslope command signal is routed to Figure 260, sheet 7 andis processed as discussed in paragraph 3.B.(8)(i).

3. B. (8) (h) Dual Couple Approach Mode

During the tracking phase of an ILS approach, the system willutilize landing aid flightpath information from both the pilotand copilot PFD. Initiation of this flight segment of theapproach phase is automatic and occurs at 1200 feet radioaltitude as shown in Figure 259.1.

The dual couple mode requires all the following conditions tobe satisfied:

. Pilot PFD NAV source must be NAV 1 or ILS 1.

● Copilot PFD NAV source must be NAV 2 or ILS 2.

● LOC/GS track mode,

. Radio altitude less than 1200 feet.

. Both NAV 1 and NAV 2 sources are valid.

/

RADIOALTITUDEBETWEEN1200

/

AND1500FEETDUALCPLMODE

I (1200RA)

I I

‘-1 ~p%4Ri%:fiD”FEETII

*I

/v v

OUTER MIODLE / /

MARKER MARKERRUNWAY

4.O-7.OMILES 3500’

Dual Couple ApproachFigure 259-.1 22-14-00

Page 298Tli9Aug 15/91


Honeywell !!#&!!!.cEAt

●

●

In

dual couple transition the following events occur:

Command bars are in-view on both pilot and copilot PFDs.

The GP-820 PFD-CMD button illuminates both L and Rannunciators.

the dual couple mode the high priority FGC (i.e., the FGCselected on the-DC-884 sensor page) averages the radio datafrom the pilot and copilot NAV sources. This allows theapproach mode to be continued in the event of a failure of onenavigation receiver or if there is an unflagged miscomparebetween the NAV receivers. The high priority FGC outputsidentical FD commands (based on the average radio deviations)to both pilot and copilot PFDs in the dual couple mode.

Navigation failures in the dual couple mode result in thefollowing automatic reconfiguration:

. NAV 1 Fails - Pilot PFD command bars are removed and thePFD-CMD reverts to R. Approach mode continues using datafrom NAV 2. If NAV 1 becomes valid the dual couple modewill automatically return.

. NAV 2 Fails - Copilot PFD command bars are removed and thePFD-CMD reverts to L. Approach mode continues using datafrom NAV 1. If NAV 2 becomes valid the dual CPL mode willautomatically return.

. Unflagged Miscompare Between NAV 1 and NAV 2 - The systemwill automatically select the NAV source with the smallestdeviation (must be greater than 1/3 DOT). The PFD-CMD logicand command bars operate the same as in the NAV failurecase. This is annunciated on the EICAS.

NOTES: 1. Although the FGC uses averaged ILS data forguidance, the EFIS comparison monitor alwaysannunciates LOC/GS miscompares based on nonaverageddata (i.e., direct comparison of NAV 1 and NAV 2).

2. The PFD-CMD pushbutton is inhibited in dual couplemode.

3. After cancellation of the APR mode the PFD-CMDstatus reverts to the side which was selected priorto the dual couple transition.

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3. B. (8) (i) Pitch Autopilot Mode Flow (See Figure

The Ditch autor.)ilot diaqram shows two

260, sheet 7.)

signal paths for thevert{cal command. The ?irst path is with the-autopilotdisengaged and routes the vertical command to the symbolgenerator for the PFD only. This path is discussed inparagraph 3.B.(8)(i)l. The second path is with the autopilotengaged, and routes the vertical command to both the symbolgenerators for the PFD and to the elevator servo drive motor.This path is discussed in paragraph 3.B.(8)(i)~.

When the autopilot is engaged and no vertical flight directormode is selected, the system will automatically revert to thebasic autopilot mode of pitch attitude hold. This is discussedin paragraph 3.B.(8)(i)~. The go-around mode of operation isdiscussed in paragraph 3.B.(8)(i)~.

~ Vertical

With thevertical

● Pitch. Pitch

Steering Command With Autopilot Disengaged

autopilot disengaged, the selected flight directorcommand is routed through the following:

Rate LimiterLimiter

. Go-Around Switch

. AP Engage Switch● Pitch Bar Bias Switch

The command is then routed to the symbol generators on theASCB.

Pitch bar bias (PBB) is a fixed voltage level used to biasthe vertical flight director command bar out of view toprevent the pilot from flying invalid data. The verticalcommand bar is biased out of view if:

● No flight director mode is selected.. The flight director is not valid.

Flight director valid is comprised of the following:

● “A” CPU heartbeat monitor valid which provides validitylogic to engage/disengage the autopilot functions. Theheartbeat monitor is hardware independent from the CPU,so that no single fault can disable both the CPU and themonitor. The heartbeat monitor output is provided as aninterrupt to the nonmonitored processor.

● Flight director flag/annunciatorvalid which is a directdiscrete output from the “A” CPU to drive the FD flag onthe EADI and enable the annunciator drivers in the sameside guidance controller channel.

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Honeywell !!!!!!%!!.”● FGC power supply valid which monitors the internal power

supply voltages for proper operating levels.

3. B. (8) (i) ~ Vertical Steering Command With Autopilot Engaged

With the autopilot engaged, the selected flight directorvertical command is routed through the:

. Pitch Rate Limiter● Pitch Limiter● Acceleration Limiter. Another Pitch Rate Limiter

The resultant signal is then summed with pitch rate. Pitchrate is supplied by the IRS and is used as a damping term tohelp control the speed at which the pitch maneuver willoccur. The vertical command is then pitch limited again andthen follows two paths.

The first path goes up through the activated AP ENG switch,TCS switch, pitch bar bias switch, and then to the symbolgenerator.

The second path routes the vertical command to a summingpoint through a t8-degree position command limiter. Thesignal is then gain adjusted and applied to a pulse widthcommand limiter and output servo amplifier.

The pulse width command limiter serves two functions.First, it is the D/A converter for the servo amplifier.Second, as a motor driver it issues a continuous string of28 V dc pulses, at a rate of one pulse approximately everytwo milliseconds. The pulse width of the pulses isdetermined by the vertical command. The pulse-width commandlimiter has its current limits established by a softwareprogram contained in the “B” CPU.

The output of the servo amplifier is sent to the SM-600elevator servo drive motor, the “A” processor currentmonitor to check for servo runaway current, and to the pitchtrim threshold sensor. The SM-600 is a permanent magnet dcmotor that utilizes a dc tachometer for rate feedback. Ithas but does not use a position feedback synchro. As theservo motor drives to position the elevator, it also drivesthe dc tach generator through mechanical coupling(represented by a dotted line). The tach generator providesa rate feedback signal that serves two functions. First, itacts as a damping term when summed with the vertical connnandinput to the pulse-width command limiter. This helps tostabilize elevator position and minimize excessive elevatortravel.

22-14-00Page 298.122


Second, the rate feedback signal is integrated to obtainposition feedback, gain adjusted and summed with thesteering command. When these signals are equal, theelevator is in the proper position to satisfy the verticalcommand. As the aircraft responds, the flight directorcommand diminishes and the position feedback signal drivesthe elevator servo back to its original position.

Should there be a mismatch between vertical command andelevator servo position, a flightpath standoff could occur.To prevent this standoff, any command at the output of thesecond pitch limiter is routed through a command rate takerand limiter.

The signal is changed to rate, and summed with tachgenerator rate feedback. The summing of these two signalsis then integrated to obtain position data and summed withthe vertical command.

3. B. (8) (i) ~ Autopilot Pitch Attitude Hold (See Figure 260, sheet 1.)

Pitch attitude hold is the basic vertical autopilot mode.It is automatically active if the autopilot is engaged andno vertical flight director mode has been selected. Priorto autopilot engagement, pitch attitude is routed through asumming point-d through the normally closed PITCH HOLD +(AP ENGAGE ● TCS ● NO VERTICAL FD MODE SELECTED) switch to asynchronizer. The output of the synchronizer is invertedand summed with pitch attitude to give a zero output fromthe summing point. If the autopilot is engaged and novertical flight director mode is selected, the synchronizerswitch opens, clamping the synchronizer with the referencepitch attitude at the time of autopilot engagement. Shouldthe aircraft deviate from the desired reference attitude,the difference from the summing point is limited and routedto Figure 260, sheet 7, as a pitch hold command signal andis processed as discussed in paragraph 3.B.(8)(i)~. Thereference pitch attitude can be changed through the use ofthe pitch wheel on the PC-880 Turn Pitch Controller, withthe autopilot engaged. Moving the pitch wheel will cause arate generator output. The direction of pitch wheel motionwill determine the polarity of the output, while the speedof pitch wheel motion will determine the amplitude of thesignal. The pitch wheel signal is then TAS gain programmedto more accurately adjust the signal as a function ofaircraft speed and barometric altitude. The pitch wheelsignal then changes the pitch attitude reference, and theoutput from the summing point is identical to what waspreviously discussed.

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3. B. (8) (i) 3 Go-Around (Pitch Axis) Mode

The go-around mode is normally used to transition from anILS approach to a climb out condition when a missed approachhas occurred. The pilot selects go around by pressing theTOGA button located on either outboard throttle handle.With go-around selected, all armed and active flightdirector modes are cancelled and the autopilot isdisengaged. The GA switch changes state, and a fixed15-degree GA up bias is routed to the pitch command cue onthe PFD. The 15-degree angle is determined by the aircraftmanufacturer and represents the best climb angle. As thepilot flys the GA command, pitch attitude information fromthe IRS nulls the command and centers the pitch command cueon the PFD. (A 12-degree GA up bias is used in the-906/-907 FGC).

The go-around mode is cancelled by selecting another pitchmode, engaging TCS, or engaging the autopilot. Thego-around mode is annunciated on the PFD by a green GA.

(j) Autopilot Pitch Trim (See Figure 260, sheet 8)

There are three modes of electric trim operation available:

● Manual electric trim. Mach trim. Autopilot trim

The trim function will control trim tab motion via digitallyprocessed closed-loop control laws which utilize motor rate(tachometer) and motor position (synchro) feedbacks. Each FGCprovides excitation for its own-side trim synchro.

Trim servo travel is limited such that the trim servo does notdrive into the mechanical stops by use of the trim limitswitches in the trim servo bracket. The manual trim wheel canbe used to override the electric trim via a slip clutch (foremergency use only). Trim override inputs via the slip clutchwill cause a change in the relative positions of the trim limitswitches. Therefore, readjustment of the servo linkage may berequired to realign the trim limit switches. When a trim limitis reached, the trim servo will not continue to drive in thecurrent direction, however, the trim servo can be driven in theopposite direction.

Manual Mode - The manual electric trim mode is activewhenever the electric trim is engaged with the autopilotdisengaged. This mode responds to trim inputs from the

pilot and copilot yoke mounted trim switches. The maximumtrim rate shall be t2° tab/see in this mode at lowairspeeds. This speed limit shall decrease with trueairspeed increasing.

22-14-00Page 298:124


Each flight guidance computer receives manual trim inputsdirectly from the pilot’s and copilot’s manual electric trimswitches (located on control yokes). Since these are splitactuated trim switches (each is dual) the flight guidancecomputer brings in a total of four analog discretes asfollows:

. Pilot’s trim up (switch 1)

. Pilot’s trim down (switch 1)● Copilot’s trim up (switch 1)● Copilot’s trim down (switch 1)

Each channel of the GP-820 Flight Guidance Controller shallreceive four analog input discretes as follows:

. Pilot’s trim up (switch 2)

. Pilot’s trim down (switch 2)● Copilot’s trim up (switch 2)● Copilot’s trim down (switch 2)

These inputs received by the flight guidance controller aretransmitted via private-line serial data link to the flightguidance computer.

Any one of the above pairs, if active together (and’edlogic), activate the manual electric trim in the directionselected.

If any two and’ed switch pairs are on at once calling fortrim in opposite directions, the electric trim will notdrive. The DAFCS monitors the pilot’s and copilot’s trimswitch pairs for disagreements and provides individual trimswitch caution flags via ASCB to the EDS system for crewalerting.

The electric trim rate is gain programmed (as a function oftrue airspeed). This data shall be received via ASCB fromthe DADCS. This trim rate programming is a factor when theautopilot is engaged (autotrim) and when the manual electrictrim is activated via the pilot or copilot trim switches.This rate programming is not to be used for Mach trim ratecontrol. When voted DADC data is not valid, the electrictrim defaults to the high trim rate.

3. B. (8) (j) ~ Mach Trim Mode - The Mach trim mode is active above 0.7 Machwhenever the electric trim is engaged with the autopilotdisengaged and the manual mode inactive (i.e., no pilot triminputs). The Mach trim mode trims the aircraft noseup forincreasing Mach number and trims the aircraft nosedown fordecreasing Mach number. This control only applies in thespeed range between 0.838 Mach and 0.95 Mach and results in

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a total linear trim tab displacement of 6 degrees whenpassing through the region. Outside this speed region theMach trim function holds trim tab position. Transition toMach trim mode ON (via autopilot disengage, trim engage, ormanual trim switches released) causes the Mach trim controllaw to synch such that the present trim tab position is thecorrect starting point for the present Mach situation. Trimrates are limited to tl.O degree tab/see in Mach trim mode.

At trim engagement, the servo control is synched to commandzero rate, and the Mach trim is synched to hold present trimposition.

At trim disengagement, no synching is required.

At channel switchover, the trim is synched thetrim engagement, as described above.

When the pilot or copilot manual trim switchesthe Mach trim position reference is synched toposition and the Mach trim control law remainsmanual trim switch is released. The Mach trimcontinues to synch for a short time afterward -allow the trim motor to coast to a stop. This

same as

are actpresentoff untcontroln order

Drevent

at

ve,

1 the1awtothe

Mach trim from causing the trim to back up following amanual trim input.

3. B. (8) (j) ~ Autopilot Trim - The elevator servo current is monitored bythe FZ-820 FGC. When the current goes above a predeterminedthreshold, the trim tab is moved in the opposite directionfrom the elevator. This reduces the load on the elevatorwhich in turn reduces the amount of current required todrive the elevator servo. A delay is required to preventtrim from driving during transitory loads.

22-14-00Page 298.126


C1OJ2B— 10

——— ———=4=mGHT GUIDANCS COMPUTER

AVERAGED

—— ——— ——— ——— ——. ——— ——— ——— ——— ———

( PITCH ATT

~

/1

1

2

0

14

Is

M

)5

1

2

F,m CROSS

r

——— ——— SIOE

PC4B0 TURN PITCH 1~ FGC

CONTROLLER129J1

PITCH WHEEL

(

+ 18

TACH NO 1 - 19

TAS GAINTO SH7

PROGRAMMER

t

~

PITCH HOLD +(AP ENGo~●NO VERT F D MODE)

II NOTE 4

I-{ PITCH ATTITUOE ~ TO SH1 AND SH7

\ /

TO SH7

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Figure 260 (Sheet 1) “ 22-14-00Page 298.127/298~128

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Aug 15/91Use or disclosure Of lflfOrm8t10n on this page IS subject to Ihe restrictions on the title page 01 this documenl.

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Figure 260 (Sheet 4) - 22-14-00Page 298. 133/298; 134

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Figure 260 (Sheet 5) - 22-14-00Page 298.135/298;136

Aug 15/91Use or disclosure 01 mtormallon on this pege is subject 10 the restrtcllons on Ihe hlle page of this document

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F1ight Director/AutopilotPitch Channel Mode Flow Diagram



3. B. (9) Yaw Channel Functional Operation (See Figure 261, sheets 1 and 2.)

The yaw axis of the autopilot provides directional stability (yawdamping) and directional control for turn coordination. The yawaxis of the autopilot receives sensor information from the IRS andthe DADC. The IRS supplies the following information through theASCB.

w Yaw rate. Roll rate. Pitch attitude. Roll attitude. Normal acceleration● Longitudinal accelerations Lateral acceleration

The DADC supplies the following information through the ASCB.

. Indicated airspeed (IAS)● True airspeed (TAS). Altitude rate (VS)

The above inputs from the IRS and the DADC are all combined in theFZ-820’S rudder command processor. (See Figure 261, sheet 2.) Therudder command processor will determine the proper rudder deflectionto maintain directional stability and control.

Yaw rate, true airspeed, roll attitude, and lateral acceleration arethe primary controlling inputs for the yaw axis. The rudder commandprocessor looks at yaw rate and computes the control responsenecessary to bring the yaw rate of the aircraft to zero. Trueairspeed, roll attitude, and lateral acceleration combine to provideturn coordination. The remaining inputs to the processor aresecondary and aid in further optimizing yaw control.

The yaw axis may be engaged by pressing the YD pushbutton on theaircraft control panel. The yaw axis is disengaged by also pressingthe same YD button. The yaw axis will automatically be engaged whenthe autopilot is engaged. The autopilot cannot be operated withoutthe yaw damper.

Upon engagement, if the rudder command processor is satisfied, itsoutput will be zero and the rudder will remain centered. When therudder command processor detects a need for yaw correction, it willroute its command signal through an easy on circuit. This easy onfunction allows the command processor signal to gradually be appliedto the rudder when the yaw damper is engaged. In the event that theprocessor were to command a large displacement at YD engagement, thepilot would have sufficient time to react and disengage the yawdamper. The yaw rate command is then adjusted to the proper gainand rate limited. The command is now changed from digital to analog

22-14-00Page 298.143

Aug 15/91Use or disclosure Of information On this page IS subject to the restrictions on the title page of this document.

form and continues on through a pulse width command limiter andservo amplifier. The pulse width command limiter is a motor driverthat issues a continuous string of 28 V dc pulses at a rate of onepulse approximately every 2 milliseconds. The pulse amplitude andpolarity are determined by the yaw command input.

The servo amplifier drives the motor of the rudder linear actuator,which in turn drives the ballscrew. The ballscrew is mechanicallycoupled to:

● LVDT (Linear Variable Differential Transformer)● Hydraulic Package

The hydraulic package drives the aircraft’s rudder control surface.As the rudder control surface responds to the command, the LVDTprovides a position feedback signal to the FZ-820.

The servo amplifier output is also routed to the “A” processorcurrent monitor to check for servo runaway current. As the actuatordrives the ball screw, the LVDT is moved and provides an output thatequates to commanded rudder position. This signal is routed to theFZ-820 where it is demodulated and then follows two paths.

The first path is through a rate taker, where the position signal isconverted to rate of position change. This signal is utilized as adamping term to minimize excessive rudder travel.

The second path from the demodulator retains position informationand is summed with the command signal. When these two signals areequal and opposite, the command has been satisfied and the rudderstops moving.

As the aircraft responds to rudder clef”movement through aerodynamic feedback,rudder command processor diminish. Thpredominate and drives the rudder back

ection, the IRS detects thisand the input signals to thes allows the LVDT signal toto its original position.

22-14-00Page 298.144

Apr 15/93IJse or disclosure of informationon this page is subject to the restrictions on the title page of this document.

10J1B;1OJ2B

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Flight Director/AutopilotYaw Channel Mode Flow Diagram


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DATA FROM ASCB’’A’’ AND ‘0”1SAVERAGED TOGETHER TO PROVIOE

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THE SWITCH NOMENCLATURE

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NOTE 1

-— ——— — . — — — . — — — —— —— — — —— —AO.13665 I .R2

22-14-00Page 298.145/298.146


——— ——— ——— ——— ——— ——— ——— ——— —.— ——— —FZdSO FL2GNT GUIDANCE COMPUTER

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F1ight Director/AutopilotYaw Channel Mode Flow Diagram

Figure 261 (Sheet 2) 22-14-00Page 298. 147/298. 148

Apr 15/93


4. FMZ-800 Fliqht Management System

A. General

The flight management system (FMS) is designed as a federated system,where there are independent components but each component performs asystem function. The FMS is comprised of three basic components. Thefirst component of the system is the CD-81O Control Display Unit (CDU).The other two components are the NZ-920 Navigation Computer and thePZ-800 Performance Computer. Since each computer is a self-containedunit, NAV only or performance only systems are possible. The key to thefederated approach is to keep the navigation and performance functions asseparate and independent as possible, and yet allow for theircoexistence.

Some performance calculations, however, utilize navigation data and somenavigation displays include performance data. The flight plan data andother joint use information are shared between the navigation andperformance computers on the ASCB; however, the majority of the NAV-PERFintegration is through the CDU. The CDU can be driven from thenavigation computer for NAV, or predominately NAV display pages. The CDUis then driven from the performance computer for PERF pages. This keepsthe computers as independent as possible. It is a design requirement tokeep aircraft variable data in the performance computer. Thearchitecture of the NAV, performance, and CDU integration is shown inFigure 262.

The CDU provides the primary means for pilot input into the system. Italso provides an important output display for the navigation andperformance computers. The CDU utilizes a full alphanumeric keyboard,with four line selection keys on either side of the CRT. Severalfunction keys are provided to allow direct access to specific displaypages. Annunciators are built into the top of the unit to advise thepilot of the system’s status. The CDU sends ASCII characterssimultaneously to both the navigation and performance computers.

The NZ-920 Navigation Computer is the component in the FMS, whichprovides both lateral and vertical navigation guidance. The databaseinside the NZ-920 is used for storage of waypoints, navaids, routes,airports, and other NAV data for easy access by the pilot. The NZ-920can interface with five long range sensors; three via ARINC 429 buses andtwo over the ASCB bus. Each navigation computer can also connect to dualProline 2 429 scanning DME receivers and VOR/ILS receivers. Theinterface to the air data, EFIS, performance computer and autopilot isover the avionics standard communications bus(ASCB). Flight Plans arealso transferred between navigation computers over the ASCB while thelink to the CDU is over a RS-422 private-line interface. To providehigh-accuracy long range navigation, the navigation computer is designedto connect to IRS, GPS, Omega/VLF sensors plus VOR/DME. With links tothe on-board navigation sensors, the navigation computer develops an FMSposition based on a blend or mix of the sensors. The FMS does not

22-14-00Page 298.149


direct-source

y display navigation maps on the CDU; however, the FMS is theof map data for other cockpit displays such as EFIS. Display of

maD data is achieved bv the utilization of the internal database and ASCB1/0. A large portion ~fthe navigation database is subject to updatingon a 28-day interval. The DL-800/900 data loader is used for thispurpose.

The navigation part of the FMS may be considered an area navigationsystem or RNAV. Its fundamental purpose is to provide navigationinformation relative to a selected geographically located point.Navigation management will allow the pilot to define a route from theaircraft present position to any point in the world. The system willoutput advisory information and steering signals to allow the pilot orautomatic flight control system (AFCS) to steer the aircraft along thedesired route. Routes are defined from the aircraft present position toa destination waypoint via a direct great circle route or via a series ofgreat circle legs connected by intermediate waypoints.

The purpose of the PZ-800 computer is to provide the pilot with detailedperformance information and automatic control of the throttles.Previously, detailed performance information was only available byspending a great deal of time in the aircraft performance manuals. Withthe performance functions of the PZ-800, the pilot now has access to acomputerized performance manual. The pilot is supplied withflight-planning information prior to takeoff such as fuel and timerequired. In flight, the system provides the pilot with real timeinformation based on current aircraft and atmospheric conditions. Incase of changes to the flight plan, the pilot can be updated immediatelyon the consequences of the changes. This capability extends the pilotsability to conduct the flight safely and economically.

The performance system is enhanced by the autothrottle. Aircraftparameters calculated by the performance functions can be coupled to theautothrottle and flight guidance computer for effortless tracking ofthese commands. Alternatively, the autothrottles can be set to maintainpilot-entered values. In addition to the PZ-800 computer, theautothrottle function requires one SM-81O Autothrottle Servo-Clutchassembly for each engine controlled. Inputs from the throttles, engines,autopilot control panel, and display controller are also required.

22-14-00Page 298.150

Aug 15/91Uss or disclosure of information on this page is subject to the restrictions on the title page of this document.

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22-14-00Page 298.151

Apr 15/93


I 4. B. NZ-920 Navigation Computer (Figure 263)

(1) Avionics Standard Communication Bus (ASCB) Interface

The navigation computer transmits flight plan data, display data,guidance commands, dual NZ-920 sync data, and navigation and customdatabases during normal operation of the bus.

The navigation computer receives dynamic information about theaircraft’s present state via the ASCB. This information consists ofattitude, headinq, velocities, accelerations, altitude and soeeds.Also, dual NZ-926’sync data is received from-the ASCB. ‘

(a) NZ-920 Navigation Computer ASCB Inputs

~ FZ-820 Flight Guidance Computer (FGC) ASCB Inputs

The NZ-920 inputs both left and right FGC data. Itprocesses and uses data from the priority FGC only. -priority FGC is established by reading WSP 1, bit 11,mode/engage priority, of both FGC’S transmissions. Aone setting of this bit identifies the priority FGC.priority bits for both FGCS are set or neither setsimultaneously, the data from each FGC is considered .and is not processed for use by the NZ-920. Only one

he

logicIf the

rivalidFGC

can be priority at a time in order for the NZ-920 to use itsdata.

~ AZ-81O Digital Air Data Computer (DADC) ASCB Inputs

The NZ-920 inputs both left- and right-side DADC data fromthe ASCB. Selection of DADC data is performed by thenavigation and guidance (LNAV and VNAV) subsystems of theNZ-920.

The navigation subsystem selects and uses its on-side DADCdata if valid. Otherwise, it will switch to the off-sideDADC and use that data if the on-side DADC is invalid. Ifboth DADCS are invalid, the NZ-920 continues to searchbetween the two DADCS until a valid sensor is found.

If both DADCS are valid, the navigation subsystem in eachNZ-920 of a dual system configuration have independentsources of air data for input into the position computation.

The guidance subsystem selects and uses the same DADC fromwhich the high priority autopilot gets its data for use inits control laws for steering the airplane. By using thisDADC, guidance is using the same altitudes and speeds as theautopilot, except during the approach mode. During anapproach, the autopilot will average the DADCS and guidancewill use the previously AFCS selected DADC state todetermine its DADC usage.

22-14-00Page 298.152


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-1)

}

ARINC 429 RCVRMLSilLSPRIMARY

L)‘1.-

RS422 XMTR

{

(H)

CDU CLK (L)

CDU VALID o I———— . J

~WMLS SELECT

‘DL;O~; — —‘~ DATA LOADER

—11

RS422 FICVR

{

(H)

‘ATA ‘OADER (L)

RS422 XMTR

{

(H)

DATA LOADER (L)““’””m+ * 32NAV NO. 1

+ * 33

RS422 XMTR

{

(H)

CLOCK (L)

}‘) ARINC 429 RCVR

NAV PRIMARYL)

DL CONNECTEU

=}-L ———. — J

ASCE INTERCONNECTSHOWN IN SECTION t

—> FIGURE 6

ASCB

{

(H)

PRIMARY BUS (L)

DAUNO. 1

4*DAUNO. 2

4bDCNO. 1

4

4 )IRUNO. 1

4 ~IRU NO.2

J)

IAZNO, 1

4 ~AZ NO.2

IPZ NO.2

m

* * 43

NAV NO. 2+ * u

H)

}

ARIM2 429 RCVR

L)NAV SECONDARY ASCO

{

(H)

SECONDARY BUS (L)

la

1s ISG NO. 2

H)

}

ARINC 429 RCVR

L)DME PRIMARY

I}

‘) ARINC 429 RCVRDME SECONOARY

L)DME NO. 2

I I SG NO. 31=

I ““l G}

‘) ARltW 429 RCVRLTS NO. 1

L)

I FC NO. 2t+ I b ‘z

VLF/OMEGA+ * 24

H)

}

ARINC 429 RCVR

;L)LTS NO.2

I FZ NO. 1 1.H)

}

ARINC 429 RCVR

[L)LTS NO.3

111.Q31=

PZ NO. 1c121JiB.54

J--r-C121J1AB

1 I

wok! .—— — —NZ-920 NAVIGATfON 1COMPUTER NO. 2TAG SYNC IN ICDU SYNC IN

TAG SYNC OUT ICDU SYNC MT I

ONSIDE TUNING CNTL

TAG SYNC OUT

CDU SYNC OUT

TAG SYNC IN

CDU SYNC IN

VERSION B ASCB

PZ INSTALLED

OP MODE IDo

OP MODE ID1

34

48

9

16

28

31

X.SIDE TuNING CNTL

NO CLOCK ASCB

RADIO CONFIG IDO

RADIO CONFIG 101

RADIO CONFIG ID2

DME SCAN NPE

RADIO BUS TYPE

I I

‘{=1 [H)

}

9

ASCB SECONDARY BUS(L)

I[H)

}ASCB PRIMARY BUS

(L) I

+

SDI NO, 1

SDI NO. 2

.—— —— ———— —.

.—— — — JAD.31610-Rl#

NZ-920 Navigation ComputerInterface Diagram

Figure 263

+

22-14-00Page 298.153/298.154


II

II

II

4. B. (1) (a) ?

I

I

~

I

Selection of DADC data by the guidance subsystem is treateddifferently between the priority (master) NZ-920 and thenonpriority (slave) NZ-920.

The priority NZ-920 guidance selects and uses the DADC thatis selected by the autopilot as previously described.

The priority (master) NZ-920 guidance subsystem determinesthe AFCS selected DADC status by reading the high priorityflight guidance computer (FGC) data.

If both FGCS are high priority or both low priority then theFGCS are considered failed and guidance defaults to the leftEFIS and uses the selected DADC. If the left EFIS is notvalid then guidance uses its on-side DADC if valid. If theon-side is not valid then it switches to the off-side to getDADC data.

If the selected DADC used by the FGC fails or the selectedEFIS DADC fails, then guidance continues to follow theselected DADC and guidance is set invalid.

The nonpriority NZ-920 (slave] guidance subsystem selectsand uses the DADC selected on the EFIS on which thenonpriority NZ-920 is displayed. Guidance continues toselect the EFIS selected DADC regardless of DADC validity.If the selected DADC is invalid, guidance is set invalid.

If the nonpriority NZ-920 is not displayed, it defaults toits on-side DADC if valid. If default DADC is invalid, theNZ-920 switches and uses the other remaining DADC if valid.If both DADCS are invalid, guidance is set invalid.

The nonpriority NZ-920 guidance subsystem reads the EFIS itis displayed on to determine DADC selection. If the NZ-920cannot determine the selection because EFIS is invalid, thenonpriority NZ-920 defaults to its on-side DADC if valid.If the on-side DADC is not valid, the off-side DADC is used.

PZ-800 Performance Computer ASCB Inputs

The NZ-920 inputs and uses only its on-side performancecomputer data. Both basic and background data is used. TheNZ-920 does not read the offside performance computer data.

Inertial Reference System (IRS) ASCB Inputs

The NZ-920 inputs both left and right IRS data. Itprocesses and uses data from both IRSS when available andvalid.

22-14-00Page 298.155


4. B. (1) (a) S

I

I

~

SG-884 Symbol Generator (SG) ASCB Inputs

The NZ-920 inputs left, right, and center SG data. Itprocesses and uses data from all three SGS when availableand valid.

The NZ-920 determines SG reversionary or backup modes (WSP1) from the SG and processes the appropriate SG data basedon this reversion data.

DA-884 Data Acquisition Unit (DAU) ASCB Inputs

The NZ-920 inputs all four DAU transmissions (1A, lB, 2A,2B) . It processes and uses the DAU data as selected by itson-side display controller. If the selected DAU is invalid,the NZ-920 switches and uses data from its correspondingsecondary channel (i.e., 1A or lB, 2A or 2B). If theselected DAU channel becomes valid, the NZ-920 switches anduses the selected channel.

DC-884 Display Controller (DC) ASCB Inputs

The NZ-920 inputs and processes on-side display controllerdata.

FC-880 Fault Warning Computer (FWC) ASCB Inputs

The NZ-920 inputs both left and right FWC data.

(b) NZ-920 Navigation Computer ASCB Outputs

The NZ-920 transmits two types of data onto the ASCB; basicdata and background data.

(2) RS-422 Synchronous Interface

This RS-422 interface is a synchronous serial digital data bus.Data is transmitted over a shielded twisted pair cable. Theimplementation of the RS-422 in the navigation computer has thefollowing parameters:

. SDLC● CRC ITT. No parity● 312.5k baud

All data is transmitted in a byte format using eleven-bit groups.Each group consists of a start bit, followed by 8 data bits,followed by a parity bit (odd), and ending with a stop bit. Thedata is transmitted least significant bit first on the bus.

22-14-00Page 298.156


The RS-422 bus is used to communicate operator-entered data, CDUmessages, and CDU display data between the NAV computer and the CDU,and between the performance computer and the CDU. Also, an RS-422interface communicates Jeppesen data and flight plans between theNAV computer and the database loader.

4. B. (3) ARINC 429 Interface

ARINC 429 is a serial digital data bus standard used betweenavionics system elements. Data transmissions are open loop, i.e.,sinks are not required to inform sources that information has beenreceived. Data is transmitted on a continuous basis from the sourcesystems at rates sufficiently high to ensure small incremental valuechanges between updates. The basic information element transmittedis a digital word containing 32 bits. This data is sent in eitherbinary, binary coded decimal (BCD) or discrete formats. The leastsignificant bit and least significant character of each word aretransmitted first. Also, the least significant bit of the word isthe most significant bit of the label and the label is transmittedahead of the data in each case. The label is used to identify thedata within the word. The bus transmits at either a 100-kHz (highspeed) rate or a 12.5-kHz (low speed) data rate.

The 429 serial bus is used to input the following data to thenavigation computer:

. VOR/ILS Radio Data

. DME Radio Data● IRS Data. OMEGA Data. GPS Datac AFIS Commands and Data

The navigation computer will output sensor initialization data,radio tuning commands, and various other required parameters to thesubsystems via two 429 buses (general bus No. 1 and No. 2). Thisdata is described in Table 206.

22-14-00Page 298.157


Honeywell !$ji$!&.

M9!ld

Remote VHF Tune

Remote ATC Tune

Remote ADF Tune

Remote VOR/ILS Tune

Remote DME Tune

Set Latitude

Set Longitude

Set Magnetic Heading

Omega Station Deselect

Date

GMT

True Airspeed

POS Latitude

POS Longitude

Groundspeed

True Heading

Uind Speed

Wind Oirection

Time to Destination

Distance to Dest

LNAV Status Uord

Lateral Oeviation

Baro Corrected Alt

Equipnent ID

Octal~

030

031

032

034

035

041

042

043

272

260

125

210

310

311

312

314

315

316

352

351

275

116

204

371

Format

BCO

BCD

BCD

BCD

BCO

BCO

BCD

BCD

01s

BCDBCOBNRBNRBNRBNRBNRBNR

BNR

BNR

BNR

DIS

BNRBNR01S

Signif Digital Approx PositiveChar** -&!19!L Resolution ~ Sense

5 118.0-135.975 --- MHz ---

(Referto Chapter3 of ARINC 429 Specification)

5 190-1750

5 108.0-117.95

5 10B.O-135.95

6 180N-180S

6 180E-180W

3 0-359

(Refer to Table 207)


5 0-23 .59.9

15 0-2048

20 *0.5

20 *1. O

15 0-4096

15 *1. O

6 0-256

6 *1. O

12 0-4096

18 h32768


15 *128

17 0-131072


---

---

---

0.1

0.1

1.0---

---

0.1

0.0625

0.00000048

0.00000095

0.125

0.0000305

1.0

0.00391

1.0

0.125---

0.004

1.0---

kHz

MHz

MHz

OEG:MIN

OEG:MIN

DEG

---

---

HR:MIN

KNOTS

oEG/180

oEG/180

KNOTS

0EG1180

KNOTS

DEG1180

MIN

NM

---

NM

FEET

---

---

---

---

N

E

---

---

---

---

ALWAYS POS

N FROM O

E FROM O

---

CW FROM N

---

CW FROM N

---

TO DEST

---

FLY LEFT

---

---

UpdateRate (SPS~

NOTE1NOTE1NOTE1NOTE1NOTE1NOTE2NOTE2NOTE2NOTE 2

NOTE 3

NOTE 3

10

10

10

10

10

10

10

10

10

10

10

10

10

ARINC 429 NAV Computer Output DataTable 206 “

NOTES : 1. Transmitted inl-second bursttransmitted at

2. Transmitted in

a l-second burst. Repetition rate during thewill vary depending upon other labels beingthat time.

a 500-ms burst. Repetition rates durina the 500-msburst will vary depending upon other labels being tran~mitted at thattime.

3. Transmitted continuously (100 ms) unless a position initialization isperformed at which time these labels are transmitted for a 500-msburst.

22-14-00Page 298.158

Aug 15/91Use or disclosure Of information on this page is subject to the restrictions on fhe title page of this document.

MAINTENANCE

Honeywell H!%!#h.Omega Station Deselect (Octal Label 272)

BIT 333222222222211111 11111000000000210987654321098765 43210987654321

P SSM Station manually deselected SDI 01011101

P - Odd Parity

SSM - 00 valid01 invalid

SDI - 00 all units01 unit #110 unit 11211 unit #3

Station Assignments - Logic 1 represents a deselect.

~ Station

12 JXN (Norway VLF)13 GOD (Anthorne VLF)

NAU (Puerto Rico VLF);: A (Norway Omega)16 B (Liberia Omega)17 C (Hawaii Omega)18 D (Dakota Omega)

E (Reunion Omega);: F (Argentina Omega)21 G (Australia Omega)22 H (JaDan Omeaa)

NWt:: NDT25 GBR

NAA:! NPM28 NSS29 NLK

~Austral~a’VLF)(Japan VLF)(Great Britain VLF) I(Maine VLF)(Hawaii VLF)(Annapolis VLF)(Washington VLF)

Output Word FormatsTable 207

22-14-00Page 298.159

Aug 15/91Use or disclosure of reformation on thm page IS subject to the restrrctlons on the title page of this document.

Date (Octal Label 260)

BIT 3332222222222111 11111110000000002109876543210987 6543210987654321

IIP SSM A B c D E F 00 00001101

P - Odd Parity

SSM - 00 valid01 invalid

Data - BCD DayAB, 1 through 31BCDMonth CD, 1 through 12BCDYear EF, 00 through 99

LNAV Status Word (Octal Label 275)

BIT 3332222222222111 11111110000000002109876543210987 6543210987654321

P SSM c N M o T/F A w 10111101I

P - Odd Parity

SSM - 00 valid

c

N

M

o

01 invalid

(CDI Flag) - ~ ~;~~$id

(Mag/True North) - 1 TrueO Magnetic

(Message Alert) - 1 MessageO No Message

(Lateral Offset) - 1 OffsetO No Offset

T/F (To/From) - ~~ $$ From or To

10 From11 Invalid

A (Approach Mode) - 1 ApproachO No Approach

W (Waypoint Alert) - 1 Waypoint AlertO Waypoint Alert

Output Word FormatsTable 207 (cent) 22-14-00

Page 298.160‘Aug 15/91

Use or disclosure of information on this page is subpcf to the restrictions on the tttle page of this document.

4. B. (4) Input and Output Control Discretes

The discrete inputs are in a ground condition when connected to anairframe dc ground, defined by a voltage of less than +3.5 volts dc.

The open condition is defined as greater than 100,000 ohms or avoltage between +18.5 and +36 volts dc.

The output control discretes are in the ground condition defined bya voltage of less than +3.5 volts dc.

The open condition is defined as qreater than 100,000 ohms or avoltage between +18.5 volts and +j6 volts dc.

Volume III, Section 6, Interconnects, Table 501,a detailed description of the control discretes.

(5) Global Data Management Unit (DMU) and DL-800/900Interconnect (See Figure 264.)

Appendix C contains

Data Loader

The DMU interface to the NAV computer is ARINC 429 as described inparagraph 4.B.(3) and the interface to the DL-800/900 Data Loader isa synchronous RS-422 serial digital data bus as described inparagraph 4.B. (2). When the data loader is connected to NAVcomputer No. 1, a ground is applied to 121J1B-83 from 123J1-E. Ifthe data loader is connected to NAV computer No. 2, a ground to

C121J1B-83 is applied from 123J1-D. Pin 123JI-F is connected if athird warm spare NAV computer is installed.

(6) Navigation Modes

The navigation computer develops a position based on a blend or mixof the sensors. The navigation modes are listed below by prioritybased upon developing the most accurate position blend:

(a) DME/DME

(b) DME/VOR

(c) IRS Only

(d) IRS/Omega Mix

The FMS freezes IRS drift in memory once IRS updating begins.When system accuracies fall below specified levels, IRS/Omegamixing begins, if the Omega is within 12 NM of the average IRSposition.

When the Omega is within 12 NM of the average IRS position, it

is mixed with the average IRS position, and will have the same

weight as one IRS. If the Omega position is outside the 12 NMlimit, then it will not be used in calculation of FMS position.

22-14-00Page 298.161/298.162


J_——. ————

lmJIA NZ.92D NAV COMPUTEq NO. I 1a

I ——— —121J1A 12aJl CD-S1 O CDU NO. I

mu ~ “

( 7 7gfyILJs’ f )‘MDATAT (LI IN I

;:A (w %( J )I

M M DATA

IN ~ ~ N &J”’r’ Ia121J1B I

C“u VALID lW 4 ~ C“u VALID(GNof

(RcvRm 7 e P (Nf

4 )

ICNTL M 8 XMTRCNrL

R 0-). . I

L ——. — J

I

1

-1 II ti-1

——— —

NOTETNE INTERFACE SETWEENTNE NZ-920 ANO TNE OL40CVS09USES A FIS422 ‘US. TNENTERFACE SETWEEN TNENZ-92D AN” W “MU USESAN AJw4c 42s ‘us,

I /$.——————

Cl 21JlA NZ42D NAV COMPUTER NO.2 1 II

IIF

————77L-SOO/900 DATA LOADE 1~1

— J————

cl~l CD-O1OCDU NO.21

1FATA ‘US OUT

H

(xMTR[W 32 m LrCNTL N .J3 * v

C“U vALID Im - iL

C121J1A

C“’r 0+) es

( 1

M

;ATA o-l 52 N

IIIII

I II pil,)(u RcvR cm

ICDU VALIO @NOI

LOAO LEFT E

i LOAD RIGNT o * 23 OATA LOAOER CONNECTED

1-

- 102 Am’ ENABLE

——. . =L

)(m DATA

,(Q “lJ’IIIIL ———————JGlobal DMU (AFIS) and DL-800/900

Data Loader InterconnectsFigure 264

AO-W3114+1

22-14-00Page 298. 163/298. 164

Apr 15/93

Useordisclosure01informationon this page is subject to the restrictions on the title page of this document,

4. B. (6) (e) VLF/Omega Only

The current Phase II certification allows GPS data to be viewedon the CDU, however the navigation computer does not use GPSdata in determining its position.

The -976 Navigation Computer (9101 software level) can use GPSposition information blended with other sensors in deriving theFMS position. GPS data may also be viewed on the CDU. The GPSinformation can come directly from the GPS sensors or throughhybrid IRS/GPS (GPIRS) units to the FMS. Regardless of how the~~~ information gets to the FMS, it is treated the same by the

The GPIRS units provide three sets of data (pure IRS,hyb~id IRS/GPS, pure GPS) to the FMS. The FFISuses the pureIRS data just as IRS has been used in the past. The FMS usesthe hybrid IRS/GPS or purse GPS as the GPS sensor information.By priority, the FMS uses the hybrid IRS/GPS information ifavailable and pure GPS only if the hybrid is not available.

With GPS position blending o erational, the FMS computes anaircraft reference position Eased on the best available sensorsexcluding GPS. The reference position is compared to GPSposition. If the positions agree within predefine tolerances,GPS position is blended with the reference position to form theFMS position. GPS data is not blended if it fails the positioncomparison.

Tolerances for com aring the reference and GPS positions wereIestablished with t e FAA. The tolerances for NZ-9101 software

are: 18.0 nautical miles (NMI) oceanic, 4.3 NMI enroute, 2.6NMI terminal area, and 0.4 NMI on approach (increased to 0.7NMIwhen using VOR/DME updating).

When the FMS is in radio updating (VOR/DME or DME/DME), theaverage GPS and radio positions are equally weighted inderiving the FMS position. Thus, in radio updating, the FMSposition could deviate a maximum 1/2 the above listedtolerances due to GPS.

When GPS is blended with other long-range navigation (LRN)sensors, each GPS and IRS contribute an equal share to the FMSposition. In the case of VLF/Omega, the avera e VLF/Omega

iposition is counted equal to a single GPS or I S sensor. Forexample, two

Ture IRS position inputs blended with a single GPS

position resu ts in an FMS position deviation due to GPS isone-third of the above listed tolerances. With three IRSS andtwo GPS sensors, and no other position information, the averageGPS position would account for 40 percent of the resulting FMSposition.

GPS is certified for supplement navigation only. This meansthat it cannot be used as the sole source of osition. Theremust be other means of determining osition,

7Iowever, where all

other sensors are deselected or fai ed, the FMS can navigatewith GPS as the only position sensor. In this situation, DGRADis illuminated and NO REQUIRED SENSORS is displayed on thecontrol display unit indicating the FMS cannot guarantee therequired accuracy for the present phase of flight.

22-14-00Page 298.165


4. B. (7) Configurations

When installed in an aircraft, the FMS may be configured with asecond FMS. There are four configuration states that the FMS mayexist in: dual, initiated transfer, independent and single.

. Dual mode allows all active flight plan entries and performanceinitialization data to be transferred to the off-side systemwithout the need for a separate pilot action.

● Independent mode exists only when two systems are installed. Inthis mode only off-side tuning information is transferred betweensystems.

. Initiated Transfer mode is applicable only when two systems areinstalled. In this mode off-side radio tuning commands and pilotdefined database entries are transferred automatically. Activeflight plan and performance initialization data are transferredonly on command either through an aircraft mounted switch or viaa line select key prompt on the last page of the active flightplan.

● Single is the configuration of a single FMS system or a defaultmode when two systems in a duplex installation are notcommunicating with the offside FMS.

~: Some installations have three (3) FMS systems installed.However, the third FMS interfaces with the totalintegrated system as a hot spare.

The FMS will operate in dual or initiated transfer modes only whenthe following criteria are met:

. Both systems must have the same software version (SW PROGRAM).

c Both systems must have the same navigation database region andeffective dates (NAV DB and DB CYCLE).

● Both systems must have the same programming pin configurations(CONFIG PIN).

● Both custom databases must be identical (CUSTOM DB).

● Both FMSS must have the same present position with a 10 NMtolerance (PPOS DIFF).

● Both FMSS must be in agreement as to the master-slaverelationship (SLAVE).

The parenthetical information above will be displayed on the CONFIGPROBLEMS page if the system is not operating in the mode for whichit has been configured.

22-14-00I

Page 298.166Apr 15/93

Use or disclosureof information on thispage issubject to the restrictions on the title page of this document.

4. c. PZ-800 Performance (Autothrottle) Computer

The PZ-800 provides two distinct functions. They are described in thissection separately as performance function and autothrottle function.

(1) Performance Function

The PZ-800 contains a digitized flight manual. Entries via the CDUallow the PZ-800 to calculate the desired performance values.

(a) Performance Initialization

Performance initialization requires entries from the pilot inorder to enable the performance calculations. These entriesare in addition to the active flight plan which containslateral information and may contain vertical information.Flight planning, climb, cruise, descent, and engine-outportions of the flight can be calculated using the performanceinitialization entries. If the calculation of takeoff andlanding data is desired, additional entries are required.

(b) Performance Planning

The performance planning function for the active flight planhas two primary activities. Prior to takeoff, the performanceplanning function calculates and displays the fuel and timerequired for the flight using the selected cruise altitude,step climb increment, and selected speed mode. If the optimumcruise altitude is used, it is calculated and displayed. Thesecalculations are based on the data entered during performanceinitialization. Following takeoff, the performance planningfunction updates the fuel and time calculations to reflectactual conditions and changes that the pilot may enter.Display of performance planning information includes predictedfuel remaining and estimated time enroute (ETE) at eachwaypoint.

Performance planning for a stored flight plan can be obtained.This is done by specifying the initialization data to be usedon the the stored flight plan to be used. This function couldbe utilized by the pilot to examine alternative flight routingsor future flight plans.

The PZ-800 in conjunction with the NZ-920 Navigation Computerform the core of the Honeywell flight management system (FMS).The FMS can accept flight plans from an external source such aslockheed jetplan. Once the flight plan is entered as theactive flight plan, the performance planning function is thesame as for a pilot-entered flight plan. Following performanceinitialization, the performance planning calculations areperformed for the flight plan which has been loaded from theexternal source. The computer does not use the performancevalues (e.g., fuel flow and fuel remaining) computed by theflight planning service. 22-14-00

Page 298.166.l/298.166~2Apr 15/93


Honeywell !!!!#f#.’E4. C. (1) (c) Takeoff Calculations

Following entry of the takeoff initialization data, the takeoffcalculations are performed. Takeoff initialization data can bechanged after the calculation of the takeoff data. If a changeis made, the takeoff data is removed from display until updatedto the change. The following takeoff data are calculated anddisplayed:

● Takeoff weight or most restrictive maximum takeoff weight

. Required takeoff field length (the longer of theaccelerate-go and the accelerate-stop (emergency) distances)

● Takeoff engine limit (EPR)

● Accelerate-go and

. Takeoff Speeds

- decision speed- decision speed- rotation speed- safety speed -

steadv initial

accelerate-stop (emergency) distances

- V1range - Vlmax, Vlmin*- VrV2climb speed - V4 (UK CAA only)

- final segment speed - Vfs --.

- single-engine speed - Vse- reference speed (landing configuration) - Vref

* A choice of V1 between Vlmax and Vlmin, inclusively, ispermitted if the case is not field-length-limited. Vlmaxis the default value for V1.

Once the takeoff initialization is complete, the takeoffconfiguration comparison logic is active. The configurationcomparison logic continuously compares the initializedconfiguration for bleed air, ground spoilers, anti-skid, andflap-position against the actual aircraft configuration.Barometric altitude, pressure altitude, and outside airtemperature are also continuously compared against theinitialized conditions. If the initialized configuration doesnot match the actual configuration, the display of takeoffspeeds is confined to the CDU and glareshield-mounted displaycontroller. When the initialized configuration matches theactual configuration, the takeoff speeds are available fordisplay on the primary flight display (PFD) from the avionicsstandard communication bus (ASCB). When both engines areadvanced above 1.17 EPR and the configuration comparison logicis satisfied, the logic is latched. This means that while thelogic is latched, the display of takeoff data is available fromthe ASCB regardless of changes to the actual aircraftconfiguration. If both engines are reduced below 1.17 EPR, thecomparison logic is unlatched and is again active.

22-14-00Page 298.167


4. C. (1) (d) Climb Calculations

During performance initialization, the pilot is required toaccept the default climb mode or select the desired climb mode.This is the mode used for the flight planning function.However, each climb mode is available for selection at anytime. For the active climb mode, the performance computercalculates and displays the following climb parameters:

. Top-of-climb (TOC) altitude

● Estimated time enroute (ETE) and estimated time-of-arrival(ETA) at the TOC altitude

. Command speed

. Distance-to-go (DTG) to the TOC altitude

. Command engine pressure ratio (EPR) setting

. Current fuel remaining

(e) Cruise Calculations

While in the cruise mode, the performance function calculatesspeed and recommended power settings. The autothrottlecontrols speed of the aircraft. Altitude is flown by theflight control system on the elevator. The followingparameters are calculated and/or displayed on the CDU cruisepages for the active cruise modes: “ -

. Cruise altitude*

c Optimum altitude

● Step altitude (based on initialized step alt”

c Command speed

● ETE and ETA to step climb point

. Command EPR setting

tude increment)

c DTG for step point

. Range to reserve

● Time to reserve


* Cruise altitude is the higher of the entered cruisealtitude or the preselect-altitude until matched by actualaltitude. 22-14-00

Page 298.168Aug 15/91


Honeywell &&!j!j#.cE4. C. (1) (f) Descent Calculations

During descent the following parameters are calculated and/ordisplayed for the active descent mode:

. Bottom-Of-Descent (BOD) altitude

. ETE and ETA to BOD altitude

● Command speed

. DTG to BOD altitude

. Command EPR setting*

* The descent EPR value is removed when flaps or landinggear are extended.

(g) Landing Calculations

With this information and that already in the computer,calculations are made for the landing and go-around parameters.The following landing parameters are calculated and displayed:

. Landing weight

. Landing distance (non-UKCAA -

. Landing configuration approach(non-UK CAA)

. Target threshold speeds - VatO

“dry”, UK CAA - “wet”)

speed, Vref = 1.3 V stall

and Vatl (UK CAA only)

s Maximum threshold speeds - all-engine and single-engine (UKCAA only)

. Dry runway landing field length (non-UK CAA)

. Wet runway landing field length (non-UK CAA)

Data are computed for weights greater than the certifiedmaximum landing gross weight, but these data are accompaniedwith a message denoting exceeding the maximum landing grossweight and the most restrictive applicable condition. Therestrictive conditions applicable may include: takeoff-climb-limited, approach-climb-limited,landing-climb-limited, field-length-limited, or certified maximum gross weight exceeded.

22-14-00Page 298.169

Aug 15/91lJseor disclosure of lnfOITrIatiOn on this page is subjecf to the restrictions on the title page of this document.

4. C. (1) (h) Go-Around Calculations

In addition to the landinggo-around calculations are

. Maximum landing weight*

. Approach climb speed inengine inoperative

calculations, the followingdisplayed:

approach configuration with one

. Landing climb speed in landing configuration with bothengines operative (non-UK CAA)

● Go-around EPR rating

. Climb EPR rating

* There is only a check again of the maximum landing weight(58,500); it is not separately displayed.

(i) Single-Engine

Single-engine calculations are made by the performance functionof the computer. There is no requirement for additional inputfrom the pilot. If an engine becomes inoperative in flight,the climb, cruise, and descent modes are automatically switchedto single-engine with the associated calculations beingperformed. The following information is available whileoperating with both engines, for fixed altitude cruise and forcruise-climb:

. Single-engine cruise altitude

● Single-engine drift up/down speed

● Active average flight plan ground range and time to reservefuel

● Active average flight plan ground range and time to zerofuel

(J) WtlAT-IF Mode

The WHAT-IF mode is used for investigating the results ofchanges to the performance inputs. Questions such as what if ahigher altitude is selected are answered by the WHAT-IF mode.The WHAT-IF mode includes climb, cruise, and descent phases offlight.

Any of the performance initialization inputs may be changed forthe WHAT-IF case. The default initialization values are thecurrently active values.

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Honeywell !!!$!!~.c’The following are calculated and/or displayed for the activeflight plan under the WHAT-IF conditions:

. Cruise altitude and ceiling altitude

● Step increment

. ETE to destination for active and WHAT-IF flight plans

● Fuel required to destination for active and WHAT-IF flightplans

● Optimum altitude

● Step altitude (based on initialized step altitude increment)

. Command speed

● ETE and ETA to step climb point

. Command EPR setting

. DTG for step point

. Range to reserve

. Time to reserve


. Cruise pages under WHAT-IF conditions

4. C. (1) (k) Stored Flight Plan Data

Performance data can be calculated for a stored flight plan.The performance initialization inputs may be changed to reflectthe expected conditions for the flight. The defaultinitialization values are the currently active values exceptfor fuel weight. The following is calculated and/or displayedfor the stored flight plan:

. Cruise altitude and ceiling altitude

● Step increment

. Time to destination

● Fuel required to destination

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Honeywell &!#!u!e.cE

4. c. (1) (1) Performance Priority Selections

The priority PZ-800 Performance Computer will follow the NAVpriority. In order to determine PZ-800 priority with on-sideNAV invalid, off-side NAV is also checked. When both NAVcomputers are invalid, the FMS displayed on the PFD selectedEFIS determines priority. With both NAVS and autopilotsinvalid, the FMS displayed on the pilot’s displays determinespriority. When one NAV only is valid, the PZ-800 on-side withthat NAV is the only PZ-800 that can set priority.

The priority PZ-800 uses the DADC as indicated by the FGC.When the FGC is invalid, the left EFIS DADC display determinesthe DADC selection.

For the flap settings, a comparison of both ADC flap settingsis made provided both DADC flap settings are valid. Thiscomparison has a one-second debounce. That is, itwill take 10miscompares (based on receiving the DADC every 100 ms) beforethe PZ-800 will declare flap position invalid. If they do notagree for the full second, the flap setting is marked invalid.If there is only one valid flap setting, that flap setting isreceived and marked valid.

FGC data comes from the priority FGC provided the data isvalid. If both FGCS are high priority or both are lowpriority, the FGC is considered invalid.

The performance computer receives the PFD selected DC-884Display Controller (DC) if valid. If the selected DC isinvalid, the other side DC is read.

If the PFD selection is neither or both (dual) the previouslyselected DC remains selected. The previous DC is initializedto left DC on power-up.

If there is no autopilot, the left DC is selected if valid.otherwise the right”

If there are no DCS

PERF1PERF2

DC is selected.

the following DAU channel

Default DAU Channel

(left) CHANNEL A(right) CHANNEL B

defaults are set:

If the left or right EPR value is manually set on the dis~laycontroller on the-second page of the TRS pages, then that”EPRis used as the rating. When the DC is in the DUAL manual EPRmode, the two EPR values will be the equal. Only when the modeis SPLIT can the two values differ. The lowest of the twovalues will always be used.

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If a single rating is selected on the DC TRS page 1, thisrating will be used as the rating by the perf computer. Thechoices are:

. Takeoff● Flex. Go-around● Climb● Cruise● Max Continuous

A DAU channel switch takes place if one or morefollowing parameters are invalid, as determined

of theby the validity

or the applicable range and rate tests: Nl, N2, EPR, fuelflow,and fuel quantity. The switch only takes place if thenonselected channel has all valid data. N2 is included in thislist to allow perf and autothrottle to use a common selectionlogic.

The DAU N1 reasonable test consists of range and rate checks.For the range test, the minimum rpm is O and the maximum rpmwill be 110 percent. The rate test will be failed if a rate inexcess of 5 percent per sample is detected. IfNl isconsidered invalid for a given channel of a DAU for 400milliseconds, then the data from the other channel will beused.

The DAU EPR reasonable test consists of range and rate checks.For the range test, the minimum EPR is 0.8 and the maximum EPRwill be 2.0. The rate test will be failed if a rate in excessof 0.06 EPR per sample is detected. If EPR is consideredinvalid for a given channel of a DAU for 400 milliseconds, thenthe data from the other channel will be used.

Fuel quantity has no validity so a simple range and rate testis performed. For the range check the minimum quantity is O lband the maximum is 16,500 lb for each DAU. We would expect theperformance computer to continue to operate properly with thesevalues since they are valid inputs to the tables. The flightmanual lists a maximum value of 14,750 lb. The total fuelquantity is the sum of the two separate values from the twoDAUS. The rate check fails if the change is greater than 875lb per sample. The validity should be set valid only if bothDAUS have valid data.

Fuel flow has no validity so a simple range and rate test isperformed. The fuel flow must be between O and 10,485 lb/hr(per the ASCB specification). The max value in the FUELFLOWtable is 10,355 lb/hr, and the performance calculations willoperate without error up to this value. Total fuel flow is thesum of the two separate values from the two DAUS. The ratetest fails when the change is greater than 320 lb/hour persample. The validity should be set valid only if both DAUShave valid data. 22-14-00

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4. C. (2) Autothrottle Function

(a) Introduction

The functions which the autothrottle is capable of providingvary depending on the phase of flight and operating modes ofother systems which interface with the autothrottle (i.e.,navigation computer, flight guidance computer (FGC)). Theautothrottle has four basic modes of operation: takeoff,go-around, flight level change (FLCH) mode, and speed hold.

When engaged and operating in the FLCH mode (FLCH mode is setby the flight guidance computer) the autothrottle commands theenergy required by the flight guidance computer to accomplishthe FLCH climb or descent.

When engaged and operating in speed mode, the autothrottleadjusts engine thrust via separate throttle servos (one servofor each engine) to control aircraft speed to the target speedwhich is displayed in the cockpit. The target speed may bemanually input or slaved to a speed determined by theperformance function depending on performance mode and pilotselection.

Other autothrottle functions such as maximum operating speed(VN~MO) protection, flap placard protection, extended landinggear ~lmits, engine limit protection, and minimum thrustcontrol are performed in parallel with the basic modes. Theautothrottle also provides engine low-rotor speed (%N1) orhigh-rotor speed (%N2) synchronization in speed-hold mode as afunction of pilot selection.

The only valid bleed configurations for autothrottle operationare: (1) cowl and wing anti-ice off, (2) cowl anti-ice on, (3)cowl and wing anti-ice on. The environmental control system(ECS) is assumed to be on and the isolation valve closed at alltimes. If an invalid aircraft bleed configuration is sensed,the autothrottle disconnects and an appropriate advisorymessage is displayed. An appropriate delay is provided whenchanging bleed configurations to prevent nuisance disconnects.

The autothrottle performance limits are listed in Table 208 andthe advisory messages which appear on the EICAS display arelisted in Table 209.

22-14-00Page 298.174


Ambient temperature dictated engine limitations are determinedvia the aspirated ambient air temperature probe only. Pilotentry of static ambient temperature is not used by theautothrottle.

Descriptions of specific functional capability duringrepresentative flight phases are given below:

4. C. (2) (b) Takeoff

When the autothrottle is engaged and takeoff is initiated(requires cockpit input, minimum engine power setting, andsatisfaction of enabling conditions), the throttles areautomatically advanced to the required takeoff engine settingbased on a full-thrust (MIN EPR) takeoff or a reduced-thrust(FLEX EPR) takeoff. A FLEX EPR takeoff requires either anoperating performance function and a CDU selection by the pilotduring performance initialization,or a pilot input of enginepressure ratio (EPR) command on the manual thrust referencepage of the display controller. At 60 knots the autothrottleservos are depowered with the clutches remaining engaged(throttle hold) and remain depowered until the aircraft hasclimbed to at least 400 feet above lift-off altitude (seeFigure 264.1). An indication of the throttle hold condition isdisplayed to the pilot. At 400 feet, the servos are repoweredafter a positive pilot action (required) is made indicating atransition from takeoff (or if the flight director captures theselected altitude). This action will be the selection of aflight guidance computer mode that causes the takeoff discreteto be dropped.

(c) Climb

During the climb phase, the autothrottle mode is determined byflight guidance computer operation. If the autothrottle isengaged and the autopilot is either in vertical speed mode orno mode is selected, the autothrottle adjusts thrust withinengine limits to control aircraft speed. Enginesynchronization is provided in speed hold mode as a function ofcockpit inputs.

If the FGC is in flight level change mode, the autothrottleincreases thrust (limited to the selected engine rating) toprovide the aircraft with the energy required to accomplish theFLC climb.

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4. C. (2) (d) Cruise

During cruise, the autothrottle adjusts thrust within thelimits of minimum thrust and the selected engine rating tocontrol aircraft speed.

(e) Descent

I)uring the descent, the autothrottle mode is determined byflight guidance computer (FGC) and navigation computeroperation. If the autothrottle is engaged and the FGC is in

vertical speed, VNAV path, or no mode is selected, theautothrottle adjusts thrust within the limits of minimum thrustand the selected engine rating to control aircraft speed.Engine synchronization in speed hold mode is provided as afunction of cockpit inputs.

If the FGC is in flight level change mode, the autothrottledecreases thrust as necessary to provide the autopilot with theenergy required to accomplish the FLC descent. If the descentis being made with vertical navigation (VNAV) guidance, minimumthrust control is provided.

(f) Approach

When the FGC is in glideslope, path, or vertical speed modeduring approach, the autothrottle adjusts thrust to controlaircraft speed.

(g) Go-Around

When engaged and go-around is initiated (requires cockpitinput), the autothrottle automatically increases thrust to thego-around engine limit.

22-14-00Page 298.176


FvT IN TAKE OFF (TO) MODEAUTO EPR RATING - TO/FLEXAT Ml KTS IAS AUTOTHROITLE CLAMPS THROITLES (HOLD MODE) NOTE 1

NO NT REMAINS IN HOLD MODEAUTO EPR RATING REMAINS TO I FLEXV SPEEDS ARE DISPIAYED

I YES

A/T REMAINS IN HOLD MODEAUTO EPR RATING REMAINS TOIFLEXV SPEEDS REMAIN UNTIL Vse + 50 (-906 PZ) OR 200 KTS (-910 PZ)

IYES

~

V SPEEDS ARE CANCELLED IN -906 PZV SPEEDS ARE CANCELLED AT 200 KTS IN -910 PZ

A/l MAINTAINS EPR RATING SELECTEDAUTO - CLB OR GA (GA MODE SELECTED)

I A/T MAINTAINS SPEED AND LIMITS EPR TO RATING SELECTEDAUTO - CLB OR GA (PITCH HOLD MODE) NOTE 3

I

NOTES1. WHEN IN HOLD MODE, THE AUTOTHROITLE DOES NOT CORRECT

FOR ANY CHANGES IN EPR.

2. TO MODE CAN BE CANCELLED BY SELECTING ANOTHER VERTICALFLIGHT DIRECTOR MODE OR AUTOMATICALLY BY CAPTURING ASEL.

3. ENGAGING THE A/P WHEN IN TO MODE WILL CHANGE THE VERTICALMODE TO PITCH HOLD AND AUTO EPR RATING TO GA. SELECTING TCSWILL CANCEL TO MODE AND ALSO SETS THE AUTO EPR RATING TO GA. AD-34324 @

Takeoff Mode Flow ChartFigure 264.1 22-14-00

Page 298.176.1/298.176.2Apr 15/93



SPEED DADC Mach Engage Range 0.4 to 0.88 M(IAS/MACH)

Mach Hold Error tool M

IAS Engage Range 80 to 340 kts (CAS)

IAS Hold Error 25 kts

Throttle Control 2“ to 39.5° PLAAuthority

Throttle Rate Limit f7° pLA/SECOND

FLCH EPR EPR Control Range 0.85to 2.0 EPRCLIMB

EPR Hold Error ~().olEpR

Throttle Control 2° to39.5 PLAAuthority

Throttle Rate Limit ~7” pLA/SECOND

FLCH PLA PLA Setting 2° PLADESCENT (Power Lever Angle)

TAKEOFF EPR EPR Control Range 0.85 to 2.0 EPR(TO)

EPR Hold Error ~().ol EpR

Throttle Control 2° to 39.5” PLAAuthority

Throttle Rate Limit ~lo” pLA/SECOND

GO-AROUND EPR EPR Control Range 0.85to 2.0 EPR(GA)

EPR Hold Error ~oool” EpR

Throttle Control 2° to39.5 PLAAuthority

Throttle Rate Limit f7° pLA/SECOND

Autothrottle Performance LimitsTable 208

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MessageTimed Out

Messages (5 See) Description

AT 1(2) FAIL No A/T has failed.

AT ENGAGEINHIBIT Yes Manual selection is inhibited due to engageinterlock.A/T cannot be engaged under the followingconditions:

. Aircraft on ground (TO not selected orEPRc 1.17)

. EPR invalid for thrust modes

● IAS/MACH target invalid for speed modes

● A/T quick disconnect switch depressed

. Valid preselect altitude

● EPR rating not selected

. LP (Nl) split of 20 percent

. Bleed air isolation valve open

. Engine out (HP (N2) < 41 percent)

NOTE: This message will not be active unlessthe A/T ARM mode is selected.

AT NOT IN HOLD No Message appears with A/T engaged in TO modewhen CAS is greater than 60 knots and theHOLD mode is not active. Message automat-ically cancels when RAD ALT is greater than400 ft AGL.

Advisory Messages (Blue)Table 209

22-14-00Page 298.178


4. c. (3) Autothrottle System Description

Each PZ-800 Performance Computer contains an input/output card, anautothrottle card, and a servo card. Two PZ-800 PerformanceComputers are included for redundancy; however, only one computercontrols both throttles when selected. The PZ-800 processes inputsfrom many different sources. The primary interface is the ASCB.This high-speed bidirectional bus supplies a wealth of informationfrom the rest of the SPZ-8000 Digital Automatic Flight ControlSystem (AFCS). Other interfaces include hardwired aircraftdiscretes, analog input/output for servo control, and RS-232 fortest and debug.

The GP-820 Flight Guidance Controller is centrally located on theglareshield and is used to control both the autopilot and theautothrottle. An A/T ARM pushbutton is provided to enable theautothrottle system. Speed and altitude selections are tieddirectly to the AFCS and passed to the autothrottle via ASCB.Likewise, several flight guidance mode selections are available;these are also relayed to the autothrottle by the AFCS. Thisensures that autothrottle control modes are always coordinated withAFCS modes. Speed selection can be made in terms of calibratedairspeed or Mach number, or if FMS mode is active, the speed commandis determined by the flight management system and displayed in thesame window on the GP-820 Flight Guidance Controller.

In addition to the arm button on the GP-820 Flight GuidanceController, there are several cockpit switches which directly affectautothrottle operation. Engage/disengage, positive-disconnect, andtakeoff/go-around switches are located on the throttles. Thepositive disconnect switch provides a direct-to-hardware means fordisengaging the autothrottle servos. The takeoff/go-around buttonis used to select takeoff on the ground or go-around mode in the

air. Engine synchronization switches (overhead panel) select engine

synchronization and the engine parameter to be equalized.

The autothrottle uses information broadcast on ASCB by the FZ-820Flight Guidance Computer and other systems. The other systemsinclude DC-884 Display Controllers, Inertial Reference Systems,AZ-81O Digital Air Data Computers, NZ-920 Navigation Computers,PZ-800 Performance Computers, DA-884 Data Acquisition Units, andFC-880 Fault Warning Computers.

Display controllers (DC) are located to the left and right of theflight guidance controller on the glareshield; engine limitsselected on the DC are used by the autothrottle. Inertial referencesystems (IRS) provide accelerationmeasurements, as well as certainaircraft parameters (e.g., pitch angle, roll angle). Digital airdata computers (DADC) supply altitude, temperature, and airspeeddata, as well as flap position and V~O values. Navigation andperformance computers optimize aircraft performance and can becoupled to the autothrottle (and AFCS) for automatic flight. Dataacquisition units relay important engine data and aircraftdiscretes. Fault warning computers interface with the autothrottleto log faults and run ground maintenance tests.

22-14-00Page 298.179


Finally, two SM-810 Servos are included toThese hiqh-reliability servos each consist

drive the throttles.of a servo motor, qear

train, electrical eng~ge clutch, mechanical slip clutch, and ~ drumgrooved for aircraft cable. The dual-servo configuration allowsaccurate and independent thrust setting on both engines in takeoff,go-around, and flight level change modes. Synchronization of enginerotor speeds, side to side, is made possible by independent controlof each throttle. When the autothrottle is engaged, the PZ-800Performance Computer outputs analog rate commands to the servos tocontrol thrust. Hardware and software monitoring is used to ensurethe health of the system.

4. c. (4) Autothrottle Modes

Autothrottle modes are closely tied to flight guidance modes and canalso be influenced by the flight management system (navigation andperformance computers). Basically, the autothrottle sets enginethrust for takeoff, flight level change (FLCH), and go-around modes.At all other times it controls airspeed.

The best way to describe autothrottle modes is to step through atypical flight and discuss system operation along the way. Figure265 gives an outline of flight guidance and autothrottle modes foreach phase of flight.

(a) Takeoff

As the aircraft taxis out to the runway for takeoff, the pilotenables the autothrottle system by pressing the arm button onthe flight guidance controller. Takeoff mode is selected bypressing the takeoff/go-around button on either throttle. Oncethe aircraft has been cleared for takeoff, the pilot moves thethrottles above a minimum engine pressure ratio (EPR) thresholdand engages the autothrottle. Throttles are automaticallyadvanced to the required engine power setting for takeoff. ThePZ-800 Performance Computer has data stored internally todetermine the appropriate EPR for a full-power takeoff (as afunction of altitude and temperature). However, if a reducedthrust flex takeoff has been selected, the performance computerdoes the necessary computations and supplies the EPR value tothe autothrottle or the pilot could dial in a manual EPR valuethrough the DC-884 Display Controller.

Each throttle is driven independently by a separate servo whichallows accurate closed-loop EPR control of both engines. At 60knots, the servos are depowered and remain in throttle-holduntil the aircraft has climbed to at least 400 feet aboveground level. An indication of the throttle-hold condition isdisplayed to the pilot and monitored by the FC-880 FaultWarning Computer. The servos are repowered, above 400 feet,when a mode other than takeoff has been selected by the pilotor ASEL is automatically captured.

22-14-00Page 298.180


Honeywell !!!!!!%5.”

FLIGHT FLIGHT DIRECTOR AUTOTHROITLE AUTOPILOTPHASE MODE FUNCTION FUNCTION

1 Takeoff Takeoff Maximum rated thrust OR NIA“FLEX” reduced-thrust ORmanually-set thrust

2 Climb-Out Flight Level Takeoff to climb Airspeed

Change thrust transition control

3 Small-Step Flight Level Reduced climb thrustClimb

Airspeed

Change for smooth transition control

4 Large-Step Flight Level Full climb thrust AirspeedClimb Change oontrol

5 Top-of- Altitude Transition to AltitudeClimb Capture airspeed control capture

6 Cruise Altitude Airspeed control Altitude

Hold hold

7 Top-of- Flight Level Transition to minimum AirspeedDescent Change thrust control

8 Descent Flight Level Maitiln minimum AirspeedChange thrust oontrol

9 Approach Glideslope Airspeed control. GlideslopeTrack Proteot flap/gear track

limits.

10 Landing NIA D@engage Disengage

11 Go-Around Go-Around Maximum rated Disengagethrust

AD-219$

Basic Autothrottl e FunctionsOver the FIight Prof i1e

Figure 265 22-14-00Page 298.181

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4. c. (4) (b) Climb

Once the aircraft is airborne, the pilot normally engages theAFCS (autopilot) and selects flight level change. When FLCHclimb is initiated, the autothrottle smoothly sets thrust to anappropriate level to accomplish the climb; the AFCS controlsairspeed through the elevator. In a large-step climb, theautothrottle sets and maintains thrust at the selected enginerating (usually maximum climb EPR). Without an autothrottle,the pilots would have to constantly monitor engine parametersand make throttle adjustments throughout the climb. In asmall-step climb, a unique proportionality scheme provides anappropriate energy level while always ensuring a minimum rateof climb. This reduced-thrust-level feature for small climbshas been popular with the pilots.

Other flight guidance modes can be used for climbing. Pitchhold is the default mode for the AFCS. If the autopilot is

engaged without selecting a mode, it maintains current pitchangle while the autothrottle controls airspeed (within enginelimits). When vertical speed mode is selected, the AFCScontrols the aircraft rate-of-climb and the autothrottlemaintains selected airspeed. If the autothrottle is unable toachieve the desired airspeed without exceeding engine limits, apower-limit message is displayed to the pilot.

(c) Cruise

As the airplane nears the selected cruise altitude, the AFCSautomatically initiates an altitude capture and theautothrottle smoothly transitions to controlling airspeed.Control laws have been developed to give a smooth transitionwith good speed control and without excessive throttlemovement. Altitude hold by elevator control (AFCS) and speedhold by throttle control (autothrottle) are the normal modes ofoperation in cruise.

Several features of the autothrottle control laws contribute tothe overall low-activity design. Complementary filtering usesmeasured airspeed together with inertial acceleration invarying proportions as a function of altitude and phase offlight. Aircraft drag information included in the autothrottleis used to anticipate thrust change required to compensate forconfiguration changes (flaps, landing gear, and turns).

Also, control of aircraft acceleration and deceleration isprogrammed to smoothly capture a new selected airspeed. Thesefeatures, together with appropriate gains, provide accuratespeed control with minimal throttle activity.

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Honeywell !?!!!i~.c’4. C. (4) (d) Descent

To descend, the pilot may select flight level change mode.When FLCH is selected, the autothrottle smoothly reduces thethrottle setting to a minimum thrust level to allow theaircraft to achieve a reasonable rate of descent. The AFCScontrols airspeed through the elevator, as in FLCH climb.

Other potential flight guidance modes for descent includevertical speed and pitch hold. Also, a VNAV path mode isavailable for descent where the FMS coordinates with the AFCSto maintain a programmed path angle. The autothrottle controlsairspeed (within engine limits) in all three of these descentmodes.

(e) Approach

During an instrument landing system (ILS) approach, the AFCSworks with the FMS to track the localizer and glideslope, whilethe autothrottle controls airspeed. If the speed selector isin the automatic (FMS) mode, speed targets are supplied by thePZ-800 Performance Computer. As the aircraft nears thedestination, the speed target is automatically set to 10 knotsbelow the limit for 10 degrees of flaps. When the pilotselects the 10-degree flap position, the speed target is set toallow 20 degrees of flaps. As the approach progresses, thespeed target is reduced until it is set to Vref + 10 knots forfinal approach. The autothrottle is designed to accuratelycontrol airspeed without excessive throttle activity orsensitivity to turbulent conditions. The Gulfstream IVautothrottle was certified for category II weather minimums.

(f) Go-Around

When engaged in any mode, autothrottle go-around can beinitiated by a single push of the takeoff/go-around button.The autothrottle provides an automatic thrust increase to theinternally-computedmaximum engine rating. Go-around mode isdiscontinued when another flight guidance mode is selected bythe pilot. The AFCS disengages when go-around is initiated andthe flight director provides command bars on the primary flightdisplay.

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4. c. (5) Other Autothrottle Functions

Several autothrottle functions are performed in addition to thebasic operating modes:

(a) Engage/Disengage

While normal engagement and disengagement is by the switch onthe throttles, other means are available to disengage theautothrottle. As mentioned earlier, a positive disconnectbutton provides a direct-to-hardware path for disengagement,or the pilot can manually move the throttles, causing anautothrottle disconnect. Logic is included in the autothrottleto detect an engine failure and immediately disengage. If afault is detected in the autothrottle system or in a systemproviding required inputs, the autothrottle will alsodisconnect.

(b) Airspeed Limit Protection

Maximum airspeed limits are observed by the autothrottle whenit is operating in speed hold mode. These limits includemaximum operating speeds (VM~M ), as well as placards for all

‘!“flap and landing gear conflgura ions. Minimum speed protectionis Drovided whenever the system is ot)eratinain the automatic(FM~) speed selection mode: In this’mode,set by the PZ-800 Performance Computer anddrop below Vref.

(c) Engine Limit Protection

~he speed command isis never allowed to

The autothrottle uses the selected EPR rating as an upper limitfor control in all modes of operation. Automatic ratingselection is available, where the PZ-800 Performance Computerchooses the engine rating based on the phase of flight. Also,minimum engine limits are observed to avoid the nonlinearflat-response area near idle throttle settings.

(d) Synchronization of Engine Rotor Speeds

Engine rotor speed synchronization is provided by theautothrottle when it is engaged in speed hold mode.Low-pressure rotor speed (Nl) or high-pressure rotor speed (N2)can be selected by the pilot for synchronization. Normally, N1synchronization is selected in cruise to prevent any potentialbeating noise from the engines. If neither rotor speed isselected, a default EPR equalization is in effect.

22-14-00Page 298.184


Honeywell !l!!!!b.c’4. c. (5) (e) Mode Annunciation

Autothrottle modes are annunciated on the primary flightdisplay. Basic operating modes are displayed as either armed(white) or captured (green). Several limits are alsoannunciated. If the autothrottle is unable to achieve theselected airspeed, it will annunciate the reason. For example,the active window might show: VMO, MMO, FLAPS, GEAR, or POWER.The first four indicate speed limiting for specific aircraftconfigurations. POWER is shown when an engine limit ispreventing the autothrottle from reaching the desired speed.

(6) Autothrottle Mode Engagements

(a) Autothrottle Engage Logic (See Figures 266 and 267.)

Autothrottle controls are located on the GP-820 Flight GuidanceController and on each power lever.

Autothrottle (A/T ARM) Button - This pushbutton is locatedon the GP-820 and is used to arm/disarm the engage/disengageswitches on the power levers. This button must be armedbefore the A/T recognizes any activation from the throttlemounted switches.

Autothrottle Engage/Disengage (A/T ENG/DISENG) Switches -Each power lever handle has an A/T ENG/DISENG switch locatedon the under side of the power lever handles (see Figure266). When the A/T is armed, activation of either switchwill engage the A/T provided the engage criteria is met.

Autothrottle Quick Disconnect (A/T DISC) Switches - Eachpower lever handle has an A/T DISC switch located on thefront of the power lever handle (see Figure 266). Thisswitch is used to disconnect the A/T and also cancels theA/T OFF light.

22-14-00Page 298.185


bTO/GA

\

t

AI’-rENGI

DISENG

JLm

EN&DISENG

VfEW FROM REAR(PILOT SEAT)

AD.20S07

Autothrottle Switch Locations on Power LeversFigure 266

The autothrottle can be engaged, via the A/T ENG/DISENG switch,any time the following criteria is met:

● A/TARM light is annunciated on the GP-820 Flight GuidanceController.

. A speed target is displayed in the GP-820 speed window (forspeed mode only).

. An EPR limit rating has been selected (MAN or AUTO) fordisplay on the DC-884 Display Controller.

. Bleed air isolation valve must be CLOSED.

. N1 cannot be split by more than 20%.

● No engine out condition (N2 < 41%).

. Valid preselector. 22-14-00Page 298.186


Honeywell !!i!r.”For engagement, the A/T requires the following componentsvalid:

. FGC 1 or FGC 2

c IRS 1 or IRS 2 (or both)

. DADC 1 or DADC 2 (or both)

. DAU 1A or DAU lB (or both)

. DAU 2A orDAU 2B (or both)

. DC 1 orDC 2 (or both)

● At least one bus controller

For engagement on the ground, the following additional criteriamust be met:

. The flight director takeoff mode must be selected with thepower lever mounted TO/GA switches.

. Engine EPRmust be set greater than 1.17 EPR.

The fault warning computer provides a signal to the tonegenerator for manual and automatic A/T disconnects.Additionally, an A/T OFF light is annunciated for eachdisengagement. The light is active for 1 second on a pilot-activated disconnect and flashes steady for any automaticdisconnect. The steady annunciation can be cancelled byengaging the A/T or depressing the A/T DISC switch on eitherpower lever.

In the event of failure of the selected autothrottle, an amberAT annunciation is displayed in the autothrottle modeannunciation box (center) on the PFD.

22-14-00Page 298.187/298.188


———— ———— ———— ———— ———— ———— ———— ———— ——. .2+00 PERFORMANCE COMPUTER (PILOTS)

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Autothrottle Engage Logic DiagramFigure 267 22-14-00

Page 298. 189/298. 190Apr 15/93


4. C. (6) (b) Autothrottle Arm, Takeoff, and Hold Mode Select (Figure 268).

~ Autothrottle Arm Mode

With A/T ARM ON on the GP-820 Flight Guidance Controller,and the autothrottle disengaged, the A/T outputs white armmode annunciations in the A/T mode annunciation window onthe PFD. These annunciations turn green when anautothrottle mode is engaged.

~ Autothrottle Takeoff Mode

Takeoff (TO) mode is used to advance and set takeoff r.)ower.On

●

●

TO

the g~ou~d, this mode is armed by:

Selecting A/TARMON on the GP-820

Selecting flight director TO mode with TO GA switch oneither power lever.

is annunciated in green in the PFD vertical mode windowand in white in the PFD autothrottle mode window.

Takeoff is captured following ARM by:

. Advancing both power levers above 1.17 EPR

. Selecting A/T ENG with either power lever mountedENG/DISENG switches.

The white TO annunciation in the autothrottle window changesto green and is boxed for 5 seconds.

~ Autothrottle Hold Mode

The hold mode is automatic during takeoff. The autothrottleclamps the power levers when 60 knots is achieved. Thismode is annunciated on the PFD in green as HOLD. The powerlevers remain in HOLD until 400 feet AGL as a minimum andwill stay in HOLD until another mode is selected.

22-14-00Page 298,191/298.192


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Autothrottle Arm, Takeoff, and Hold ModeSelect Diagram

b——— —,——— —___ ____, JAD.30613.Rl#

Figure 268 22-14-00Page 298. 193/298. 194

Apr 15/93Use Dr disclosure of information on this page is subject to the restrictions on the title page of this document.

4. c. (6) (b) ~ The autothrottles are disengaged by:

. Moving either power lever mounted A/T ENG/DISENG switch

. Pressing either power lever mounted A/T disconnect switch

. Manually moving the power levers

. Deselecting A/T ARM button on the GP-820 Flight GuidanceController

● Selecting the cross-side A/T via the DC

● Deselecting the EPR limit rating on the DC


. Engine out (N2 < 41 percent)

● Bleed air isolation valve open

. FGC 1 and 2 fail

. A/T fail

. An autothrottle limit mode will drop the correctautothrottle mode but will not disconnect theautothrottles.

(c) Autothrottle F1ight Level Change, Speed (IAS/MACH), andGo-Around Mode Select (Figure 269).

~ Autothrottle Flight Level Change Mode

This mode of the autothrottle is active whenever the flightdirector is in a flight level change (FLCH) mode (IAS, MACH,VIAS, VMACH). This mode is annunciated as FLCH, in green,in the PFD autothrottle mode annunciation window. In thismode, the autothrottle is setting power (a target) for theclimb or descent, which is normally full climb power oridle, as appropriate. If the amount of climb is less than6000 feet, at low altitudes, the autothrottle sets less thanfull climb power. The amount of thrust provided is limited,in all cases, by the EPR LIMIT rating value selected fordisplay on the engine EPR display.

To

●

●

●

engage the mode:

A/T ARM is pressed on and A/T is engaged.

Press the FLCH button. If the performance computer hasbeen initialized, that speed target will automaticallyappear as the FLCH speed target, otherwise a manual speedset may be used.

A/T will provide the correct amount of thrust.

22-14-00Page 298.195


4. C. (6) (c) Z Autothrottle Speed Mode

The speed mode is the primary mode of the autothrottle.This mode is active anytime:

. A/T is engaged

and

. Flight director modes are not

- takeoff (TO)- go-around (GA)- airspeed or Mach (IAS, MACH, VIAS, VMACH).

Once active, the A/T speed mode holds the speed targetdisplayed in the GP-820 Flight Guidance Controller. Thespeed target comes from:

. FMS (performance)

. Pilot selecting the SPD manual (MAN) mode and setting avalue on the GP-820 Flight Guidance Controller.

This mode is annunciated in green as IAS or MACH dependingon the speed target.

~ Autothrottle Go-Around Mode

The go-around mode is used to rapidly advance and setgo-around power. This mode is selected by:

. A/T is engaged

. Select flight director GA mode with TOGA switch on eitherpower lever.

GA is annunciated in green in the PFD vertical mode windowand in green in the PFD autothrottle mode window. This moderemains active until another flight director mode isselected.

22-14-00Page 298.196


4. c. (6) (c) ~ The autothrottle is disengaged by:

● Moving either power lever mounted A/T ENG/DISENG switch

● Pressing either power lever mounted A/T disconnect switch

● Manually moving the power levers

. Pressing A/T ARM button on the GP-820 Flight GuidanceController

. Selecting the cross-side A/T via the DC

. Deselecting the EPR limit rating on the DC

● Loss of EPR rating


. Engine out (N2 < 41 percent)

. Loss of engine data

. A/T Fail

● FGC 1 and 2 Fail

G DADC 1 or DADC 2

. DClorDC2

. IRS 1 or IRS 2

● OAT< -70 “C (-80 “C for -910 PZ)

● Bleed air isolation valve open

. Speed target invalid (speed hold mode only)

● An autothrottle limit mode will drop the currentautothrottle mode but will not disconnect theautothrottles.

(d) Manual Override

If during autothrottle operation the pilot manually moveseither of the power levers, the A/T autothrottle will disengageand momentarily annunciate OVRD on the PFD in green.

22-14-00Page 298.197/298.198

Apr 15/93Use or disclosure of information on this page is subject to the restrictions on the title page ofthis document.

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Autothrottle Flight Level Change, Speed (IAS/MACH),and Go-Around Mode Select Diaqram

Figure 269 22-14-00Page 298. 199/298.200

Apr 15/93Use or disclosure Qf information on this page is subject to the restrictions on the title page of this document.

4. C. (6) (e) A/T

The

(7)

Limit Mode Operation

autothrottle is Drocwammed to t)rotectsr)eedand thrust1 imits during the va~iofis phases of flight.“ As speedapproaches the appropriate limit, the A/T active mode, shown onthe PFD, is turned white and moved to the arm side of theA/Tannunciation window. The appropriate limit is annunciated ingreen on the PFD. The limits protected by the A/T arebelow with the mode annunciation:

Limit Annunciation

Landing Gear Operating Speed(VL~MLO) GEAR

Flaps Extended Speed(vFE/MFE) FLAPS

Power (selected EPR) Limited POWER

Maximum Operating Limit Speed(V~~MMO) VMO or MMO

shown

When the autothrottle performance limiting factor is corrected,the previously active mode will be restored in green on thePFD. All mode transitions are shown boxed for 5 seconds.

A/T Priority Description

The autothrottle system uses two identical performance computers.Only one computer is active at any one time. The pilot can manuallyselect A/T 1 or A/T 2 through the display controller sensor page.This selection can be made on either the pilot or copilot displaycontroller (both controllers always show the same A/T selection).

Full system performance capabilities are provided independent ofwhich A/T is active (i.e., there is no operational advantage in thepilot manually selectingA/T 2 insteadofA/T 1). The system willpower-up with A/T 1 selected. Since the autothrottle system is failpassive, failure of the active A/T performance computer or selectionof the other A/T performance computer via the DC will result in anA/T disconnect.

22-14-00Page 298~201

Aug 15/91Use or disclosureof informationon thispageis subject to the restrictions on the title page of this document.

Honeywell

4. c. (8) Engine Sync

The engines may be selected


to sync to either N1 (LP) or N2 (HP).This w;ll be djsplayed on the engine instruments display. Wheneverthe autothrottles are engaged, the aircraft engine sync system isdisabled and the autothrottles will sync LP or HP for the engineswhen in the autothrottle speed hold mode, annunciated as IAS, or

MACH in the autothrottle window on the PFD. The autothrottleswillequalize EPR when in the autothrottle FLCH, TO, or GA modes. Table210 lists how engine synchronization is performed by theautothrottles.

Flight Director Autothrottle Parameter System ProvidingMode Mode Sync’ed Sync Function

N/A Not Engaged LP or HP Aircraft System

DESCENTIAS, MACH, 2 DEG PLAVIAS, VMACH FLCH CLIMB Autothrottle

EPR

I TO I TO I EPR I Autothrottle I

I GA I GA I EPR I Autothrottle IALT, VALTASEL, VASEL IAS or MACH LP or HP Autothrottle

VsVPATHGS

PITCH HOLD

Engine SynchronizationTable 210

22-14-00Page 298.202


4. c. (9) Autothrottle Mode F1ow

Figure 270 shows a schematic of control laws for the four basicmodes of autothrottle operation.

(a) Takeoff/Go-Around Mode

When the autothrottle is operating in either takeoff orgo-around mode, the throttles are independently adjusted tocapture the commanded EPR. The control achieves this EPR levelby commanding a throttle rate proportional to the EPR error.The increased bandwidth requirement during takeoff (thecommanded thrust must be attained prior to throttle-hold) callsfor a high loop gain. This could result in thrust overshootingthe command if additional compensation were not applied. Thecompensation used to preclude overshoots is in the form ofcommand biases which are proportional to EPR rate and throttlerate.

(b) F1ight Level Change (FLCH) Mode

In flight level change (FLCH) mode, the throttles arecontrolled to maintain a commanded EPR. During FLCH climb, theEPR command is determined from the selected engine rating andthe magnitude of the requested altitude change. For largeclimbs, the full-thrust rating is commanded, while smalleraltitude changes result in an EPR command which is proportionalto the size of the climb increment. During FLCH descents, aminimum thrust level is commanded. When FLCH mode isinitiated, the FLCH EPR command processor smoothly adjusts EPRcommand to the appropriate level. This gradual change incommanded thrust minimizes the effect on autopilot speedcontrol. The required EPR is maintained by commanding athrottle rate which is proportional to the EPR error. Unliketakeoff operation, additional compensation is not required dueto the lower loop gain in FLCH. The requirement for a smoothtransient during FLCH allows for a decreased bandwidth.

(c) Speed Hold Mode

In speed hold mode, the control modulates the throttles tomaintain the commanded airspeed.

The airspeed target input to the autothrottle is firstevaluated by the speed command processor. The speed commandprocessor ensures that maximum airspeeds are not violated(flap, gear, and v~~MMo 1imits). Also, a second-order filterwith rate limiting is applied to speed target changes tosupport the design goal of smooth throttle movements.

22-14-00Page 298.203/298.204

Aug 15/91Use or disclosure of infOrrrIatiOn on this page is subject to the restrictions on the title page of this document.

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Page 298.205/298.206Aug 15/91

Use or dwcloswe Of Information on this page IS subject 10 the restrictions on the htle page Of this document.

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Autothrottle Mode F1OW Diagram


22-14-00Page 298.207/298.208

Apr 15/93

Use of disclosure of information on this page is subject to the restrictions on the title page of this document,

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Page 298.209/298.210Aug 15/91

Use or d!scloswe of Inlormal!on on this page IS subject 10 the restrlct)ons on the title page 01 this document

The airspeed feedback term used in the speed-hold loop isprocessed by a complementary filter to reduce the effects ofwind gusts. A complementary filter is used when rateinformation must be extracted from a signal containing unwantedhigh-frequency noise terms. These unwanted terms must befiltered out without the loss of actual signal informationabove the cutoff frequency of the filter. The requiredcomplement signal is the derivative of the control signal. Itis important that the derivative be obtained from a differentsource to ensure that it does not contain the unwanted noise.In this application, the prime signal is true airspeed and thecomplement signal is aircraft inertial acceleration along theflightpath. An EPR rate command is calculated by applyinglead-lag compensation (with a gain based on aircraft response)to the speed error. This command is biased by a thrust-prediction term which anticipates required thrust adjustmentsthat result from changes in aircraft drag. Thus, compensationis provided for flap, gear, and turn effects on aircraft drag.A second thrust-prediction term, which compensates for changesin flightpath angle, is also included during certain aircraftmaneuvers. The EPR rate command is then integrated and boundedby engine limits to determine the commanded EPR level.

The closed EPR inner loop provides an increased degree ofstability and accuracy to the autothrottle system during speedhold operation. This loop is configured similar to the FLCHcontrol. The resulting throttle rate command is proportionalto EPR error.

4. D. Target Speeds

(1) Automatic Speed Targets

The speed target is set by the speed schedules initialized on theCDU page PERF INIT 2/5. The active speed target follows the flightphase (refer to paragraph 4.D. (l)(d)); in climb the active speed isset by the climb schedule, in cruise it is set by the cruiseschedule, and finally in descent the active speed target is set bythe initialized descent schedule. In cruise-climb andcruise-descent the speed target is the cruise speed.

(a) The current speed command is displayed on page 1 of the activeflight plan. A CAS and MACH are both displayed when climbingor descending; otherwise, the cruise speed target (either a CASor a MACH) is displayed. The active speed target, which is thelesser of CAS or MACH, is shown in large characters. Theactive speed target is also displayed on the GP-820 FlightGuidance Controller when MAN on the GP-820 is not selected.

22-14-00Page 298.211


4. D. (1) (b) The current speed target observes placard speeds and minimumspeed, and is the lesser of the following:

I ● Flap or gear limit speeds● V~O/MO● SDee$’limit below an altitude● blaypoint constraint speeds

. Speed schedule for phase of flight (climb, cruise, descent)

. Latched speeds

(c) The FMS automatically changes the speed target throughout theflight to accommodate aircraft configuration and phase offlight. This automatically controlled speed target isavailable for use by the autopilot or autothrottle, independentof VNAV being selected, for speed control as long as MAN is notselected on the GP-820. If MAN is selected on the GP-820, thepilot controlled speed is used by the autopilot or autothrottlefor speed control when engaged. The automatic FMS speed targetfor a typical flight is changed as follows:

● Prior to takeoff, the speed target is set to the placardspeed for the aircraft configuration.

. During initial climb, as the gear and flaps are retracted,the speed target is adjusted to the placard speed forcurrent configuration.

● Following aircraft cleanup, the speed target is set to 10knots below the speed/altitude limit (250/10,000), ifentered, or the climb speed. If a speed/altitude limit isentered, the speed target is set to climb speed afterpassing the limiting altitude.

● During climb, the speed target is the climb speed scheduleselected during initialization and the transition to MACH isaccomplished automatically. This speed schedule is observedduring any intermediatelevel off.

. The speed target is changed to the cruise speed as theautopilot transitions to ASELor VASEL before reaching thecruise altitude.

. The cruise altitude can be manually entered on PERF INITpage 4, or the FMS will automatically compute an optimumcruise altitude if the OPTIMUM option is used. However, thecruise altitude will be set equal to the altitudepreselector should this be set higher than the entered orcomputed altitude. Cruise altitude entries that are lowerthan the preselector altitude will result in a RESET ALTSELECT message on the CDU scratchpad and will not beaccepted. Once the performance initialization is completed,the cruise altitude can also be entered directly on the PERFDATA page 1. Entering DELETE on either of these two pageshas the effect of recomputing the optimum altitude. Once incruise, entries are ignored, and the cruise altitude willonly be redefined when the altitude preselector is armed.

22-14-00Page 298.212


Honeywell

● The speed target is


changed to the descent speed, providedthe preselector is dialed down, just prior to the TOD if adeceleration is required or at the TOD if no deceleration isrequired. For any descent greater than 6500 feet, anydescent within 100 miles of the TOD, or any descent past theTOD, the speed target is changed to the descent speed (atthe beginning of the descent). Descents started more than100 miles prior to the TOD and descents with less than 6500feet altitude change are considered cruise descents, and thespeed target remains at the cruise speed.

● During descent, the speed target is the descent speedschedule selected during initialization, and the transitionto CAS is accomplished automatically. This speed scheduleis observed during any intermediate level off.

. If a speed/altitude limit is entered, the speed target ischanged to the limit minus 10 knots prior to the limitingaltitude to allow for deceleration, provided the preselectoris set below the limiting altitude.

c When the aircraft is within 20 NM flight plan distance fromthe destination and less than 10,000 feet above thedestination elevation, the speed target is changed to theplacard speed for 10-degree flaps minus 10 knots. Thisallows for the pilot to start extending the flaps whenrequired.

. As the gear and flaps are extended, the speed target ischanged to the placard minus 10 knots for the next expectedflap selection. When landing flaps are selected, (39degrees), the speed target is set to Vref plus 10 knots.Typically, the speed would change to 240 knots at 20 NM, 210knots when 10-degree flaps are selected, 160 knots when20-degree flaps are selected, and some value between 133-158knots when 39-degree flaps are selected.

. Vref displayed on the display controller and on the PFD iscalculated by the FMS and is always based on current weightand configuration. This Vref is accurate and valid evenbelow the 45,000 pound minimum landing weight.

. If go-around is selected, the speed target is changed asdescribed for climb.

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4. D. (1) (d) Phase of Flight Logic

~ The Climb Phase

While on the ground, the power-up flight phase is set toclimb, by default. Level-offs, while in climb, do notchange the flight phase. (See Figure 270.1).

If a descent is started prior to reaching the computedtop-of-climb (TOC) altitude, the flight phase changes todescent. The flight phase does not change back to climbuntil climb is resumed (see Figure 270.2).

The normal way to transition from the climb to the cruisephase is to reach the TOC altitude. There is one exceptionto this rule; 30 NM prior to top-of-descent (TOD) the TOCaltitude will come down to the present altitude if thevertical mode is altitude hold (see Figure 270.3). Thiscauses the flight phase to change to cruise.

~ The Cruise Phase

When cruise altitude is reached, the flight phase stays incruise until another climb, or the descent is started. Aclimb or descent is recognized by an ASEL ARM towards a newpreselected altitude. If a climb is initiated, a newTOCaltitude is placed at the preselector, but if a descent isstarted the cruise altitude does not change.

There are two sub-phases to the cruise phase calledcruise-climb and cruise-descent. These sub-phases aredefined as climbs or descents less than 6,500 feet (seeFigure 270.4). Note that when a cruise-descent takes placethe cruise altitude is moved down. Also, a cruise-descentis only allowed above 15,000 ft.

Descents within 100 NM of the planned TOD are always treated

as a descent, regardless of the altitude change (see Figure

270.5) .

TOC

CLIMB

AD-31723@

Climb Phase With No DescentsFigure 270.1 22-14-00

Page 298.214Apr 15/93


TOC

AD-31729@

Climb Phase With DescentFigure 270.2

TOC

ITOC ~----------- ., ‘*D● ●

CLIMB .=”” CRUISE ●9’*

AD-31730@

Transition From Climb to Cruise PhaseFigure 270.3

TOC

CRUISE

CRUISE /xK~:ANCRUISE-DESCENT

CRUISE-CLIMB ABOVE 15,000 FT

AD-31 731@

Cruise-Climb and Cruise-Descent SubphaseFigure 270.4 22-14-00

Page 298.214.1Apr 15/93


+LESS THAN 100 NM

CRUISE . . . . . . . . . . . . . ...-..* . . . . .

AD-31 732@

Descents Within 100 NM of TODFigure 270.5

4. D. (1) (d) ~ The Descent Phase

The flight phase changes to descent once ASEL ARM towards alower preselector altitude is detected, unless acruise-descent is flown. After leveling off, while indescent, the flight phase will change back to cruise, if thelevel segment is predicted to be at least 100 NM (seeFigures 270.5 and 270.6). Typically, the level segments areshorter, and the flight phase stays in descent all the waydown.

Once in descent, the flight phase can go back to climb if aclimb is initiated (see Figure 270.7).

If a missed approach is executed (see Figure 270.8), theflight phase transitions to climb. A new cruise altitude isdefined at the preselector altitude, and a shortcruise-segment will be flown before flight phase transitionsto descent.

~ The Auto EPR Rating Transitions

The auto EPR rating is set to TO on the ground. FLEX isalso boxed if selected on the TAKEOFF INIT pages. As longas the flight guidance computer (FGC) mode is TO the autoEPR rating will remain TO as well. The TO rating will bekept until passing 400 ft. above the field elevation underall circumstances.

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Apr 15/93


The transition to CLB takes place when a vertical mode(VNAV, FLC, ALT or VS buttons) is engaged. Capturing analtitude will also cause the auto EPR rating to transitionto CLB. Simply engaging the autopilot in pitch hold willnot cause a transition to CLB. The auto EPR rating goes toGA because the FGC TO mode is cancelled. The auto EPRrating is set to CLB as long as the flight phase is climb,even during level-offs, prior to TOC.

When the flight phase is cruise, the auto EPR rating is CRZ,except for cruise-climbs then the phase becomes CLB. When

the flight phase is descent, the auto EPR rating stays with

CRZ, except when the gear comes down for approach. The autoEPR then changes to GA.

Whenever the FGC GA mode is selected, the auto EPR ratingwill go to GA. If an engine failure occurs, the auto EPRrating goes to MCT, provided the rating was not TO or GAwhen the engine failure occurred. Upon landing, the autoEPR rating returns to TO.

TOD

CRUISE

CRUISE TOD

MORETHAN1OONM

AD-31733@

Descent to Cruise Phase When MoreThan 100 NM from TOD

Figure 270.6

22-14-00Page 298.214.3


TOD

CUMB

AD-31734@

Descent to Climb PhaseFigure 270.7

TOC CRUISE TOD TOC TOD

AD-31735@

Missed Approach Flight PhaseFigure 270.8

(2) Speed Constraints

(a) Speed constraints may be entered at waypoints and are displayedin large characters, unless a descent angle has been manuallyentered, in which case the speed constraint is in smallcharacters above the angle and altitude entry.

(b) Climb speed constraints are observed until reaching theconstraining waypoint.

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Apr 15/93


4. D. (2) (c) Cruise speed constraints greater than the perf init cruise

speed are observed after passing the constraining waypoint.

(d) If a cruise or descent speed constraint requires a

deceleration, the deceleration begins prior to the constraining

waypoint to be on speed at the waypoint.

(e) Any speed constraint can be deleted by selecting “*DELETE*” tothe appropriate line on the active flight plan pages. Thisalso deletes any altitude constraint.

(f) If the airspeed is not slowing sufficiently to meet a lowercruise/descent speed constraint, a message CHECK SPEEDCONSTRAINT will be generated until the aircraft is anticipatedto meet the speed constraint at the waypoint or the message iscleared using the CDU CLEAR button.

(3) Latched Speeds

If VNAV automatically transitions to the flight level change mode,the current speed may be latched as the speed target. For example,when flying a path with the preselector set lower than the last

altitude constraint, as the aircraft passes the last altitude

constraint, VNAV automatically transitions to the flight level

change mode and continues the descent toward the preselector. Toavoid possible pitch changes, the current airspeed is latched as thespeed target if the current speed is 5 knots or more less than thecurrent target speed. This is noted on the active flight plan pagedisplay of the speed target by changing the title to LATCHED. Thelatched speed can be removed by selecting *DELETE* to the speedcommand line.

(4) Speed Protection

The FMS provides speed protection for the aircraft. Protectedspeeds include:

. V~~MO and buffet limits● Spee$ limit below an altitudec Placard speeds● Minimum speeds (1.3 Vs)

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4. D. (5) Vertical Flight Plan (Figure 271)

The Honeywell FMS is structured such that the pilot can enter allinformation for the entire flight prior to takeoff. In flight, theoperation of the FMS can be limited to checking the progress of theflight and updating the system with any in-flight changes asnecessary. During ground operations, the pilot is prompted in thelower right-hand corner of the screen on the normal fourpredeparture steps. The four steps are:

. Verify date, time, and database

. Initialize position

. Enter flight plan

. Initialize performance

An 8-waypoint example flight will be used. The performanceinitialization data used are as follows:

● Climb Speed Schedule 300/0.75● Cruise Speed Schedule 0.78● Descent Speed Schedule 300/0.80. Transition Altitude 18,000● Speed/Altitude Restriction 250/10,000. Cruise Altitude 45,000

Pressing the PERF DATA prompt allows the PZ-800 to calculate theperformance values for the entire flight plan. Airspeed constraintswill be entered into the flightplan at waypoints 2, 3, 4, 5, 6, and7 as follows:

WPT 2 320 ktWPT 3 280 ktWPT 4 0.80 MWPT 5 0.75 MWPT 6 0.82 MWPT 7 0.78 M

Figure 271 shows the vertical flight plan for the above example.

22-14-00Pa~e 298.215‘Aug 15/91


CRUISE SPD

45,000 FT TOC 0.78M WPT 4 A WPT 5 TOD

CRUISE ALT T DESCENT SPD

0.80M 0.75M280 KTS

0.82M

320 KTS

CLIMB SPD0.78M

300/0.7518,000 FT

10,000 Fr250 KTS 250 KTS

WPT 8

AD-3061 6#

Example Vertical Flight PlanFigure 271

The SPD/ALT restriction entry made is the speed and altitude belowwhich a speed limit applies. The speed target of the FMS is 10knots below the value entered, so in this example the speed targetbelow 10,000 feet will be 240 knots.

Table 211 shows the climb and descent schedule for flaps as well asthe gear limitation speed.

Flar)s(deci) Climb Phase (kt) Descent Phase (kt)

250 210;: 220 16039 170 VREF+1O

Gear = 225

Climb and Descent ScheduleTable 211

~: The flaps down speed schedule is based on theof flight (refer to paragraph 4.D.(l)d.). IfASEL is armed (60 ft/min for 3 seconds in the

current phaseon an approachdirection of

the altitude p~eselect value) with the altitude preselectorabove the aircraft then the phase of flight will change toclimb and the speed schedule will revert to the climbschedule. Capturing GS or arming ASEL to a preselectedaltitude below the aircraft will change the flight phase backto descent and return the descent speed schedule.

22-14-00Page 298.216


4. D. (5) (a) Climb Phase

Climb speeds can be altered on the following CD-81O CDU pages:

● Initialization climb speed - PERF INIT 2/5 or CLIMB PAGE

. Constraint speed - FLIGHTPLAN X/X

. Restriction speed - PERF INIT 4/5

Climb speed rules:

. Rule 1 - Speed constraints > initializationclimb speeds areignored.

● Rule 2 - Speed constraints < initializationclimb speeds arehonored until the constraint waypoint is sequenced. Oncethe constraint waypoint is sequenced the target speedreturns to the initialization climb speed.

. Rule3 - Initialization climb speeds can be changed at anytime and will cause the active target speed to change if nospeed constraint is active.

Figure 272 shows the climb phase of the flight. The followingparagraphs refer to the circled numbers shown.

22-14-00Page 298.217


MAINTENANCE

Honeywell !!!!t!!!&.

WPT

TOC45,000 Fr

WPT 3

WPT 2

18,000 FT

10,000 Fr

1AD-30617#

Climb PhaseFigure 272

At point @the aircraft is on the ground ready for departure,with flaps at 20. The speed target at this point will be 220knots. Since the airport traffic area speed restriction isusually 200 knots, a manual speed target must be entered untiloutside this area, or 200 may be entered as the desired climbspeed on the performance initialization, page 2.

9At point 2 the aircraft is in a climb and the flaps areraised to O degrees. Although the flap placard for flaps 10is 250 knots, the speed target changes to 240 knots due to thespeed restriction below 10,000 feet at 250 knots (the systemalways uses 10 knots less than the entered speed restriction toprevent overshoots).

9At point 3 the speed target would change automatically to 250knots (FLA S 10 placard). When the flaps are raised to zerodegrees, the speed target will automatically change to theclimb speed schedule of 300 knots as entered in i)erformanceinitialization. 22-14-00

Page 298.218Aug 15/91


At point @ all altitudes will be displayed as flight levelsin the FMS except for descent predictions less than 18,000feet. Altitudes greater than 18,000 feet are displayed asflight levels even when the aircraft is on the ground. If thebaro correction is not changed to 29.92, the message CHECK BAROSET will be displayed in the CDU scratchpad. This message isdisplayed under the following conditions:

●

●

In a climb, if the baro set is not 29.92 and the aircraftaltitude equals the transition altitude (18,000 feet in thisexample) plus 3000 feet, or the aircraft levels off betweenthe transition altitude plus 3000 feet above the transitionaltitude.

In a descent, if the baro set is 29.92 and the aircraftaltitude equals the transition altitude minus 3000 feet, orthe aircraft levels off between the transition altitude minus3000 feet below the transition altitude.

At point ~ a constraint speed of 320 knots is entered on theactive flight plan page adjacent to WPT2, but since 320 islarger than the entered 300 knots entered in PERF INIT, thespeed constraint is ignored and 300 remains as the speed target(refer to rule 1).

At point ~ a constraint speed of 280 knots is enteredadjacent to WPT3. The active speed target changes to 280 knots(refer to rule 2).

At point ~ the active speed target changes back to theinitializationclimb speed of 300 knots when the waypoint issequenced (refer to rule 3).

The speed target will automatically change to a Mach target(0.75 M) at the appropriate point or if selected on the GP-820Flight Guidance Controller via the IAS/MACH pushbutton.

At point ~ the aircraft levels off at the cruise altitude of45,000 feet and the speed target automatically changes to thecruise speed of 0.78 M as entered in the performanceinitialization.

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Apr 15/93Use or disclosure of information on this page issubject to the restrictions on the title page of thks document.

Honeywell !!%!!~.c’4. D. (5) (b) Cruise Phase

Cruise speed rules:

● Rule 1 - All constraint speeds are honored.

Q Rule 2 - When a constrained waypoint is sequenced, theactive target speed is latched to the constraint speed.

. Rule 3 - Entering DELETE on line lR of the active flightplan page (CMD SPD) will delete the latched target speed andreturn the target speed to the initial cruise speedschedule.

. Rule4 - Cruise speeds can be changed at any time. Theactive target speed will not change to a new cruise speedschedule if the current target speed is latched to aconstraint speed.

Figure 273 shows the cruise phase of the flight. The followingparagraphs refer to the circled numbers shown.

6363 @

TOCA

‘WPT 4

AD-3081 8#

Cruise PhaseFigure 273

22-14-00Page 298;220


Honeywell ?##~&wcE

At point @ the current target speedcruise speed schedule).

is 0.78 M (from the

At point @ the target speed changes (and latches) to 0.80 M.

At point ~ the target speed changes to 0.75 M to allow theaircraft to decelerate into WPT5.

At point ~ the target speed is now latched to 0.75 M.

PAt point 13 , if it is desired to return to the cruise speedschedule o 0.78 M, line select DELETE to lR on the activeflight plan page (CMD SPD).

At point @ the aircraft begins its descent and the speedtarget automatically changes to the descent speed schedule of0.80 M as entered in the performance initialization.

4. D. (5) (c) Descent Phase

Descent speed rules:

. Rule 1 - Only constraint speeds less than the descent speedschedule are observed.

● Rule 2 - Target speeds will latch to the constraint speed.

. Rule3 - The descent speed schedule can be changed at anytime. If the current target speed is latched, only changesto the descent speed schedule less than the latched speedwill be honored. Speeds entered above the latched speed arelimited to the latched speed. If the current latched speed

is deleted then all changes to the descent speed schedulewill be honored, assuming speed limits are not violated.

Figure 274 shows the descent phase of the flight. Thefollowing paragraphs refer to the circled numbers shown.

22-14-00Page 298.221


Honeywell $%!!!.”

VdPT 8

AD-30619#

Descent PhaseFigure 274

At point @ the top of descent (TOD) has been reached and theaircraft begins its descent. The target speed automatical1ychanges to the descent speed of 0.80 M.

At point ~ there is no change to the current target speed(refer to rule 1).

At point @ the target speed changes to 0.78 M to allow theaircraft to decelerate into WPT 7.

At point ~ the current speed target is now latched to0.78 M, but switches over to 300 KCAS when reaching the0.78 M/300 crossover altitude.

At point ~ deleting the SPD CMD on the active flight planpage returns the target speed back to the descent speedschedule of 300/0.80.

22-14-00Page 298.222


9At point 20 the speed target automatically changes to 240

knots to a low the aircraft to decelerate to the SPD/ALT LIMIT.

9At point 21 the aircraft is within 20 NM from the destination

(WPT 8) an less than 10,000 feet above the destination

elevation so the speed target changes to the 10-degree flapplacard minus 10 knots (240 knots). In this example the speedtarget would already be 240 knots, so no change would benoticed.

At point @ the flaps are lowered to 10 degrees, which causesthe speed target to automatically change to 210 knots, which is10 knots below the 20-degree flap placard.

At point @ the flaps are lowered to 20 degrees, which causesthe speed target to automatically change to 160 knots, which is10 knots below the 39-degree flap placard.

At point @ the flaps are lowered to 39 degrees, which causes

the speed target to automatically change to VREF plus 10 knots.

The speed targets, whether manual or automatic, can be flown bythe flight guidance system or the autothrottle system.Whenever the autopilot is controlling speed, such as FLCH orVFLCH, the autothrottle controls thrust. The autothrottle willcontrol speed if engaged for all other flight director modes,except TO/GA.

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Apr 15/93Useor disclosure of information onthispage issubject to the restrictionsonthetitle page of thisdocument,

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